Beechcraft 1900 Airliner Pilot Training Manual – PDF Download

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BEECH 1900 AIRLINER PILOT TRAINING MANUAL VOLUME 2 AIRCRAFT SYSTEMS

FlightSafety International, Inc. Marine Air Terminal, LaGuardia Airport Flushing, New York 11371 (718) 565-4100 www.flightsafety.com

Courses for the Beech 1900 Airliner and other Beech aircraft are taught at the following FlightSafety learning centers:

Wichita (Raytheon) Learning Center 9720 East Central Avenue Wichita, KS 67206 Phone: (316) 685-4949 Toll-Free: (800) 488-3747 Fax: (316) 685-2476 LaGuardia Learning Center Marine Air Terminal LaGuardia Airport Flushing, NY 11371-1061 Phone: (718) 565-4170 Toll-Free: (800) 749-8818 Fax: (718) 565-4174 Toronto Learning Center 95 Garratt Boulevard Downs View Ontario Canada M3K 2A5 Phone: (416) 638-9313 Toll-Free: (877) 359-32748 Fax: (416) 683-3348

Copyright Š 2000 by FlightSafety International, Inc. All rights reserved.

Printed in the United States of America.

F O R T R A I N I N G P U R P O S E S O N LY

NOTICE The material contained in this training manual is based on information obtained from the aircraft manufacturer’s pilot manuals and maintenance manuals. It is to be used for familiarization and training purposes only.

At the time of printing it contained then-current information. In the event of conflict between data provided herein and that in publications issued by the manufacturer or the FAA, that of the manufacturer or the FAA shall take precedence.

We at FlightSafety want you to have the best training possible. We welcome any suggestions you might have for improving this manual or any other aspect of our training program.

F O R T R A I N I N G P U R P O S E S O N LY iii

CONTENTS SYLLABUS Chapter 1 AIRCRAFT GENERAL Chapter 2 ELECTRICAL POWER SYSTEMS Chapter 3 LIGHTING Chapter 4 MASTER WARNING SYSTEM Chapter 5 FUEL SYSTEM Chapter 6 AUXILIARY POWER UNIT Chapter 7 POWERPLANT Chapter 8 FIRE PROTECTION Chapter 9 PNEUMATICS Chapter 10 ICE AND RAIN PROTECTION Chapter 11 AIR CONDITIONING Chapter 12 PRESSURIZATION Chapter 13 HYDRAULIC POWER SYSTEMS Chapter 14 LANDING GEAR AND BRAKES Chapter 15 FLIGHT CONTROLS Chapter 16 AVIONICS Chapter 17 MISCELLANEOUS SYSTEMS WALKAROUND APPENDIX ANNUNCIATOR PANEL INSTRUMENT PANEL POSTER

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CHAPTER 1 AIRCRAFT GENERAL CONTENTS Page INTRODUCTION .................................................................................................................. 1-1 GENERAL .............................................................................................................................. 1-1 AIRPLANE SYSTEMS .......................................................................................................... 1-2 General ............................................................................................................................ 1-2 Chapters........................................................................................................................... 1-2 STRUCTURES ....................................................................................................................... 1-4 Beechcraft 1900 Airliner Description ............................................................................. 1-4 Beechcraft 1900 Airliner Configurations ........................................................................ 1-6 STANDARD EQUIPMENT................................................................................................... 1-7 DIMENSIONS AND SPECIFICATIONS.............................................................................. 1-8 Airplane Dimensions ....................................................................................................... 1-8 Specifications .................................................................................................................. 1-8 DOORS AND INTERIORS.................................................................................................. 1-14 Cabin Entry and Exits.................................................................................................... 1-14 Airstair Locking Mechanism......................................................................................... 1-14 Airstair Door Operation................................................................................................. 1-15 Emergency Exits............................................................................................................ 1-16 Cargo Door .................................................................................................................... 1-16 Cabin Compartments..................................................................................................... 1-17 Flight Deck .................................................................................................................... 1-19 Control Surfaces ............................................................................................................ 1-32

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TIEDOWN AND SECURING............................................................................................. 1-33 TAXIING.............................................................................................................................. 1-34 SERVICING DATA............................................................................................................. 1-34 PRODUCT SUPPORT ......................................................................................................... 1-36 EXTERIOR INSPECTION .................................................................................................. 1-38

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ILLUSTRATIONS Figure

Title

Page

1-1

1900 Airliner General Arrangement—Series UE.................................................... 1-4

1-2

1900 Airliner General Arrangement—Series UB.................................................... 1-5

1-3

Stabilons and Tailets................................................................................................ 1-6

1-4

1900 Airliner Three-View Diagram—Series UB and UC....................................... 1-9

1-5

1900 Airliner Three-View Diagram—Series UE .................................................. 1-10

1-6

Rear Three-Quarters View..................................................................................... 1-12

1-7

Cabin Pressure Vessel............................................................................................ 1-12

1-8

Cabin Altitude Chart.............................................................................................. 1-14

1-9

Entrance and Exit Provisions—Series UB and UC ............................................... 1-15

1-10

Cargo Door ............................................................................................................ 1-16

1-11

Cabin Areas ........................................................................................................... 1-17

1-12

Cabin Seating Layouts........................................................................................... 1-18

1-13

Flight Deck ............................................................................................................ 1-19

1-14

Control Wheels ...................................................................................................... 1-20

1-15

Right Side Panel—Series UA, UB, UC................................................................. 1-21

1-16

Fuel Control Panel ................................................................................................. 1-21

1-17

Instrument Panel .................................................................................................... 1-22

1-18

Instrument Panel Diagram ..................................................................................... 1-23

1-19

Engine Instruments—Series UA, UB, UC ............................................................ 1-24

1-20

Engine Instruments—Series UE-1 through UE-92................................................ 1-25

1-21

Engine Instruments—Series UE-93 and Subsequent ............................................ 1-26

1-22

Instrumentation—Series UA, UB, UC .................................................................. 1-27

1-23

Instrumentation—Series UE .................................................................................. 1-28

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1-24

Power Control Quadrant and Pedestal .................................................................. 1-29

1-25

Pilot’s and Copilot’s Subpanels—Series UA, UB, UC......................................... 1-29

1-26

Pilot’s and Copilot’s Subpanels—Series UE ........................................................ 1-30

1-27

Annunciators—Series UA, UB, UC...................................................................... 1-31

1-28

Lighting Controls—Series UA, UB, UC............................................................... 1-31

1-30

Flight Control Locks ............................................................................................. 1-32

1-29

T-Tail Configuration ............................................................................................. 1-32

1-31

Airplane Secured ................................................................................................... 1-33

1-32

Ground Turning Clearance—Series UA, UB, UC ................................................ 1-34

1-33

Ground Turning Clearance—Series UE................................................................ 1-35

1-34

Danger Areas......................................................................................................... 1-36

1-35

Servicing Data ....................................................................................................... 1-37

1-36

Exterior Inspection ................................................................................................ 1-38

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CHAPTER 1 AIRCRAFT GENERAL

INTRODUCTION A good, basic understanding of the airplane will help in studying individual systems and their operations. This chapter provides preliminary background information related to airplane systems, operational considerations, and performance, which are presented in other chapters of this training manual. Also, this chapter presents an overall view of the airplane, including external familiarization, cabin arrangements, and cockpit layout. In this chapter, you will find diagrams and data describing the airplane in general, including systems not found in the POH/AFM.

GENERAL There are four series of the Beech 1900 Airliners and several cabin configurations. The primary difference between the UA and UB serial numbered airplanes and the UC and UE series is in the fuel system and the caution/advisory annunciator panel arrangement. The UC and UE series have a wet-wing fuel system, while UA and UB series have a series of bladder tanks. The UB, UC, and UE series have cargo doors, while the

UA series airplanes have an airstair door in the rear. This training manual covers all four series of the 1900 Airliners.

Reference material in this training manual is organized into 15 chapters (with two unused tabs) covering all airplane systems. Each chapter is complete and independent and can be referred to in any sequence.

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Following are brief descriptions of the subject matter in each chapter. All material is discrete to the Beechcraft 1900, 1900C, and 1900D Airliners.

AIRPLANE SYSTEMS General The POH/POM Systems Description section briefly summarizes all 1900 Airliner systems.

Additional descriptions and details are included in separate chapters of this training manual.

POH/POM information is updated as required and always supersedes any information in the training manual.

CHAPTERS

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Master Warning System Chapter 4, “Master Warning System,” presents a description and discussion of the warning, caution, and advisory annunciator panels. Each annunciator is described in detail, including its purpose and associated cause for illumination.

Emphasis is on corrective action required by the pilot if an annunciator is illuminated.

Fuel System Chapter 5, “Fuel System,” presents a description and discussion of the fuel system. The physical layout of fuel cells and normal/abnormal fuel system operations are described. Correct use of the boost pumps, fuel transfer system, and firewall shutoff valves are discussed. Locations and types of fuel drains and correct procedures for taking and inspecting fuel samples are detailed.

This chapter includes a list of approved fuels and procedures for fuel servicing.

Aircraft General Chapter 1, “Aircraft General,” presents an overall view of the airplane. This includes external familiarization, cabin arrangement, and cockpit layout. In this chapter you will find additional general descriptions, diagrams, and data which may not be found in the Pilot or Airplane manuals.

Electrical Power System Chapter 2, “Electrical Power System,” describes the airplane electrical system and its components. The electrical system is discussed to the extent necessary for pilot management of all normal and emergency operations. The location and purpose of switches, indicators, lights, and circuit breakers are noted. DC and AC generation and distribution are described. This chapter also includes electrical system limitations and a discussion of potential electrical system faults.

Lighting

Powerplant Chapter 7, “Powerplant,” presents a discussion of the Pratt & Whitney PT6A-65B and -67D turboprop engines. Engine theory and operating limitations are described, and normal pilot procedures are detailed. Crewmembers must have sufficient knowledge of the PT6A series engines to understand all normal and emergency procedures.

This chapter also describes propeller construction and system components. Location and use of propeller controls, governor operation, overspeed protection, reversing, manual feathering, and autofeather are discussed.

Fire Protection Chapter 8, “Fire Protection,” describes the fire warning and protection systems. Operation and testing information for the fire detection and fireextinguishing systems is included.

Chapter 3, “Lighting,” discusses cockpit lighting, cabin lighting, and exterior lighting. All lights are identified and located. The location and use of controls for the lighting system are also included.

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Pneumatics

Landing Gear and Brakes

Chapter 9, “Pneumatics,” presents a discussion of pneumatic and vacuum systems. Sources and operation of pneumatic and vacuum air are described. Acceptable gage readings and normal and abnormal system indications are outlined.

The bleed-air warning system is described as an integral function of the pneumatic system.

Chapter 14, “Landing Gear and Brakes,” outlines control and operating limitations of the hydraulic landing gear system. Normal and abnormal indications are discussed. Procedures are described for use of the alternate landing gear extension s y s t e m i n t h e ev e n t o f n o r m a l s y s t e m malfunction.

Ice and Rain Protection

This chapter also details the power steering system and includes procedures for normal and abnormal indications.

Chapter 10, “Ice and Rain Protection,” describes theory and operation of anti-ice and deice systems. Each ice protection device in the aircraft is described individually, showing each component and control location. Normal and abnormal operations of each anti-ice and deice system are considered. The purpose of this section is to acquaint the pilot with all systems used for flight into icing conditions or heavy rain. Information concerning preflight deicing and defrosting includes a discussion of appropriate fluids approved for these procedures.

Air Conditioning Chapter 11, “Air Conditioning,” presents a description of air-conditioning, bleed-air heating, and fresh air ventilation systems. The air cycle machine is discussed in detail as an integral function of the entire system. Each environmental subsystem description includes general operating practices, control of the system, and emergency procedures.

Pressurization System Chapter 12, “Pressurization,” presents a description of the pressurization system components.

Component locations and system operational controls are discussed. The importance of airstair and cargo compartment door security is noted, and emergency procedures are described. Where necessary, references are made to the environmental system as it affects pressurization.

Correct use of the wheel brakes and parking brakes are described in this chapter. System components and brake wear indications are also detailed. Safety procedures related to system operation are considered.

Flight Controls Chapter 15, “Flight Controls,” describes the foursegment Fowler-type flap system. System controls and limitations are considered, and operations are outlined as referenced in the POH/AFM. All normal and abnormal operating procedures are discussed.

Avionics Chapter 16, “Avionics,” describes the standard avionics installation for the 1900 Airliner. Communication and navigation radio equipment and audio panel layout are detailed. Non-standard avionics are not covered here; however, for airc r a f t e q u i p p e d w i t h o p t i o n a l av i o n i c s installations, supplements are available to describe equipment operation. A glossary of avionics terminology is included in an Appendix at the back of this training manual.

This chapter also presents a discussion of the dual pitot-static system. Pitot-static system input to flight instruments and related safety features are described. Operational principles, pitot-static air sources, and component locations are given for the system. Pilot and copilot alternate static air systems are also covered.

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Miscellaneous Systems Chapter 17, “Miscellaneous Systems,” presents a summary of the oxygen system and its components. General descriptions, operating principles, system controls, and emergency procedures are included. Use of the oxygen duration chart is outlined, and oxygen availability is determined, using practice problems to illustrate various flight situations. FAR requirements are discussed as they relate to crew and passenger oxygen needs, including types and availability of oxygen masks. The Pilot’s Operating Handbook and Pilot’s Operating Manual are referred to for descriptions of system servicing procedures.

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STRUCTURES BEECHCRAFT 1900 AIRLINER DESCRIPTION The Beechcraft 1900 Airliner (Figures 1-1 and 1-2) is a high-performance, pressurized, twinengine, turboprop airplane designed and equipped for day or night flight in IFR conditions, high-density air traffic zones, and known icing conditions. It is also capable of operating in and out of small, unimproved airports within POH/AFM operating limits.

The 1900 Airliner design blends a highly efficient airframe with proven high-technology components to provide a reliable, economical,

Figure 1-1 1900 Airliner General Arrangement—Series UE

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and versatile airplane. The 1900 Airliner is a growth version of the Beechcraft Super King Air B200 corporate aircraft. Many King Air features were retained, including the flight deck design, wing airfoil section, classic T-tail, and PT6 freeturbine engine.

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The structure is an all-metal, low-wing monoplane. It has fully cantilevered wings and a T-tail empennage. Efficient, high aspect ratio wings provide an excellent combination of low drag for cruise conditions and easy handling at slow speeds for airport operations.

Figure 1-2 1900 Airliner General Arrangement—Series UB FOR TRAINING PURPOSES ONLY

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BEECHCRAFT 1900 AIRLINER CONFIGURATIONS The Beechcraft 1900 Airliner is certified to carry up to 19 passengers. In addition to standard airplane configurations, Beechcraft offers optional items at additional cost and weight. Basic configurations, dimensions, weights, and specifications are summarized later in the chapter. Refer to the current airplane POH/AFM for detailed, up-todate information.

The wing is fabricated as a one-piece unit consisting of a one-piece spar, center wing, and two outboard wing assemblies. The center wing forms an integral part, providing structural support for engine nacelles and outboard wing assemblies. The UE Series has incorporated winglet design based on NASA research. This enhances climb performance and extends the effective wing span to improve range.

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effective horizontal surface, which is able to control greater excursions in CG loading. In “free” air, trim changes resulting from changes in power settings or flap positions are minimized or eliminated entirely. Material fatigue, present when a surface is embedded in the high-energy propeller slipstream, is also eliminated.

Small horizontal surfaces called stabilons (Figure 1-3) are mounted on both sides of the fuselage just aft of the cargo door. Two smaller vertical fins, called tailets, are mounted on the lower horizontal stabilizer tips.

Faired, oval nacelles of minimum frontal area are installed on the center wing to house the engine and landing gear. Nacelle location and design maximize propeller-to-ground clearance, minim i z e c a b i n n o i s e , a n d p r ov i d e l ow d r a g powerplant cowlings. The upper nacelle surface fairs into the upper surface of the wing near the rear spar. The lower nacelle surface extends aft below the wing to house the retracted main landing gear. Nacelle surfaces are flush-riveted for aerodynamic smoothness.

The T-tail on the Beechcraft 1900 Airliner was designed to improve aerodynamics, lighten control forces, and increase CG range. Extensive wind tunnel and flight tests and the experience of thousands of pilots have confirmed that the combination of T-tail and straight wing provides excellent control characteristics in all flight regimes.

The T-tail configuration was chosen for the 1900 Airliner for several reasons. Although distinctive appearance and excellent handling qualities were important objectives, the operator of the 1900 Airliner needs flexibility in overall airplane usage and loading. By putting the horizontal tail at the top of the vertical fin, the tail surface “flies” in less disturbed air than it would if located lower on the fin, or on the fuselage. The result is a more 1-6

Figure 1-3 Stabilons and Tailets

Stabilons were added to improve pitch stability for the unusually wide CG range (4 to 40% MAC). They provide exceptionally positive recovery characteristics from the deep stall flight regime. For additional directional stability, tailets are teamed with a pair of vortex generators, ahead of the wing-fuselage junction. Vortex generators reduce stall speeds and interference drag to increase directional stability. This package of small aerodynamic surfaces—stabilons, tailets, and vortex generators—provides the 1900 Airliner with highly desirable stability characteristics without using artificial and expensive electronic stability augmentation systems.

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The fuselage is a conventional monocoque structure of high-strength aluminum alloys. The basic cabin shape is a favorable compromise between passenger comfort and efficient cruise performance. The cabin profile is squared-oval, not round. Passengers can sit upright comfortably without leaning to accommodate sloping walls.

The floors are flat from side to side for passenger ease in entering and leaving the cabin. In the UE Series, the cabin height has been raised to 71 inches.

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●

The basic King Air B200 fuselage was retained, adding windows to accommodate the increase in passenger seats. The cabin is pressurized for operation at its most efficient cruise altitude.

These design characteristics result in a structure that is optimized for pressure loads, is lightweight, and will carry 19 passengers comfortably.

Principal differences between UA/UB Series, UC Series, and UE Series Airliners are as follows: ●

❍

Has PT6A-67D engines rated at 1,279 shp.

❍

Maximum Ramp Weight increased to 17,230 pounds.

❍

Maximum Takeoff Weight has increased to 17,120 pounds.

❍

Maximum Landing Weight has increased to 16,765 pounds.

❍

Maximum Zero Fuel Weight has increased to 15,165 pounds.

❍

Cockpit has changed to Collins EFIS 84 system with flight directors as standard equipment.

has

Other improvements and differences are included in the various chapters of this training manual.

UC Series: ❍

❍

●

UE Series:

Has wet-wing fuel system with capacity increased to 667 gallons usable fuel.

Has new style ENGINE FIRE TEST switches. The UA/UB series airplanes have a mod kit available to change the rotary switch with the new style arrangement.

UC/UE Series: ❍

❍

❍

Fuel control panel on the pilot’s left side has added auxiliary fuel pump switches and MAIN or AUX fuel quantity selection switch.

STANDARD EQUIPMENT The following is standard equipment on the Beech 1900 Airliner: ●

Reversible pitch propellers

●

Propeller synchrophaser and synchroscope

●

Autofeather propeller system

●

Complete anti-ice and deice systems:

Has additional caution/advisory annunciators in a slightly modified panel.

Has a modified switch panel arrangement on the control pedestal.

●

Revision 1

❍

Surface deice boots

❍

Electric propeller anti-ice

❍

Engine anti-ice

❍

Dual heated windshields

❍

Brake deice

❍

Heated pitot tubes

Dual bleed-air heat

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Air-conditioning system:

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No-smoking sign with chime

❍

Air cycle machine

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Forward baggage compartment

❍

Freon air conditioner

●

EFIS 84 system (UE Series)

❍

Aft cabin blower

NOTE

●

Dual-pane cabin windows

●

Fire protection system: ❍

Cabin fire extinguisher

❍

Cockpit fire extinguisher

❍

Engine fire detection

❍

Engine fire extinguishers

Rear cargo door 52 x 52 inches is standard on 1900C configuration.

DIMENSIONS AND SPECIFICATIONS AIRPLANE DIMENSIONS

●

19 removable, track-mounted, high-back seats

Figure 1-4 shows airplane dimensions for Series UB and UC, while Figure 1-5 shows airplane dimensions for Series UE.

●

Fixed-step forward airstair door with entrance step lights

SPECIFICATIONS

●

Oxygen system—Dual 76.5-cubic-foot cylinders and 21 oxygen outlets

Minimum crew—FAA certificated: Series UA, UB, UC ................................ One Series UE................................................ One

●

Exterior lighting:

Maximum passengers—FAA certificated: Series UA, UB, UC ................................... 19 Series UE................................................... 19

❍

Wing deice lights

❍

Tail logo lights

❍

Strobe lights

❍

Wing-mounted landing lights

❍

Nosewheel single taxi light

❍

Recognition lights (UE Series)

Engine: Series UA, UB, UC ................... 2PT6A-65B Series UE................................... 2PT6A-67D Propellers—Four-blade, reversible: Series UA, UB, UC .................. Two Hartzell Series UE.................................. Two Hartzell

●

Control wheel-mounted clocks—Pilot and copilot

●

Dual instantaneous indicators

●

Flight hour recorder

●

Electric and manual elevator trim system

vertical

speed

Landing gear—Retractable, tricycle, dual main wheels: Series UA, UB, UC ....................... Hydraulic Series UE....................................... Hydraulic Wing area: Series UA, UB, UC ........... 303.0 square feet Series UE........................... 310.0 square feet

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Figure 1-4 1900 Airliner Three-View Diagram—Series UB and UC

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Figure 1-5 1900 Airliner Three-View Diagram—Series UE

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Maximum Certificated Weights Maximum ramp weight: Series UA, UB, UC................ 16,710 pounds Series UE ............................... 17,230 pounds

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Cargo door height aft: Series UE .................................... 53.8 inches

Maximum Weight in Baggage Compartments

Maximum takeoff weight: Series UA, UB, UC................ 16,600 pounds Series UE ............................... 17,120 pounds

Nose: Series UA, UB, UC..................... 150 pounds

Maximum landing weight: Series UA, UB, UC................ 16,100 pounds Series UE ............................... 16,765 pounds

Forward cabin compartment: Series UA, UB, UC..................... 250 pounds Series UE .................................... 250 pounds

Maximum zero fuel weight: Series UA, UB, UC................ 14,000 pounds Series UE ............................... 15,165 pounds

Hanger—Forward cabin compartment: Series UA, UB, UC..................... 100 pounds Series UE .................................... 100 pounds

Typical operating weight: Series UE ............................... 10,550 pounds

Aft baggage compartment—Forward Section: Series UA, UB, UC..................... 250 pounds Series UE ................................. 1,000 pounds 1900C.......................................... 880 pounds

Cabin and Entry Dimensions Cabin width (maximum): Series UA, UB, UC........................ 54 inches Series UE ....................................... 54 inches Cabin length (maximum between pressure bulkheads): Series UA, UB, UC................... 473.5 inches Series UE .................................. 473.5 inches Cabin height (maximum): Series UA, UB, UC........................ 57 inches Series UE ....................................... 71 inches Airstair entrance door width (minimum): Series UA, UB, UC................... 26.75 inches Series UE .................................. 26.75 inches Airstair entrance door height (minimum): Series UA, UB, UC........................ 51 inches Series UE ....................................... 62 inches Cargo door width (minimum): Series UA, UB, UC........................ 52 inches Series UE ....................................... 52 inches

Aft baggage compartment—Aft Section: Series UA, UB, UC..................... 565 pounds Series UE .................................... 630 pounds 1900C.......................................... 630 pounds

Specific Loadings Wing loading (pounds per square foot): Series UA, UB, UC.................... 54.8 pounds Series UE ................................... 54.7 pounds Power loading (pounds per shaft horsepower): Series UA, UB, UC...................... 7.5 pounds Series UE ................................... 6.62 pounds

Loading Wing loading at gross weight .............. 54.7 pounds/square feet Power loading at gross weight ......... 6.6 pounds per horsepower

Cargo door height forward (minimum): Series UA, UB, UC........................ 52 inches Series UE ....................................... 57 inches

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Baggage Compartment Volumes Under seat stowage......................... 32 cubic feet (capacity 190 pounds) Forward cabin baggage .................. 17 cubic feet (capacity 250 pounds) Aft cabin baggage......................... 175 cubic feet (capacity 1,630 pounds) TOTAL ................................ 213.3 cubic feet (capacity 2,340 pounds)

Cabin Compartment Volumes Crew station.................................. 103 cubic feet Forward entrance and baggage area ............................ 56 cubic feet Main cabin.................................... 584 cubic feet Rear baggage area ........................ 175 cubic feet TOTAL ................................... 918 cubic feet

Pressurization 5.1 psi pressure differential provides a sea level cabin up to 11,000 feet and a 9,000-foot cabin at 25,000 feet (Figures 1-6 and 1-7).

Oxygen System High-pressure continuous flow with 152-cubicfoot capacity with outlets for all cabin occupants.

Figure 1-6 Rear Three-Quarters View

1-12

Figure 1-7 Cabin Pressure Vessel

Cabin Equipment Dual bleed-air heat system 64,000 BTU capacity, with environmental air provided through floor registers and individual air outlets for all occupants.

Cooling—39,000 BTU (ground) and 46,000 BTU (flight)—is provided by a combination of air cycle and vapor cycle systems.

All-Weather Equipment ●

Flush fuel vents

●

Engine fuel heaters

●

Engine inlet screen anti-ice protection system

●

Exhaust heated engine inlet lips

●

Fuel vent heaters

●

Electric propeller deicing, wing ice lights

●

Radar

●

Windshield, electric heat and surface deicer boots

●

Brake deice (optional)

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Fuel and Oil Capacity Main tank capacity............................ 481 gallons Auxiliary tank capacity..................... 184 gallons Total fuel capacity (wing tanks, usable) .......................... 665 gallons Oil capacity (each engine) ................. 3.6 gallons

Fuel and Oil Specifications Fuel .......... JP-4, JP-5, JP-8, Jet A, Jet A-1, Jet B Hydraulics (brakes, struts, and landing gear retract system...... MIL-H-5606

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Series UA, UB, UC .................

Series UE.................................

Retraction: Series UA, UB, UC .................

Series UE.................................

180 KIAS 180 KIAS 180 KIAS 180 KIAS

Maximum flap extension speed (VFE): Takeoff: Series UA, UB, UC ................. 198 KIAS Series UE.......................................... N/A Approach: Series UA, UB, UC ................. 168 KIAS Series UE............... 188 KIAS (flaps 17°) Landing: Series UA, UB, UC ................. 153 KIAS Series UE............... 143 KIAS (flaps 35°) (154 KIAS for UE-79 and after) Stall (100% flaps, power off): Series UA, UB, UC......................... 88 KIAS Series UE ........................................ 84 KIAS

Landing Gear Wheels and brakes ................ Lorol multiple disc (each main wheel) Main wheel tire size........................ 22 x 6.75-10 (tubeless 10-ply rating) Nosewheel tire size ...................... 19.5 x 6.75-10 (tubeless 10-ply rating)

Operating Speeds The 1900 Airliner qualifies as one of the most maneuverable airliners of its size in the world.

Insistence on handling ease in all flight regimes and tough construction techniques contribute to the following figures (calculated at maximum takeoff weight of 16,600 pounds for UA, UB, UC Series and 16,950 pounds for UE Series):

Air minimum control speed (VMCA): Flaps up: Series UA, UB, UC ................... 96 KIAS Series UE................................... 92 KIAS Flaps takeoff: Series UA, UB, UC .................... 91KIAS Series UE.......................................... N/A Flaps approach: Series UA, UB, UC ................... 89 KIAS Series UE................. 92 KIAS (flaps 17°)

Rates of Climb The 1900 Airliner delivers an extra margin of reliability with its powerful PT6A series jetprop engines. The following figures are calculated at maximum takeoff weight:

Maximum operating speed (VMO): Series UA, UB, UC ....................... 247 KIAS Series UE................................ 17,120 pounds To 13,200 feet.......................... 248 KIAS 13,200 to 25,000 feet.... 248 to 195 KIAS

Two engines (sea level, standard day)...................... 2,350 fpm

Maneuvering speed (VA) (16,600 pounds): Series UA, UB, UC....................... 188 KIAS Series UE ...................................... 178 KIAS

One engine (5,000 feet elevation, standard day) ....... 390 fpm

One engine (sea level, standard day)......................... 500 fpm

Maximum landing gear operating speed (VLO): Extension: Revision 1

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Service Ceiling

DOORS AND INTERIORS

At maximum takeoff weight, over-the-weather capabilities and greater mission dependability are possible with the 1900 Airliner (Figure 1-8):

CABIN ENTRY AND EXITS

Two-engine maximum operating altitude............................... 25,000 feet Cabin pressurization: Series UA, UB, UC ............................ 4.8 psi Series UE............................................ 5.1 psi

A fixed-step cabin airstair door is installed forward of the propeller plane on the left side of the fuselage (Figure 1-9). The door opening is 26.75 inches wide by 51 inches high (UC Series). The door opening is 26.21 inches wide by 62.64 inches high in the UE Series.

The 1900C is configured with a 52-inch by 52inch cargo door (52 inches wide, 57 inches high in the UE Series) aft of the passenger cabin on the left side of the airplane. The cargo door is hinged at the top and can be opened from inside or outside. A partition separates the cargo compartment from the cabin area.

The pressure vessel is sealed with inflatable rubber strips, installed around each doorframe, which allow pressurized cabin air to seep through small holes in the sides of the seal. The higher the differential pressure, the tighter the seal.

Figure 1-8 Cabin Altitude Chart

AIRSTAIR LOCKING MECHANISM

Range A typical stage length for regional airline flights is approximately 200 nautical miles. A 1900 Airliner, with 19 passengers on board, can handle three such stage lengths at maximum cruise power without refueling. You may never need the maximum range of the 1900 Airliner, but using this capability will help save time between stages by reducing turn-around time to only minutes.

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The 1900C and 1900D have three emergency exit hatches. Two are on the right side of the fuselage at the leading and trailing edges of the wing, and one is on the left side of the fuselage at the trailing edge of the wing. The 1900 configuration without a cargo door has two emergency exit doors on the right side only. The aft passenger airstair door serves as the second emergency exit on the left side.

The door-locking mechanism is operated by either of two vertically staggered handles, one inside and the other outside the door. The inside and outside handles are mechanically interconnected. When either handle is rotated, three (four on UE Series) rotating camlock latches on either side of the door capture posts mounted on the fuselage side of the opening. The door can be locked with a key to secure the aircraft when parked.

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AIRSTAIR DOOR OPERATION A release button adjacent to the door handle must be depressed before the handle can be rotated to unlock the door. As an additional safety measure, a differential-pressure-sensitive diaphragm is incorporated into the release button mechanism to prevent inadvertent opening of the door when the aircraft is pressurized.

To close the door from outside the airplane, lift the free end of the airstair door, and push it up against the doorframe as far as possible. Next, rotate the handle clockwise as far as it will go; this will allow the airstair door to move into the

PILOT TRAINING MANUAL

closed position. Then rotate the handle counterclockwise until the release button pops out; the handle should be pointing aft. Check airstair door security by attempting to rotate the handle clockwise without depressing the release button; the handle should not move.

For the UA, UB, and UC Series, to close the door from inside the airplane, grasp the handrail cable, and pull the airstair door up against the doorframe; then rotate the handle counterclockwise as far as possible, continuing to pull inward on the door. Next, rotate the handle clockwise until the release button pops out; the handle should now be pointing down.

Figure 1-9 Entrance and Exit Provisions—Series UB and UC

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For the UE Series, to close the door from the inside, grasp the rear handrail cables, and pull the airstair door up against the airframe; then rotate the handle up as far as possible, continuing to pull inward on the door. Next, rotate the handle down until the release button pops out; the handle should now be pointing down.

To check airstair door security, attempt to rotate the handle counterclockwise, or UP in the UE Series, without depressing the release button; the handle should not move. Next, ensure the red safety lock is in position around the diaphragm shaft when the handle is in the locked position.

The safety lock can be observed by depressing a red switch near the small viewing window to illuminate a lamp inside the door. The final check is to ensure that each orange stripe on each of the six (eight in the UE Series) rotary camlocks is aligned with its corresponding notch on the doorplate. If any condition specified in this doorlocking procedure is not met, DO NOT TAKE OFF. The “Cabin/Cargo Door Annunciator Circuitry Check” in the Normal Procedures section should be performed prior to the first flight of the day. In addition, a security check of the airstair door should be performed by a qualified crewmember before each flight.

Never attempt to unlock the door or even to check door security in flight. If the FWD CABIN DOOR annunciator illuminates in flight, or if the pilot has any reason to suspect the door may not be securely locked, the cabin should be depressurized slowly (considering altitude first), and all occupants should be instructed to remain seated with their seat belts fastened. After the airplane has made a full-stop landing, only a crewmember should check the security of the airstair door.

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UA, UB, UC Series To prevent entry from outside when the aircraft is secured, the exit can be locked with a flushmounted, pull-out handle near the inside exit release. When locked, with the inside lock-lever down, a red placard is visible to warn that the exit is not accessible from outside. Removing the hatch from inside with the EXIT-PULL handle is possible at any time, since the handle is not locked by the lock-lever. However, the lock-lever should be up, in the unlocked position, prior to flight.

UE Series To prevent entry from outside when the aircraft is secured, the exit can be locked by inserting a locking pin in the hole next to the release handle.

In this condition, the exit cannot be opened from the inside or outside until the pin and flag are removed.

CARGO DOOR The 1900C and D models are equipped with a cargo door to provide access for loading large or bulky items. The swing-up cargo door is hinged at the top and is equipped with latching mechanisms at the bottom (Figure 1-10).

EMERGENCY EXITS The emergency exit hatches (Figure 1-9) can be released from inside with a pull-down handle placarded “EXIT-PULL.” From outside, the doors are released with a flush-mounted pull-out handle. The nonhinged, plug-type hatches remove completely from the frame into the cabin when the latches are released.

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Figure 1-10 Cargo Door

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After initial opening force is applied, gas-operated springs open the door automatically. The door is counterbalanced to remain open, and two support rods secure the door open. Once closed, the gas springs apply a closing force to assist in latching the door. A rubber seal around the door is inflated with cabin air to seal the pressure vessel in flight.

To open the door from outside the airplane, depress the release button adjacent to the door handle, and rotate the handle clockwise. Pull out at the bottom of the door until the gas springs take over to lift it to the fully open position.

To close the cargo door from outside the airplane, pull down the free end of the cargo door, and push it against the doorframe as far as possible.

Next, rotate the handle counterclockwise until the release button pops out and the handle points aft. Check cargo door security by attempting to rotate the handle clockwise without pressing the release button; the handle should not move.

Finally, note alignment of the orange stripe and pointer by looking through a small inspection window in the lower right (lower left in UE Series) corner of the door.

Although normal access is from outside, the cargo door can also be opened or closed from inside; however, never attempt to unlock or check door security in flight. If the AFT CABIN DOOR annunciator illuminates, observe the same cautions noted in procedures for illumination of the FWD CABIN DOOR annunciator.

CABIN COMPARTMENTS The main cabin carries 19 passengers. Standard lightweight commuter seats are arranged in a single row along each cabin wall, with three passenger seats across the aft cabin partition (Figures 1-11 and 1-12). The center aisle provides easy access to all seats. Carry-on baggage space is available under each seat.

An unpressurized, 13-cubic-foot nose baggage compartment accommodates 150 pounds of baggage. Access is through an upward-opening door on the forward left side of the fuselage. This compartment has been deleted in the UE Series.

Figure 1-11 Cabin Areas

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Figure 1-12 Cabin Seating Layouts

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To accommodate additional carry-on baggage, a 15-cubic-foot baggage compartment is installed opposite the forward door, aft of the crew compartment. Its total 250-pound structural capacity includes a coat rack allowance of 100 pounds for hanging clothing.

A pressurized baggage compartment at the extreme rear of the 1900C cabin is separated from the passenger compartment by a solid bulkhead. Nylon webbing provides restraint for loose items and divides the aft compartment into two sections. Baggage or cargo space totals 154 cubic feet, distributed as 880 pounds (1,000 pounds in UE Series) forward and 630 pounds aft of the nylon webbing.

FLIGHT DECK The efficient, comfortable flight deck is arranged for convenient use by a single pilot or a two-pilot crew (Figure 1-13). Pilot and copilot sit side-byside in individual chairs, separated by a control pedestal. Seats are adjustable fore and aft, as well

PILOT TRAINING MANUAL

as vertically. Seat belts and inertial-reel shoulder harnesses are provided for each seat. Conventional dual controls allow the airplane to be flown by either pilot (Figure 1-14).

Most aircraft system circuit breakers are located on the main circuit-breaker panel at the copilot’s right side (Figure 1-15). The fuel control panel (Figure 1-16), located on the pilot’s left sidewall, includes fuel quantity gages, switches, and a single row of circuit breakers.

The instrument panel (Figures 1-17 and 1-18) contains flight instruments, engine instruments, and the avionics panel. Engine instruments (Figures 1-19, 1-20, and 1-21) are mounted in a vertical double row next to the avionics panel.

Other instrumentation is shown in Figures 1-22 and 1-23. The center avionics panel contains nav/comm controls and a weather radar unit. Two static air source selector switches are mounted outboard of the pilot’s and copilot’s flight instruments in the UA, UB, UC Series and in the lower side wall in the UE Series aircraft.

Figure 1-13 Flight Deck FOR TRAINING PURPOSES ONLY

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Figure 1-14 Control Wheels

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Figure 1-15 Right Side Panel—Series UA, UB, UC

Figure 1-16 Fuel Control Panel FOR TRAINING PURPOSES ONLY

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Figure 1-17 Instrument Panel

The power control quadrant and pedestal extends back from the center subpanel (Figure 1-24).

Controls for powerplant operation, flap system, trim tabs, yaw damp, parking brake, pressurization, and stall warning test are mounted on the pedestal. Optional systems on the lower pedestal, if installed, include flight director, autopilot, antiskid, and power steering.

Just below the instrument panel are the pilot’s subpanel on the left and the copilot’s subpanel on the right (Figures 1-25 and 1-26). Aircraft system controls, engine switches, master switch, and landing gear controls are located on these subpanels.

The UA, UB, UC Series annunciator system (Figure 1-27) consists of a warning annunciator panel (red display) in the center glareshield and a caution/advisory annunciator panel (yellow for caution, green for advisory) on the center sub-

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panel. The UE Series annunciator system consists of a warning panel (red display), a caution panel (yellow display), and an advisory panel (green and white display). Conditions covered by the annunciator system generate signals which illuminate the appropriate annunciator in a given situation.

The lighting control panel is in the overhead area, between the pilot and copilot (Figure 1-28).

Rheostat controls for flight deck and instrument lighting are mounted on this panel, convenient to both the pilot and copilot. Also mounted on this panel are the windshield wiper control, generator load and voltage gauges, and inverter monitoring gauge. Airspeed and operating limitations are also placarded on this panel. In the UE Series, all external lighting control switches have been relocated from the pilot’s right subpanel to the overhead panel.

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Figure 1-18 Instrument Panel Diagram

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Figure 1-19 Engine Instruments—Series UA, UB, UC

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Figure 1-20 Engine Instruments—Series UE-1 through UE-92

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Figure 1-21 Engine Instruments—Series UE-93 and Subsequent

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Figure 1-22 Instrumentation—Series UA, UB, UC

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Figure 1-23 Instrumentation—Series UE

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Figure 1-24 Power Control Quadrant and Pedestal

Figure 1-25 Pilot’s and Copilot’s Subpanels—Series UA, UB, UC Revision 1

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Figure 1-26 Pilot’s and Copilot’s Subpanels—Series UE

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Figure 1-27 Annunciators—Series UA, UB, UC

Figure 1-28 Lighting Controls—Series UA, UB, UC Revision 1

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CONTROL SURFACES Ailerons, rudder, and elevators are cable-operated by conventional dual control wheels. The Ttail horizontal stabilizer and elevator are mounted at the extreme top of the vertical stabilizer (Figure 1-29).

Control surfaces are cable-operated by conventional dual controls in the flight deck. Control locks (Figure 1-30) should be installed to prevent potential wind damage to controls or control surfaces whenever the airplane is parked.

Figure 1-29 T-Tail Configuration

Figure 1-30 Flight Control Locks

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TIEDOWN AND SECURING

and the propeller gearbox. Engine inlet covers provide additional protection, especially during blowing dust or rain conditions.

For overnight parking, or during high winds, tie the airplane with suitable chain or rope at the mooring points, and install protective covers (Figure 1-31). Place wheel chocks fore and aft of the main gear wheels and nosewheel. Except in severe conditions, or for temporary parking, do not set the parking brake. Be sure flaps are up and control locks are installed. Since propellers rotate freely even when the engine is shut down, they should be secured with restraints. Allowing propellers to windmill without lubrication for extended periods can damage engine bearings

Securing procedures require particular attention to use of the parking brake and the rudder gust lockpin. Before towing the airplane, the parking brake must be released (brake handle pushed in).

The rudder gust lockpin must be removed on aircraft equipped with manual steering. On aircraft equipped with power steering, removing the lock is unnecessary, since the nosewheel is free to turn when power steering is off.

Figure 1-31 Airplane Secured

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TAXIING

SERVICING DATA

Partial braking and differential power are used to achieve maximum ground turning radius during taxi (Figures 1-32 and 1-33). Use caution, however, since locking the inside brake can cause tire or strut damage. When taxiing, turning, and starting engines, be aware that propeller windstream and exhaust can be hazardous to persons or parked airplanes in areas directly behind the engines (Figure 1-34). Although velocities and temperatures cannot be accurately measured, reasonable care should be taken to prevent incidents within these danger areas.

The Handling, Servicing, and Maintenance section of the POH/AFM outlines procedures for maintaining the 1900 Airliner in its originally manufactured condition. This information sets time intervals for periodic servicing or preventive maintenance. All limits, procedures, safety practices, time limits, servicing, and maintenance requirements contained in the POH/AFM are mandatory. This section of the POH/AFM includes a Consumable Materials chart, which lists approved and recommended servicing materials. Figure 1-35 illustrates servicing points and materials required. This chart is for reference only and is always superseded by the POH/AFM.

Figure 1-32 Ground Turning Clearance—Series UA, UB, UC

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Figure 1-33 Ground Turning Clearance—Series UE

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Figure 1-34 Danger Areas

PRODUCT SUPPORT

contact the appropriate specialist to return your call.

The Beechcraft Commercial Service Department maintains a staff of service engineers to provide technical assistance whenever required.

At Beechcraft, the Commercial Service Department (CSD) is the single focal point for airline service requirements. When necessary, the CSD will involve other factory departments (Engineering, Quality Control, etc.) for assistance. Product support is truly a team effort.

Airframe, engine, avionics, and electrical specialists are just a phone call away. During normal working hours, phone (316) 676-7016 for access to our Airliner Service Group. The specialist most familiar with your particular question will be assigned to handle your request immediately.

After hours and on weekends or holidays, our service engineers may be reached through the “Service Hot-Line,� (316) 676-7111. This number will reach a Beech dispatcher, who will

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Beechcraft currently offers start-up service during the delivery process. This program provides technical assistance and spare parts when advance notice of customer requirements is furnished. Complete details are available from your sales representative.

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Figure 1-35 Servicing Data

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EXTERIOR INSPECTION 1. Cockpit Check 2. Nose Section 3. Right Wing, Landing Gear, Engine, Nacelle, and Propeller 4. Empennage and Tail

NOTE The preflight inspection has been divided into five areas as shown (Figure 1-36). The inspection procedure in the POH/AFM begins in the flight compartment, proceeds aft, then moves clockwise around the aircraft, discussing the left wing, landing gear, left engine and propeller, nose section, etc.

5. Left Wing, Landing Gear, Engine, Nacelle, and Propeller

Figure 1-36 Exterior Inspection

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CHAPTER 2 ELECTRICAL POWER SYSTEMS CONTENTS Page INTRODUCTION .................................................................................................................. 2-1 GENERAL .............................................................................................................................. 2-1 ELECTRICAL SYSTEM ....................................................................................................... 2-1 DC Power Distribution.................................................................................................... 2-7 Battery ........................................................................................................................... 2-12 Starter/Generators.......................................................................................................... 2-13 AC Power Distribution (Series UA, UB, UC)............................................................... 2-15 AC Power Distribution (Series UE) .............................................................................. 2-20 EXTERNAL POWER........................................................................................................... 2-39 AVIONICS MASTER POWER ........................................................................................... 2-40 CIRCUIT BREAKERS......................................................................................................... 2-41 ABNORMAL INDICATIONS ............................................................................................. 2-42 Generator Inoperative.................................................................................................... 2-42 Battery Charge Annunciator Illuminated ...................................................................... 2-42 Circuit Breaker Tripped................................................................................................. 2-42 Triple Fed Bus Fault...................................................................................................... 2-42

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ILLUSTRATIONS Figure

Title

Page

2-1

Electrical System Schematic Symbols..................................................................... 2-2

2-2

Electrical System Component Locations................................................................. 2-3

2-3

Pilot’s Subpanel and Overhead Panel—UA, UB, UC ............................................. 2-4

2-4

Electrical System Buses and Feeders—UA/UB ...................................................... 2-6

2-5

Right Circuit Breaker Panel—UA, UB, UC .......................................................... 2-10

2-6

Battery.................................................................................................................... 2-12

2-7

Starter Generator.................................................................................................... 2-13

2-8

Inverter Schematic ................................................................................................. 2-16

2-9

Inverter Schematic—Condition 1 .......................................................................... 2-17

2-10

Inverter Schematic—Condition 2 .......................................................................... 2-18

2-11

Inverter Schematic—Condition 3 .......................................................................... 2-19

2-12

AC Schematic—Inverters ON (Series UE) ........................................................... 2-21

2-13

AC Schematic—No. 1 Inverter Transfer (Series UE) ........................................... 2-22

2-14

AC Schematic—Load Shed (Series UE) ............................................................... 2-23

2-15

DC Electrical System Diagram.............................................................................. 2-24

2-16

DC System—Battery OFF..................................................................................... 2-25

2-17

DC System—Battery ON ...................................................................................... 2-26

2-18

DC System—Generator Ties Man Closed............................................................. 2-27

2-19

DC System—RH Engine Start............................................................................... 2-28

2-20

DC System—RH Generator ON............................................................................ 2-29

2-21

DC System—LH Engine Cross Start..................................................................... 2-30

2-22

DC System—Both Generators ON ........................................................................ 2-31

2-23

DC System—Generator Ties OPEN...................................................................... 2-32

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2-24

DC System—Bus Sense Test ................................................................................ 2-33

2-25

DC System—Both Generators Failed ................................................................... 2-34

2-26

DC System—RH Generator Bus Short ................................................................. 2-35

2-27

DC System—Center Bus Short ............................................................................. 2-36

2-28

DC System—TFB Short ....................................................................................... 2-37

2-29

DC System—External Power Applied.................................................................. 2-38

2-30

Avionics Master Power Schematic ....................................................................... 2-40

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Fuel Panel Circuit Breakers................................................................................... 2-41

2-32

Right Circuit Breaker Panel .................................................................................. 2-41

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Generator Annunciator Lights............................................................................... 2-42

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Battery Charge Annunciator Light........................................................................ 2-42

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CHAPTER 2 ELECTRICAL POWER SYSTEMS

G EN PL #1 IL O DC #1 EN G O RV M T SE TE S 1 # SY HO T T BA

T BA

FF

O

ACEN G

INTRODUCTION Understanding the airplane electrical system will ease pilot workload for normal operations and during an electrical system or component failure. The pilot should be able to locate and identify switches and circuit breakers quickly, and should also be familiar with appropriate corrective actions in emergency situations.

GENERAL The Electrical System section of the workbook presents a description and discussion of the airplane electrical system and components. The electrical system is discussed to the extent necessary for the pilot to cope with normal and emergency operations. The location and purpose of switches, indicators, lights, and circuit breakers are described. DC and AC generation and distribution is detailed. This section includes electrical system limits and descriptions of system and component faults. Figure 2-1 shows symbols used in the electrical system schematics.

ELECTRICAL SYSTEM The airplane electrical system is a 28 VDC (nominal) system with the negative lead of each power source grounded to the main airplane structure.

Direct Current (DC) electrical power is provided by one 24 VDC, 34-amp-hour nickel-cadmium battery, and two 28 VDC, 300-amp starter/generators connected in parallel. This system is capable of supplying power to all subsystems necessary for normal airplane operation. Figure 2-2 shows the location of the electrical system

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components. The Master Switch on the pilot's left subpanel controls battery and generator power to the airplane electrical system (Figure 2-3).

The battery is connected to the hot battery bus.

Both are located in the right center wing. Operation of equipment on the hot battery bus does not depend upon battery switch position. The battery switch closes a battery bus tie and a battery relay, connecting the battery to the rest of the electrical system. A circuit containing isolation diodes permits the battery relay to be energized by external power or by generator power from the center bus if battery voltage is insufficient to activate the relay.

Individual generator control units regulate output to supply constant voltage to the buses, compensating for variations in engine speed and electrical loads. The load on each generator is indicated by left and right loadmeters on the overhead meter panel. A normal system voltage of 28.25 (Âą0.25) volts maintains the battery at full charge.

The 1900 Airliner utilizes a multi-bus system.

The main buses are left and right generator buses, center bus, triple-fed bus and hot battery bus. All switches in the cockpit that receive power from the center OR triple-fed buses are identified by a white ring around the switch. Electrical loads are divided among the buses as noted on the Electrical System Buses and Feeders Chart (Figure 2-4). Equipment is arranged so that all items with duplicate functions (such as right and left landing lights) are connected to different buses.

During normal operation, all buses are automatically tied together, supplying power through individual protective devices. The hot battery bus is always connected directly to the battery. The triple-fed bus is powered by the battery and by both generator buses. Left and right generators supply power to their respective left and right generator buses. The center bus is fed by two generator buses and by the battery. All three buses automatically connect whenever the bus ties are closed. These are shown in the DC Electrical System Diagram (Figure 2-15).

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Figure 2-1 Electrical System Schematic Symbols

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Figure 2-2 Electrical System Component Locations

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Figure 2-3 Pilot’s Subpanel and Overhead Panel—UA, UB, UC (1 of 2)

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Figure 2-3 Pilot’s Subpanel and Overhead Panel—UE (2 of 2)

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Bus volts may be individually monitored on the voltmeter by selecting the appropriate bus with the VOLTMETER BUS SELECT switch located in the overhead panel (Figure 2-3).

The electrical system provides maximum protection against electrical power loss should a ground fault occur. High current (Hall Effect) sensors, bus tie relays, and current limiters are provided to isolate a fault from its power source. The electri-

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cal system bus arrangement provides multiple power sources for all circuits.

NOTE Figures 2-15 through 2-29 (Pages 2-24 through 2-38) represent electrical system power distribution schematics for various conditions. These illustrations will be referred to throughout this discussion.

Figure 2-4 Electrical System Buses and Feeders—UA/UB (1 of 3)

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DC POWER DISTRIBUTION The DC power distribution system is commonly called a “triple-fed” system since most buses receive power from three sources. For this reason, a back-up power source is available to most of the aircraft electrical systems.

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Three in-flight DC power sources are available: ●

One 24 VDC, 34-amp-hour Nickel-Cadmium Battery

●

Two 28 VDC Starter/Generators

When the battery switch is turned ON, both the battery relay and the battery bus tie relay close (Figure 2-17). Battery power is routed through the battery relay to the triple-fed bus, and through

Figure 2-4 Electrical System Buses and Feeders—UC (2 of 3)

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Figure 2-4 Electrical System Buses and Feeders—UE (3 of 3)

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the battery bus tie relay to the center bus and to both starter relays. Battery power is now available for starting either engine. Neither generator bus is yet powered since the generator bus ties are normally open until a generator is brought on line. Power can be provided to generator buses, if desired, by manually closing the GEN TIE switch. However, generator bus power is not required for engine start.

After either engine has been started, and a generator switch has been moved to RESET, the Generator Control Unit (voltage regulator) will bring the generator up to voltage. Releasing the springloaded switch to the center ON position closes the generator line contactor and both generator ties, automatically supplying power through two 250-amp current limiters (UA, UB, UC to both generator buses. Generator output will then be routed through the center bus and through the battery bus tie to permit battery charging. As soon as one generator is providing power to the electrical system, the opposite generator bus and the triple-fed bus are also powered, supplying 28 VDC power to all airplane systems (Figure 2-20).

When both generators are operating, each generator directly feeds its respective generator bus.

The generator buses, hot battery bus, and battery are tied to the center bus. The triple-fed bus receives power from the battery and from each generator bus. Three 60-amp current limiters and three diodes provide fault isolation protection for the triple-fed bus.

Bus Tie System The electrical system is protected from excessively high current flow by the bus tie system.

Three current sensors and their associated bus tie relays are electrically connected through a solidstate control circuit board. The current sensors (known as Hall Effect Devices) are located between both generator buses and the center bus, and between the battery and the center bus. The current sensors serve only to isolate a bus from the rest of the electrical system without any action by the pilot. The control circuit board also permits a functional check of the current sensor

PILOT TRAINING MANUAL

system by checking their ability to sense a ground fault and open their respective bus tie relays.

A current flow of at least 275 (Âą5) (325 UE) amps instantly activates an internal solid-state switch within the sensors, causing them to open their associated bus tie relays. Once activated, the sensors prevent the bus tie relays from closing until the circuit is reset. When a bus tie opens, the control relay will illuminate the appropriate GEN TIE or BAT TIE OPEN annunciator. After a bus tie relay is activated by a fault, it may be reset by momentarily pressing the BUS SENSE-RESET switch on the pilot's left subpanel.

Two switches on the pilot's left subpanel control the bus tie system. One switch, placarded BUS SENSE-TEST-RESET, is spring loaded to the center NORM position. Pressing the switch to TEST connects bus voltage to each current sensor test circuit (Figure 2-24), causing test current to travel directly through the Hall Effect sensing device. The solid-state sensor switch energizes the control relay and opens the bus tie relays, illuminating GEN TIE OPEN and BAT TIE OPEN annunciators. The reaction time of the sensor modules is approximately 0.010 seconds for generator current sensors and 0.120 seconds for the battery current sensor. Only momentary activation of the TEST switch is required. Voltage is continuously applied to the test circuit when the switch is held in TEST; therefore, prolonged application of test voltage will damage or destroy the sensor module.

Momentarily activating the switch to RESET deenergizes the control relays, closes the bus tie relays, and extinguishes the annunciators.

Located adjacent to the bus sense switch, the generator bus tie control switch is placarded GEN TIES-MAN CLOSE-OPEN. The GEN TIES control switch must be lifted out of the lever-locked NORM position to either OPEN or MAN CLOSE. The switch locks in OPEN, but springs back to NORM from MAN CLOSE, and manually controls only the generator bus tie relays. Manually closing the generator bus tie relays connects generator buses to the center bus,

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and permits battery power to feed the entire electrical system (Figure 2-18).

Bus Isolation

Momentarily placing the GEN TIES switch in CLOSE completes a latching circuit, closing both generator bus ties, and illuminating the MAN TIES CLOSE annunciator. The generator bus tie relays cannot be manually closed if the tie opened due to a ground fault; however, reset may be attempted with the BUS TIES-TEST-RESET switch by momentarily placing the switch to RESET.

Automatic bus isolation capability is provided by high current sensing devices (Hall Effect Devices) which protect generator and center buses. If excessive current flow is sensed, the HED opens its associated bus tie, isolating that particular bus from its power source. The two remaining buses continue to operate as a system.

Current sensors are deactivated during engine starts to prevent high current flow from opening the bus ties when the starter is energized.

When the generator ties are already closed, the GEN TIES switch can be used to open the generator bus ties. When the GEN TIES switch is positioned to OPEN, power is removed from the generator bus tie relay circuit.

Series UA, UB, UC A 250-amp current limiter (large slow-blow fuse) is located in the circuitry between the center bus and each of the generator buses. Since HED's sense high current in only one direction, current limiters are needed to provide protection for current flow in the opposite direction. If an overcurrent situation causes a current limiter to open, it also will cause bus isolation.

Figure 2-5 Right Circuit Breaker Panel—UA, UB, UC (1 of 2)

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Series UE In the UE Series, the 250-amp current limiters have been removed and the HED's in the generator circuits have been changed to bidirectional 325-amp units. If an overcurrent situation occurs, one or both of the generator ties will be opened

PILOT TRAINING MANUAL

causing bus isolation. The operation of the battery tie HED remains the same as earlier models.

Current protection for the triple-fed bus is provided exclusively by 60-amp current limiters.

Triple-fed bus isolation will occur only if all three of these limiters open (Figure 2-26).

Figure 2-5 Right Circuit Breaker Panel—UE (2 of 2)

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Load Shedding

BATTERY

The electrical system features automatic load shedding capability. When the battery is the only power source, the system automatically isolates both generator buses. If both generators are secured (or fail), both generator bus ties open, shedding all electrical loads on the generator buses (Figure 2-25). The battery will continue to power the center, triple-fed, and hot battery buses.

The 20-cell, nickel-cadmium (Ni-Cad) battery is located in the right center wing in an air-cooled box (Figure 2-6). The battery relay and charge monitor shunt are mounted in the battery compartment immediately forward of the battery.

Power to the main electrical buses is routed from the battery through the battery relay and battery bus tie, both controlled by the BAT-ON-OFF switch on the pilot's left subpanel.

If necessary, power to the generator buses can be restored by manually closing the generator ties.

However, if both generators have failed in flight, manually closing the generator bus ties will cause the battery to discharge at a much faster rate. If, for any reason, it becomes necessary to close the generator ties, they should be opened again as soon as practical to conserve battery power. Without an operating generator, the battery cannot be recharged in flight.

Some aircraft systems are powered directly from the hot battery bus, and can be operated without turning the battery switch on; however, when generators are not on line, these systems should be used sparingly to prevent excessive battery discharge.

A battery charge monitor system advises the pilot of battery charge/discharge status. A charge rate of more than 7 amps, for 6 or more seconds, will trigger the yellow BATTERY CHARGE annunciator light.

Figure 2-6 Battery

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The BATTERY CHARGE annunciator may be illuminated for short intervals when heavy electrical draw items are cycled. For example, following a battery-powered engine start, normal r a p i d c h a rg e r a t e s c a u s e t h e BAT T E RY CHARGE annunciator to be illuminated, providing automatic self-test of the battery charge monitor system. As the battery approaches its fully-charged level, and charge current decreases to a preset level, the annunciator will be extinguished. This will normally occur within a few minutes after engine start; however, longer charging times may be required if the battery has a low state of charge, low charge voltage per cell, or low temperature. On UC-59 and after, and UE-1 and after, a battery ammeter is fitted in the overhead panel.

Procedures for monitoring battery condition with the battery charge monitoring system are outlined in the POH/AFM Normal Procedures section.

PILOT TRAINING MANUAL

STARTER/GENERATORS The starter/generators are dual-purpose, enginedriven units (Figure 2-7). The same unit functions as a starter during engine starting and as a generator when supplying electrical power. A series starter winding is used during starter operation, and a shunt field winding is used during generator operation. Regulated generator output is 28.25 (Âą0.25) volts, and 300 amps maximum continuous load.

Starter power to each individual starter/generator is provided from the center bus through a starter relay. During engine starts, the battery is connected to the starter/generator by the starter relay.

With one engine running and its generator on line, the operating generator can be used to assist the battery in starting the opposite engine. This is called a generator cross-start. Normally the first engine is started on battery power alone, and the second engine is started using a generatorassisted cross-start.

Figure 2-7 Starter Generator

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The starter turns the engine compressor through accessory gearing, and is controlled by left or right engine ignition switches. Both switches, located on the pilot's left subpanel are placarded IGNITION AND ENGINE START, ON-OFFSTARTER ONLY. Actuating the switch to either position energizes the starter through the start relay. The generator control relays disable the field shunt and prevent generator operation during the start cycle.

During a cross-start (Figure 2-21) the operating Generator Control Unit will control generator output to approximately 400 amps, preventing transient surges from opening the 250-amp current limiter (UA, UB, UC). In addition, whenever a starter is selected, all bus tie sensors are deactivated to prevent bus tie relays from opening.

To prevent damage to starter motors, starter operation time limitations must be observed. Consult the POH/AFM for starter cycle limitation applicable to the model being operated.

DC Generation The generator system consists of the starter/generator units, generator control switches, generator control units (GCUs), line contactors and loadmeters.

Generator switches, labeled OFF, ON, and RESET, are located on the pilot's left subpanel next to the battery switch. The generating system is self-exciting and does not require electrical power from the aircraft electrical system for operation. The system uses generator residual voltage for initial generator buildup.

Two generator control units, mounted below the center aisle floor, regulate generator output and provide constant bus voltage during variations in engine speed and electrical load requirements.

When both generators are operating, generator control panels should balance electrical loads between the two generators within 10 percent.

Generators are manually connected to voltage regulating circuits by GEN 1 and GEN 2 control switches on the pilot's left subpanel. Each gener-

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ator's load is indicated separately on its respective loadmeter in the overhead meter panel.

The generator control units (GCU) provide the following functions: 1. Voltage regulation and line contactor control 2. Overvoltage and overexcitation protection 3. Paralleling/load sharing 4. Reverse-current protection 5. Line contactor control 6. Cross-start current limiting

Voltage Regulation and Line Contactor Control The generators are normally regulated to 28.25 (Âą0.25) VDC. When the generator control switch is held in RESET, generator residual voltage is applied through the GCU to the generator shunt field causing the generator output voltage to rise.

As generator output approaches the 28 VDC regulator setting, the voltage regulator circuit begins controlling the generator shunt field to maintain a constant output voltage. The voltage regulator circuit varies shunt field excitation, as required, to maintain a constant 28 VDC generator output for all rated conditions of generator speed, load and temperature.

Releasing the generator control switch from reset to ON applies generator voltage to the GCU, and closes the line contactor control circuit. The GCU compares generator output voltage to aircraft bus voltage. If generator output voltage has risen to within 0.5 volts of the aircraft bus, the GCU will output a signal to the line contactor which will close and connect the generator to the aircraft bus (Figure 2-20). In addition, the line contactor control signal closes both generator bus ties, connecting the center bus and the generator buses. With the bus ties closed, the generator can recharge the aircraft battery and provide power to all aircraft electrical loads.

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During single-generator operation, the GCU opens the line contactor and isolates the inoperative generator from its bus.

The GCU's cross-start current limiting circuit reduces operating generator output to protect the 250-amp current limiter (Series UA, UB, UC.

Overvoltage and Overexcitation Protection

AC POWER DISTRIBUTION (SERIES UA, UB, UC)

The GCU provides overvoltage protection to prevent excessive generator voltage from being applied to aircraft electrical equipment. If either generator output exceeds the maximum allowable 32 volts, that generator will attempt to absorb all aircraft electrical loads. The overexcitation circuits of the GCU will detect which generator is producing excessive voltage, and will disconnect that generator from the electrical system.

Two solid state inverters (Figure 2-8) supply AC power for avionics and for AC-powered engine instruments. Only one inverter is in use at a time in Series UA, UB, UC aircraft, and either inverter will provide both 115 volts, 400 Hz for avionics equipment and 26 volts, 400 Hz for applicable engine instruments and some avionics. Output of the standard inverter is 250 volt-amps. Optional inverters with higher volt-amp ratings are available.

Paralleling/Load Sharing The paralleling circuit averages the output of both generators to equalize load levels. The paralleling circuits of both GCU's become operative when both generators are brought on line. The paralleling circuits sense generator field voltages to compare the loads of both generators. The voltage regulator circuits then increase or decrease individual generator loads until both generators share the load equally. The GCU's are designed to balance loads to within 10 percent when above 25% on the load meters.

The inverters are installed in the upper aft area of each nacelle. Inverter operation is controlled by the INVERTER NO. 1-OFF-NO. 2 select switch on the pilot's left subpanel. The switch actuates an inverter power relay, supplying the selected inverter with DC power. When DC power is supplied, an inverter select relay provides the necessary switching to permit the operating inverter to supply 26 VAC avionics and instrument power, and 115 VAC avionics and test jack power. The inverter select relay is energized when the number one inverter is selected (Figure 2-9). It is deenergized when the inverter switch selects inverter number two or off.

Reverse-Current Protection If, for any reason, a generator is unable to supply at least 28.25 (Âą.25) volts to its bus, it will begin to draw current from the aircraft electrical system. The generator with higher output voltage will attempt to feed the underexcited generator (reverse current). If the GCU senses reverse current in the generator field, it will protect that generator by opening its line contactor.

Cross-Start Current Limiting During cross-start, when the starter motor is engaged, a signal from the starter control switch is applied to the GCU of the operating generator.

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Dual sources of DC input power are available to each inverter. The power select relay for each inverter automatically selects inverter power from its respective generator bus, or from the center bus if the generator bus is not powered.

When battery power is applied to the center bus prior to engine start (Figure 2-11) inverter power is available to the power relay of each inverter through the normally closed contacts of an inverter power select relay.

PILOT TRAINING MANUAL

A failure of one inverter power source will not cause loss of the selected inverter. During normal operation, the inverter power select relay is energized, supplying power from the generator bus through a circuit breaker on the copilot's circuit breaker panel. If a fault interrupts power to that bus, the power select relay will deenergize, automatically supplying inverter input power from the center bus.

Figure 2-8 Inverter Schematic

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Figure 2-9 Inverter Schematic—Condition 1

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Figure 2-10 Inverter Schematic—Condition 2

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Figure 2-11 Inverter Schematic—Condition 3

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AC POWER DISTRIBUTION (SERIES UE) The solid state inverters (Figure 2-12) supply AC power for the EFIS System, selected avionics and the Engine Torque gauges (UE-1 thru UE-92).

Both inverters are utilized at all times during normal operations.

The standard inverters are rated at 250 volt-amp, 115 volt, 400 Hz for avionics equipment and 26 volt, 400 Hz for applicable engine instruments and some avionics. Optional inverters with higher volt-amp ratings are available.

In the case of an inverter failure, action can be taken by the pilot to transfer all of the items on the failed bus to the operational inverter without loss of any operational capability (Figure 2-13).

Dual sources of DC input power are available to each inverter. The power-select relay for each inverter automatically selects inverter power from its respective generator bus, or from the center bus if the generator bus is not powered.

When battery power is applied to the center bus prior to engine start, inverter power is available to the power relay of each inverter through the normally closed contacts of the inverter powerselect relays.

A failure of one inverter power source will not cause loss of the inverter. During normal operation, the inverter power-select relay is energized, supplying power from the generator bus through a circuit breaker on the copilot's circuit breaker panel. If a fault interrupts power to that bus, the power-select relay will deenergize, automatically supplying inverter input power from the center bus.

In the case of a dual generator failure, both inverters will receive DC power from the center bus and the AC SHED buses will be activated (Figure 2-14). This reduction in AC power requirements, along with AC bus transfer action taken by the pilot, will result in reducing the DC power requirement. The AFM should be consulted for equipment lost during AC SHED bus operation.

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NICAD BATTERY CONDITION CHECK USING LOADMETER (GROUND OPERATION ONLY) Following an engine start, battery recharge current is very high causing normal illumination of the BATTERY CHARGE annunciator. The annunciator should extinguish within five minutes. If it does not extinguish, or if it should reappear, battery charge current should be monitored until the annunciator does extinguish. For those aircraft without a battery ammeter, use the following procedure to determine the battery charge current. Do not use this procedure if the Battery Charge Annunciator has illuminated in flight, but refer to the POH/AFM Abnormal Checklist for in-flight procedures.

1. One Generator—OFF 2. Voltmeter Bus Select Switch—BATT (Read 28 volts) 3. Momentarily turn the battery switch OFF, noting change in loadmeter indication.

4. The change in loadmeter indication is the battery charge current and should be no more than 2.5 percent (only perceivable needle movement). If the results are unsatisfactory, repeat the check until the charge current decreases to less than 2.5 percent.

Battery condition can also be checked on the ground before engine shutdown. With the volt select switch in BATT, the voltmeter should indicate 28 volts. Momentarily turn the battery switch off, noting the change in loadmeter indication. If the change value exceeds 2.5 percent, allow the battery to charge, repeating the check every 90 seconds. The change in loadmeter indication (the amount of needle deflection) is directly proportional to battery charging current.

A change value of 2.5 percent is indicated by very little needle movement, since full-scale deflection represents a relative load value of 100 percent. If the change value is not less then 2.5 percent within three minutes, remove the battery and have it checked by a qualified nickel cadmium battery shop.

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Figure 2-12 AC Schematic—Inverters ON (Series UE)

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Figure 2-13 AC Schematic—No. 1 Inverter Transfer (Series UE)

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Figure 2-14 AC Schematic—Load Shed (Series UE)

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The following figures represent various configurations of the electrical system.

Figure 2-15 DC Electrical System Diagram

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Figure 2-16 DC System—Battery OFF

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Figure 2-17 DC System—Battery ON

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Figure 2-18 DC System—Generator Ties Man Closed

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Figure 2-19 DC System—RH Engine Start

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Figure 2-20 DC System—RH Generator ON

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Figure 2-21 DC System—LH Engine Cross Start

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Figure 2-22 DC System—Both Generators ON

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Figure 2-23 DC System—Generator Ties OPEN

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Figure 2-24 DC System—Bus Sense Test

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Figure 2-25 DC System—Both Generators Failed

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Figure 2-26 DC System—RH Generator Bus Short

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Figure 2-27 DC System—Center Bus Short

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Figure 2-28 DC System—TFB Short

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Figure 2-29 DC System—External Power Applied

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EXTERNAL POWER External power can be supplied to the aircraft electrical system through an external power receptacle, located under the aft portion of the left nacelle. When external power is connected, a relay in the external power sensor will close only if correct polarity is sensed (Figure 2-29).

Whenever an external power plug is connected to the receptacle the green (white - UE) EXTERNAL POWER annunciator will illuminate, whether or not the external power unit is ON. The external power switch on the pilot's left subpanel closes the external power relay. A highvoltage sensor prevents the external power relay from closing if external power is above 32 (Âą0.5) volts DC. When the switch (placarded EXT PWR - OFF) is moved to EXT PWR, external power enters the aircraft electrical system, and closes both generator bus tie relays. The entire electrical system, including the starters, is then powered.

When the external power plug is connected to the aircraft, voltage can be monitored by rotating the VOLTMETER SELECT switch in the overhead panel to the EXT PWR position. External power voltage can be monitored with the external power switch in any position.

The external power source should be capable of delivering adequate power for aircraft starts.

Using an inadequate ground power unit can cause voltage drop, which may cause the starter to intermittently drop off-line resulting in relay chatter and possible welded contacts. Prior to attempting an external power start, aircraft electrical loads should be reduced to the minimum level practical.

Observe the following precautions when using an external power source: 1. When an external power source is used, ascertain that the auxiliary power unit has a minimum capacity of 1000 amps (intermittent) and 300 amps (continuous) output at 28.0 to 28.4 VDC. Never connect an external power source to the airplane unless the air-

PILOT TRAINING MANUAL

craft battery indicates a charge of at least 20 volts.

2. Use only an auxiliary power source that has a negative ground. If the polarity of the power source is unknown, use a voltmeter to determine polarity before connecting the unit to the airplane. If the ground power unit is not fitted with a standard AN-type plug, the polarity of the plug must be checked. The positive lead from the ground power unit must be connected to the center post of the external power receptacle; the negative lead must be connected to the front post, and a positive voltage of 24 to 28 VDC must be applied to the small polarizing pin.

3. Be sure the external power unit is turned off before connecting it to the aircraft. In addition, all radio equipment, the battery switch, the external power switch and the generator switches should be off. Generators should remain off until auxiliary power has been disconnected.

Because avionics master power relays must be energized to remove power from avionics equipment, the battery should be on before external power is used for engine starts. In addition to protecting avionics from transient electrical surges, the battery will provide a backup source of power during the start cycle. If the battery is removed from the airplane, or if the battery switch is off, a battery should be connected in parallel to the external power unit prior to energizing ground power.

4. If battery voltage indicates less than 20 volts, the battery must be recharged or replaced before using external power.

5. External power voltage must be regulated to 28.0 to 28.4 volts before it is plugged in to the external power receptacle. Voltages higher than 30 volts over extended time periods can damage the battery. The overvoltage sensor will prevent external power from feeding the electrical system if voltage is greater than 32 (Âą.5) volts.

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The following list outlines basic procedures for using external power for engine starts. These procedures are detailed further in the Normal Procedures section of the POH/AFM. 1. Turn the battery switch ON. 2. Connect external power; check EXT PWR annunciator illuminated.

3. Turn the external power unit ON; check voltmeter reading in EXT PWR position.

4. Turn the external power switch ON. 5. Monitor external power voltage on the overhead panel voltmeter.

6. Observe BAT TIE OPEN and GEN TIES OPEN annunciators extinguished. Use the voltmeter selector switch to check that external power is supplied to all buses.

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AVIONICS MASTER POWER Each navigation and communication radio unit has its own on-off switch; however, for pilot convenience and to prevent wear on individual switches, an avionics master power switch is installed on the pilot's left subpanel.

The switch is powered through the Avionics Master circuit breaker on the copilot's circuit breaker panel (Figure 7-34). If all avionics drop off-line, but the circuit breaker does not trip, the trouble could be in the Avionics Master switch.

Radios can be returned to service by pulling the Avionics Master circuit breaker. An Avionics Master Power schematic is shown in (Figure 2-30).

Avionics are available in a variety of packages specific to each airplane installation. Refer to vendor supplements and to the Avionics Section of this workbook for details of the avionics system.

Figure 2-30 Avionics Master Power Schematic

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CIRCUIT BREAKERS Both AC and DC power are distributed to aircraft systems through two separate circuit breaker panels which protect most components in the airplane. Each circuit breaker is stamped with its amperage rating. The small circuit breaker panel, below the fuel management gages, contains circuit breakers for the fuel system (Figure 2-31).

The main circuit breaker panel is located to the right of the copilot (Figure 2-32).

PILOT TRAINING MANUAL

The copilot’s circuit breaker panel contains breakers for major engine-related systems, all avionics components, environmental system, lights, annunciator warning systems, and some low current draw systems. Circuit breakers for the electrical distribution system are also on this panel.

Detailed procedures for tripped circuit breakers, and other electrical system malfunctions are found in the POH Abnormal Procedures section.

Figure 2-31 Fuel Panel Circuit Breakers

Figure 2-32 Right Circuit Breaker Panel

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ABNORMAL INDICATIONS

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BATTERY CHARGE ANNUNCIATOR ILLUMINATED

Electrical fires are covered in the POH/AFM Emergency Procedures section. Abnormal electrical system situations are described in the POH/AFM Abnormal Procedures section. For all authorized procedures refer to the POH/AFM. However, the following descriptions illustrate some representative situations.

I n - fl i g h t i l l u m i n a t i o n o f t h e BAT T E RY CHARGE annunciator indicates an above-normal charge current and a possible battery malfunction. See POH/AFM Normal Procedures for complete battery condition check procedures.

An amperage of 7 amps or more for six seconds will activate this light.

GENERATOR INOPERATIVE If an L DC GEN or R DC GEN caution annunciator illuminates flight, turn that generator OFF. Wait one second, then move the switch to RESET for one second, then ON. If the generator will not reset, turn it off and rely on the other operating generator. Monitor the loadmeter to ensure generator loads remain below 100 percent.

Figure 2-34 Battery Charge Annunciator Light

CIRCUIT BREAKER TRIPPED If a system draws excessive current through its circuit breaker, the circuit breaker will trip.

Resetting a tripped circuit breaker can cause further damage to a component or system. If a nonessential circuit breaker on either circuit breaker panel trips during flight, do not reset it. If the tripped breaker protects an essential system, such as avionics, wait a few seconds and reset it. If it fails to reset, DO NOT attempt to reset it again.

TRIPLE FED BUS FAULT Figure 2-33 Generator Annunciator Lights

If either loadmeter exceeds the 100 percent maximum limit, turn the battery switch off and monitor the loadmeters. If either loadmeter continues to indicate more than 100 percent, turn off all nonessential electrical equipment. If the readings then fall below 100 percent, turn the battery switch back on. Continue to monitor the loadmeters for the remainder of the flight.

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There is no annunciation for a triple fed bus fault, but can be determined by using the voltmeter selector switch. If a triple-fed bus fault exists, the voltmeter will indicate zero volts when TPL BUS is selected with the volt select switch. Reset capability is not provided for this situation, however, the pilot can determine inoperative components by referring to the Power Distribution Schematic.

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PILOT TRAINING MANUAL

CHAPTER 3 LIGHTING CONTENTS Page INTRODUCTION .................................................................................................................. 3-1 DESCRIPTION....................................................................................................................... 3-1 COCKPIT LIGHTING ........................................................................................................... 3-3 CABIN LIGHTING ................................................................................................................ 3-3 EXTERIOR LIGHTING......................................................................................................... 3-3 Series UA, UB, and UC................................................................................................... 3-3 Series UE......................................................................................................................... 3-3

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PILOT TRAINING MANUAL

ILLUSTRATIONS Figure

Title

Page

3-1

Overhead Light Control Panel (Series UA, UB, and UC) ....................................... 3-2

3-2

Exterior Lights Group (Series UA, UB, and UC).................................................... 3-2

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CHAPTER 3 LIGHTING

EXIT INTRODUCTION The Beech 1900 Airliner lighting system consists of cockpit-controlled interior and exterior lights. Interior lights are in the cockpit and passenger cabin. Exterior lighting consists of navigation lights, entry and exit threshold lights, and cargo area lights.

DESCRIPTION The lighting system contains independently operated circuits that light the following areas of the aircraft: ●

Cockpit

●

Passenger compartment

●

Entrance and exit areas

●

Cargo area

●

Aircraft exterior

Interior lights in the cockpit illuminate the flight instruments. Reading lights and exit lights are in the passenger cabin. Exterior entrance, exit, and cargo area lights illuminate the airstair and cargo area. The exterior lights improve aircraft visibility, both on the ground and in flight, to enhance traffic avoidance.

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PILOT MIC OXYGEN MASK

CO-PILOT MIC OXYGEN MASK DO NOT OPERATE ON DRY GLASS WINDSHEILD WIPER OFF PARK SLOW

NORMAL

NORMAL

FAST

MASTER PANEL LIGHTS ON

PILOT FLIGHT LIGHTS

ENGINE INSTRUMENT LIGHTS

BRT

OFF

BRT

AVIONICS PANEL LIGHTS

OFF

BRT

INCR

INCR

OVERHEAD FLOOD LIGHTS OFF

BRT

INCR

OFF

INSTRUMENT INDIRECT LIGHTS BRT

INCR

EDGELIGHT PANEL LIGHTS

OFF

BRT

INCR

OFF

OVERHEAD SIDEPANEL SUBPANEL & CONSOLE IND LIGHTS OFF

BRT

CO-PILOT FLIGHT LIGHTS BRT

INCR

INCR

OFF

INCR

OFF

CABIN LIGHTS

EXTERIOR LIGHTS LANDING

TAXI

ICE

ANTI COLLISION BEACON STROBE FLT

NAV

RECOG

CABIN FULL

TAIL FLOOD

OFF

NORM

OFF MAXIMUM AIRSPEEDS – KIAS

THIS AIRPLANE MUST BE OPERATED AS A COMMUTER CATEGORY AIRPLANE IN COMPLIANCE WITH THE OPERATING LIMITATIONS STATED IN THE FORM OF PLACARDS MARKINGS AND MANUALS. NO ACROBATIC MANUEVERS INCLUDING SPINS ARE APPROVED. THIS AIRPLANE APPROVED FOR VFR, IFR, DAY & NIGHT OPERATION & IN ICING CONDITIONS.

GEAR EXTENSION

180

17 DEGR FLAPS

188

GEAR RETRACT

180

35 DEGR FLAPS

154

CAUTION

GEAR EXTENDED

180

MANEUVERING

178

STALL WARNING IS INOPERATIVE WHEN MASTER SWITCH IS OFF. STANDBY COMPASS IS ERRATIC WHEN WINDSHIELD ANTI-ICE OR SOME COMBINATION OF EXTERIOR LIGHTS ARE ON. SEE AFM.

L GEN

20

EMERGENCY EXIT TEST ON

RIGHT

OPERATIONS LIMITATIONS

0

FSB ON ARM

PARTIAL

GND LEFT

READING ON

40

60

20

80

DC % LOAD 100

0

40

60

DC 80

DC % LOAD 100

30 20 VOLT 10

60 + BATT 0 AMP – 60

CTR BUS EXT PWR

TPL BUS BATT

LH INV

FREQ 390 400 410 380 110 120 100 AC VOLTS

PROP AMPS 420 130

PUSH FOR VOLTS

RH INV

VOLT

SERIES UA, UB AND UC

R GEN

40 30 20 10 0

40 30 20 10 0

INSTRUMENT

OFF EMERGENCY LIGHTS

SELECT

SERIES UE

Figure 3-1 Overhead Light Control Panel (Series UA, UB ,UC AND UE)

Figure 3-2 Exterior Lights Group (Series UA, UB, and UC)

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COCKPIT LIGHTING An overhead light control panel, easily accessible to both pilot and copilot, controls all cockpit lighting systems (Figure 3-1). Each light group has its own rheostat switch placarded BRT – OFF. Although each light group has its own rheostat to control light intensity, the master panel light switch, placarded ON – OFF, can be used to simultaneously control the following: pilot and copilot flight lights, subpanel, sidepanel, overhead flood, avionics panel and instrument indirect.

CABIN LIGHTING Cabin lighting is controlled by three switches in the overhead panel placarded CABIN LIGHTS. Cabin incandescent lighting is controlled by a three-position switch placarded CABIN – FULL – PARTIAL – OFF. In the partial position, only four of the cabin lights illuminate (this switch position receives power from the HOT BATT BUS).

Passenger reading lights are controlled by a twoposition switch placarded READING – ON – OFF. When the reading light switch is on, individual reading lights can be controlled by each passenger. However, if necessary, the pilot can use the switch to extinguish all passenger lights from the cockpit. A switch to the right of the reading light switch activates the NO SMOKI N G / FA S T E N S E AT B E LT s i g n s a n d accompanying chimes. This three-position switch is placarded NO SMOKE & FSB – FSB – OFF (UA, UB, UC).

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Cargo compartment lights are controlled by a two-position switch just inside the cargo door at floor level. The three cargo compartment lights are also connected to the hot battery bus, and the lights will extinguish when the cargo door is closed.

EXTERIOR LIGHTING SERIES UA, UB, AND UC Switches for the landing lights, taxi lights, wing ice lights, navigation lights, recognition lights, rotating beacons, and wingtip and tail strobe lights are located on the pilot’s right subpanel (Figure 3-2). They are appropriately placarded.

Tail floodlights are incorporated into the horizontal stabilizers to illuminate both sides of the vertical stabilizer. A switch for these lights placarded LIGHTS – TAIL FLOOD – OFF, is on the pilot’s right subpanel (Figure 3-2).

SERIES UE Switches for the landing lights, taxi light, wing ice lights, navigation lights, anti-collision (beacon and strobe) lights and tail flood lights are located in the overhead center panel.

The threshold lights at the passenger entryway are controlled by a two-position switch on the side of the third step of the airstair door. The control switch is connected to the hot battery bus and is usable whether or not the battery switch is turned on. When the door is closed, a microswitch in the locking mechanism will extinguish the lights. However, care should be taken not to leave the door open with the light switch on, since it will cause the battery to be discharged.

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CHAPTER 4 MASTER WARNING SYSTEM CONTENTS Page INTRODUCTION .................................................................................................................. 4-1 GENERAL .............................................................................................................................. 4-1 ANNUNCIATOR SYSTEM................................................................................................... 4-2 Master Warning Flashers................................................................................................. 4-3 Master Caution Flashers .................................................................................................. 4-3 Caution/Advisory Panel (Series UE)............................................................................... 4-5 Dimming.......................................................................................................................... 4-5 Testing and Lamp Replacement ...................................................................................... 4-6 WARNING AND CAUTION/ADVISORY PANEL DESCRIPTIONS................................ 4-6

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ILLUSTRATIONS Figure

Title

Page

4-1

Annunciator System................................................................................................. 4-2

4-2

Master Warning and Caution Flashers .................................................................... 4-3

4-3

Warning Annunciator Panel—Series UA, UB, UC................................................. 4-3

4-4

Warning Annunciator Panel—Series UE ................................................................ 4-4

4-5

Caution/Advisory Annunciator Panel—Series UA, UB, UC .................................. 4-4

4-6

Caution/Advisory Annunciator Panel—Series UE.................................................. 4-5

4-7

Annunciator Bulb Replacement............................................................................... 4-6

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TABLES Table

Title

Page

4-1

Warning Annunciators—Series UA, UB, UC.......................................................... 4-7

4-2

Caution Annunciators—Series UA, UB, UC........................................................... 4-8

4-3

Advisory Annunciators—Series UA, UB, UC....................................................... 4-10

4-4

Warning Annunciators—Series UE ....................................................................... 4-11

4-5

Caution Annunciators—Series UE ........................................................................ 4-12

4-6

Advisory Annunciators—Series UE ...................................................................... 4-15

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CHAPTER 4 MASTER WARNING SYSTEM

TEST

INTRODUCTION Warning and caution annunciators may be the first indication of malfunction in an airplane system or component. Crewmembers should be completely familiar with annunciator indications and with pilot actions required to continue flight until a safe landing can be made. If a malfunction occurs before takeoff, as indicated by the annunciator system, potentially dangerous flight situations could be averted by correcting the problem while still in a safe ground environment.

GENERAL This chapter provides detailed information covering warning, caution, and advisory annunciator panels.

Descriptions of the system include purpose and associated cause of illumination for each

annunciator. Appropriate pilot action, if required, is outlined.

Also, test functions and procedures for the master warning system are described.

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ANNUNCIATOR SYSTEM The annunciator system (Figure 4-1) consists of a warning annunciator panel (red) in the center glareshield and a caution/advisory annunciator panel (yellow/green for UA, UB, UC; yellow/green and white for UE) located on the center subpanel.

A pair of master flashers are positioned on each side of the glareshield in front of each pilot. The outboard flasher on each side is the red MASTER

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WARNING flasher. The other is the yellow MASTER CAUTION flasher. A PRESS TO TEST switch is located immediately to the right of the warning annunciator panel.

In addition to color coding, annunciators incorporate word-readouts to facilitate interpretation of annunciator indications. If covered by the annunciator system, an aircraft system fault generates a signal which illuminates the appropriate warning light.

Figure 4-1 Annunciator System

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MASTER WARNING FLASHERS When a red warning annunciator illuminates, both MASTER WARNING flashers (Figure 4-2) begin flashing. A red warning annunciator indicates a system fault has occurred which requires immediate pilot attention and appropriate response (Figures 4-3 and 4-4). The warning annunciator will remain illuminated until the fault is corrected.

Pressing the face of either pilot’s MASTER WARNING flasher will extinguish the flasher even if the fault is not corrected. If the MASTER WARNING flashers have been canceled, they will again be activated when an additional warning annunciator illuminates. When the fault that tripped the annunciator is corrected, the affected warning annunciator will extinguish, but the MASTER WARNING annunciators will continue to flash until canceled.

MASTER CAUTION FLASHERS Whenever a system fault occurs that requires the pilot’s attention but not his immediate reaction, the appropriate yellow caution annunciator in the caution/advisory panel illuminates (Figures 4-5 and 4-6), and both MASTER CAUTION flashers (Figure 4-2) begin flashing. The flashing MASTER CAUTION lights can be extinguished by pressing the face of either MASTER CAUTION

MASTER WARNING

MASTER CAUTION

PRESS TO RESET

PRESS TO RESET

Figure 4-2 Master Warning and Caution Flashers

flasher to reset the circuit. If any caution annunciator again illuminates, the MASTER CAUTION flashers will be reactivated. A caution annunciator will remain on until the fault that tripped it is corrected. MASTER CAUTION flashers will continue to flash until canceled.

The caution/advisory annunciator panel also contains green advisory annunciators. There are no master flashers associated with these annunciators, since they are advisory only. Advisory lights indicate functional situations which do not demand immediate attention or reaction, but simply advise the pilot that a system has been armed or activated. An advisory annunciator can be extinguished only by changing the condition indicated by the illuminated green lens.

L FUEL PRESS

CABIN ALTITUDE

BAGGAGE DOOR

INVERTER

R FUEL PRESS

L OIL PRESS

L ENVIR FAIL

FWD CABIN DOOR

R ENVIR FAIL

R OIL PRESS

A/P DISC

R BL AIR FAIL

AFT CABIN DOOR L BL AIR FAIL

A/P TRIM FAIL

Figure 4-3 Warning Annunciator Panel—Series UA, UB, UC

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CABIN DIFF HI

R FUEL PRES LO

CABIN DOOR

R ENVIR FAIL

R OIL PRES LO

L AC BUS

CARGO DOOR

R AC BUS

A/P TRIM FAIL

ARM EMER LITES

A/P FAIL

L FUEL PRES LO

CABIN ALT HI

L OIL PRES LO

L ENVIR FAIL

L BL AIR FAIL

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R BL AIR FAIL

Figure 4-4 Warning Annunciator Panel—Series UE

L DC GEN

L FUEL QTY

L FW VALVE

L FUEL FEED

L GEN TIE OPEN

L ENG ICE FAIL

L BK DI OVHT

L AUTOFEATHER

L CHIP DETECT

L IGNITION ON L ENG ANTI-ICE

L BK DEICE ON

L ENVIR OFF

FUEL TRANSFER

BATTERY CHARGE BATT TIE OPEN

R FUEL QTY

R DC GEN

R GEN TIE OPEN

R FUEL FEED

R FW VALVE

HYD FLUID LOW

MAN STEER FAIL

R BK DI OVHT

R ENG ICE FAIL

ANTI SKID FAIL

PWR STEER FAIL

R CHIP DETECT

R AUTOFEATHER

TAXI LIGHT

EXTERNAL POWER

ELEC TRIM OFF

MAN TIES CLOSE

R IGNITION ON R BK DEICE ON

R ENG ANTI-ICE

AIR COND N1 LOW

R ENVIR OFF

R FUEL QTY

R DC GEN

SERIES UA, UB BATTERY CHARGE BATT TIE OPEN

L DC GEN

L FUEL QTY

L FW VALVE

L FUEL FEED

L GEN TIE OPEN

R GEN TIE OPEN

R FUEL FEED

R FW VALVE

L ENG ICE FAIL

L BK DI OVHT

HYD FLUID LOW

MAN STEER FAIL

R BK DI OVHT

R ENG ICE FAIL

ANTI SKID FAIL

PWR STEER FAIL ANN PWR SOURCE

L NO FUEL XFR

R NO FUEL XFR PWR STEER ENGA

L AUTOFEATHER

L IGNITION ON

TAXI LIGHT

EXTERNAL POWER

R IGNITION ON

R AUTOFEATHER

L ENG ANTI-ICE

L BK DEICE ON

ELEC TRIM OFF

MAN TIES CLOSE

R BK DEICE ON

R ENG ANTI-ICE

L ENVIR OFF

FUEL TRANSFER AIR COND N1 LOW

R ENVIR OFF

SERIES UC

Figure 4-5 Caution/Advisory Annunciator Panel—Series UA, UB, UC

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CAUTION/ADVISORY PANEL (SERIES UE)

PILOT TRAINING MANUAL

DIM mode is automatically selected whenever all of the following conditions are met:

The caution/advisory annunciator panel on Series UE airplanes also contains the green and white advisory annunciators (Figure 4-6). There are no master flashers associated with these annunciators. An advisory annunciator can be extinguished only by changing the condition indicated on the illuminated lens.

DIMMING The automatic annunciator BRIGHT and DIM functions affect the following annunciators: ●

Warning annunciators

●

Caution annunciators

●

Advisory annunciators

●

Yellow MASTER CAUTION flashers

●

Landing gear handle lights

●

Landing gear position lights

L DC GEN L FW VALVE

L FUEL QTY

STALL HEAT

●

At least one generator is on line.

●

Cockpit OVERHEAD FLOOD LIGHTS are OFF.

●

MASTER PANEL LIGHTS switch is ON.

●

PILOT FLIGHT LIGHTS are ON.

●

Ambient light level in the cockpit is below a preset value.

Cockpit ambient light level is sensed by a photoelectric cell in the overhead lighting control panel. Unless all of the above conditions are met, the BRIGHT mode will be selected automatically. The red MASTER WARNING flasher and the fire T-handles are not affected by the DIM mode; they are always set to BRIGHT.

BATTERY CHARGE PROP GND SOL

R FUEL QTY

R DC GEN

L COL TANK LOW L GEN TIE OPEN

BATT TIE OPEN

R GEN TIE OPEN R COL TANK LOW

ANTI SKID FAIL

ANN PWR SOURCE

R BK DI OVHT

R ENG ICE FAIL

PWR STEER FAIL MAN STEER FAIL

R PITOT HEAT

R FIRE LOOP

AFX DISABLE

R NO AUX XFR

R FW VALVE

L ENG ICE FAIL

L BK DI OVHT

HYD FLUID LOW

L FIRE LOOP

L PITOT HEAT

XFR VALVE FAIL

L NO AUX XFR

AUTOFTHER OFF

PITCH TRIM OFF

INBD WG DEICE

YD/RB FAIL

TAIL DEICE

L AUTOFEATHER

L IGNITION ON

PWR STEER ENGA

R IGNITION ON

R AUTOFEATHER

L ENG ANTI-ICE

L BK DEICE ON

MAN TIES CLOSE

R BK DEICE ON

R ENG ANTI-ICE

L ENVIR OFF

RDR PWR ON

TAXI LIGHT

EXTERNAL POWER

R ENVIR OFF

FUEL TRANSFER

RUD BOOST OFF OUTBD WG DEICE

Figure 4-6 Caution/Advisory Annunciator Panel—Series UE

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TESTING AND LAMP REPLACEMENT The lamps in the annunciator system should be tested before every flight, and any time the integrity of a lamp is in question. Depressing the PRESS TO TEST button illuminates all annunciator lights, MASTER WARNING flashers, and MASTER CAUTION flashers. Any lamp that fails to illuminate when tested should be replaced. All annunciator lamps, including landing gear warning and position indicators, are interchangeable. Except on D models, where position lights are not pilot servicable.

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Replacement bulbs are conveniently located within the panels, and are indicated by dashed lines on the face of each annunciator which contains spare lamps. To replace any annunciator lamp (Figure 4-7), first depress the center of the annunciator with your finger. Release, and the annunciator will pop out slightly. Remove the annunciator from the panel, and pull the lamp from the back of the annunciator. Replace the failed bulb with a spare lamp, and return the annunciator to the panel, depressing it until locked back into place.

WARNING AND CAUTION/ADVISORY PANEL DESCRIPTIONS Tables 4-1 through 4-6 list all the warning, caution, and advisory annunciators on the Beech 1900 Airliner. The cause for illumination is included beside each annunciator.

Figure 4-7 Annunciator Bulb Replacement

4-6

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Table 4-1 WARNING ANNUNCIATORS—SERIES UA, UB, UC ANNUNCIATOR

L FUEL PRESS

CAUSE FOR ILLUMINATION Loss of fuel boost pressure on left side.

CABIN ALTITUDE

Cabin pressure altitude exceeds 12,500 feet.

BAGGAGE DOOR

Nose baggage door is open or not secure.

INVERTER R FUEL PRESS

The inverter selected is inoperative.

Loss of fuel boost pressure on right side.

L OIL PRESS

Loss of oil pressure in the left engine.

L ENVIR FAIL

Left environmental system bleed-air overtemp or overpressure.

FWD CABIN DOOR

Forward cabin door is open or not secure.

R ENVIR FAIL

Right environmental system bleed-air overtemp or overpressure.

R OIL PRESS

Loss of oil pressure in the right engine.

AFT CABIN DOOR L BL AIR FAIL A/P TRIM FAIL* A/P DISC* R BL AIR FAIL

Aft cabin door is open or not secure.

Melted or ruptured left bleed-air failure warning line.

Improper trim or no trim from autopilot trim command.

Autopilot is disconnected by means other than pilot’s disconnect switch.

Melted or ruptured right bleed-air failure warning line.

* Optional equipment

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Table 4-2 CAUTION ANNUNCIATORS—SERIES UA, UB, UC ANNUNCIATORS

L DC GEN L FUEL QTY BATTERY CHARGE BATT TIE OPEN R FUEL QTY R DC GEN

Left generator is off line.

Left fuel quantity—Less than 30 minutes remaining at maximum continuous power.

Excessive charge rate on battery.

Battery isolated from generator buses.

Right fuel quantity—Less than 30 minutes remaining at maximum continuous power.

Right generator is off line.

L FW VALVE

Left fuel firewall valve has not reached its selected position.

L FUEL FEED

Low fuel level in left fuel system holding tank—Less than 2 minutes remaining at maximum continuous power.

L GEN TIE OPEN

Left generator bus is isolated from center bus.

R GEN TIE OPEN

Right generator bus is isolated from center bus.

R FUEL FEED

Low fuel level in right fuel system holding tank—Less than 2 minutes remaining at maximum continuous power.

R FW VALVE

Right fuel firewall valve has not reached its selected position.

L ENG ICE FAIL

Ice vane has not attained proper position.

L BK DI OVHT*

Melted or ruptured left brake deice plumbing failure warning line.

HYD FLUID LOW

4-8

CAUSE FOR ILLUMINATION

Landing gear hydraulic fluid is low.

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Table 4-2 CAUTION ANNUNCIATORS—SERIES UA, UB, UC (Cont) ANNUNCIATORS

MAN STEER FAIL*

CAUSE FOR ILLUMINATION Nose gear will not trail with power steering not engaged.

R BK DI OVHT*

Melted or ruptured right brake deice plumbing failure warning line.

R ENG ICE FAIL

Ice vane has not attained proper position.

ANTI-SKID FAIL*

Loss of electrical or low hydraulic oil pressure in antiskid brake system.

R CHIP DETECT**

Contamination in right engine oil is detected.

L CHIP DETECT**

Contamination in left engine oil is detected.

PWR STEER FAIL*

Loss of electrical or low hydraulic oil pressure in power steering system.

ANN PWR SOURCE**

Partial power loss to some annunciator lights.

L NO FUEL XFR

No fuel transfer from left auxiliary to main tank.

R NO FUEL XFR

No fuel transfer from right auxiliary to main tank.

* Optional equipment ** UB 54 and after; UC 1 and after

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Table 4-3 ADVISORY ANNUNCIATORS—SERIES UA, UB, UC ANNUNCIATOR

PWR STEER ENGA* L AUTOFEATHER L IGNITION ON TAXI LIGHT EXTERNAL POWER R IGNITION ON R AUTOFEATHER

Power steering is operating.

Left autofeather system is armed, with the power levers advanced above 90% N1.

Left engine igniter is powered.

Taxi light is on with landing gear up.

External power connector is plugged in.

Right engine igniter is powered.

Right autofeather system is armed, with the power levers advanced above 90% N1.

L ENG ANTI-ICE

Left ice vane is extended.

L BK DEICE ON*

Left brake deice bleed-air valve is in the open position.

ELEC TRIM OFF*

Electric trim is deenergized by control wheel trim disconnect switch with pedestal power switch on.

MAN TIES CLOSE

Manually closed generator bus ties.

R BK DEICE ON*

Right brake deice bleed-air valve is in the open position.

R ENG ANTI-ICE

Right ice vane is extended.

L ENVIR OFF FUEL TRANSFER

4-10

CAUSE FOR ILLUMINATION

Left environmental bleed-air valve is closed.

Transfer valve is open.

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Table 4-3 ADVISORY ANNUNCIATORS—SERIES UA, UB, UC (Cont) ANNUNCIATOR

AIR COND N1 LOW R ENVIR OFF

CAUSE FOR ILLUMINATION Right engine rpm is too low for air conditioner to engage.

Right environmental bleed-air valve is closed.

* Optional equipment

Table 4-4 WARNING ANNUNCIATORS—SERIES UE ANNUNCIATOR

L FUEL PRES LO

CAUSE FOR ILLUMINATION Fuel pressure is low on left side.

CABIN ALT HI

Cabin altitude exceeds 10,000 feet.

CAB DIFF HI

Cabin pressure differential is high.

R FUEL PRES LO

Fuel pressure is low on right side.

L OIL PRES LO

Oil pressure failure in left engine.

L ENVIR FAIL

Left environmental air duct overtemp or overpressure.

CABIN DOOR

Cabin door is open or not secure.

R ENVIR FAIL

Right environmental air duct overtemp or overpressure.

R OIL PRES LO L AC BUS

Oil pressure failure in right engine.

Left AC bus has inoperative inverter.

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Table 4-4 WARNING ANNUNCIATORS—SERIES UE (Cont) ANNUNCIATOR

CARGO DOOR R AC BUS L BL AIR FAIL A/P TRIM FAIL* ARM EMER LITES* A/P FAIL* R BL AIR FAIL

CAUSE FOR ILLUMINATION Cargo door is open or not secure.

Right AC bus has inoperative inverter.

Melted or failed left bleed-air failure warning line, or system is off.

Improper trim or no trim from autopilot trim command.

Emergency light controls are disarmed.

A failure has occurred in the selected APC-65 computer.

Melted or failed right bleed-air failure warning line, or system is off.

* Optional equipment

Table 4-5 CAUTION ANNUNCIATORS—SERIES UE ANNUNCIATORS

L DC GEN

Left generator is off line.

L FUEL QTY

Fuel quantity is below 324 pounds of usable fuel.

STALL HEAT

Insufficient current to provide heat on stall warning transducer to prevent icing.

BATTERY CHARGE PROP GND SOL

4-12

CAUSE FOR ILLUMINATION

Excessive charge rate on battery.

One or both ground idle low-pitch-stop solenoids are malfunctioning.

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Table 4-5 CAUTION ANNUNCIATORS—SERIES UE (Cont) ANNUNCIATORS

R FUEL QTY R DC GEN L FW VALVE L COL TANK LOW L GEN TIE OPEN BAT TIE OPEN R GEN TIE OPEN

CAUSE FOR ILLUMINATION Fuel quantity is below 324 pounds of usable fuel.

Right generator is off line.

Left fuel firewall valve has not reached its selected position.

Left fuel system collector tank is below 53 pounds of usable fuel (eight minutes cruise at 400 pounds per hour).

Left generator bus is isolated from the center bus.

Battery is isolated from the generator buses.

Right generator bus is isolated from the center bus.

R COL TANK LOW

Right fuel system collector tank is below 53 pounds of usable fuel (eight minutes cruise at 400 pounds per hour).

R FW VALVE

Right fuel firewall valve has not reached its selected position.

L ENG ICE FAIL

Left ice vane malfunction. Ice vane has not attained the proper position.

L BK DI OVHT*

Melted or failed left brake deice plumbing failure warning line.

HYD FLUID LOW

Landing gear hydraulic fluid level is low.

ANTI SKID FAIL*

Electrical failure or low hydraulic oil pressure in the antiskid brake system.

ANN PWR SOURCE R BK DI OVHT*

Partial power loss to some annunciator lights.

Melted or failed right brake deice plumbing failure warning line.

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Table 4-5 CAUTION ANNUNCIATORS—SERIES UE (Cont) ANNUNCIATORS

CAUSE FOR ILLUMINATION

R ENG ICE FAIL

Right ice vane malfunction. Ice vane has not attained the proper position.

L FIRE LOOP

4-14

Left engine fire detection sense loop is open.

L PITOT HEAT

Insufficient current to provide heat on left pitot to prevent icing.

XFR VALVE FAIL

Fuel cross-transfer valve is not fully open or fully closed for two seconds or more.

PWR STEER FAIL*

Electrical failure or low hydraulic oil pressure in power steering system.

MAN STEER FAIL*

Nose gear will not free-caster with power steering not engaged.

R PITOT HEAT

Insufficient current to provide heat on right pitot to prevent icing.

R FIRE LOOP

Right engine fire detector sense loop is open.

L NO AUX XFR

No fuel transfer from left auxiliary to main tank.

AUTOFTHER OFF

Autofeather system is turned off with landing gear extended.

PITCH TRIM OFF*

Electric trim is deenergized by a trim disconnect switch on the control wheel, with the system power switch on the pedestal turned on.

AFX DISABLE

Autofeather system is not capable of feathering the propellers.

R NO AUX XFR

No fuel transfer from right auxiliary to main tank.

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Table 4-5 CAUTION ANNUNCIATORS—SERIES UE (Cont) ANNUNCIATORS

YD/RB FAIL** RUD BOOST OFF

CAUSE FOR ILLUMINATION A failure has occurred in the selected FYD-65 computer.

Rudder boost system is turned off.

* Optional equipment ** On airplanes without an autopilot

Table 4-6 ADVISORY ANNUNCIATORS—SERIES UE ANNUNCIATOR

CAUSE FOR ILLUMINATION

INBD WG DEICE

Pressure in left and right inboard wing deice boots is sufficient to deice.

TAIL DEICE

Pressure in tail deice boots is sufficient to deice.

OUTBD WG DEICE

Pressure in left and right outboard wing deice boots is sufficient to deice.

L AUTOFEATHER

Left autofeather system is armed, with the power levers advanced above approximately 89 to 91% N1.

L IGNITION ON PWR STEER ENGA R IGNITION ON R AUTOFEATHER

Left engine igniter is powered.

Power steering is operating.

Right engine igniter is powered.

Right autofeather system is armed, with the power levers advanced above approximately 89 to 91% N1.

L ENG ANTI-ICE

Left ice vane is extended.

L BK DEICE ON*

Left brake deice bleed-air valve is in the open position.

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Table 4-6 ADVISORY ANNUNCIATORS—SERIES UE (Cont) ANNUNCIATOR

MAN TIES CLOSE

CAUSE FOR ILLUMINATION Manually closed generator bus ties.

R BK DEICE ON*

Right brake deice bleed-air valve is in the open position.

R ENG ANTI-ICE

Right ice vane is extended.

L ENVIR OFF RDR PWR ON* FUEL TRANSFER TAXI LIGHT EXTERNAL POWER R ENVIR OFF

Left environmental bleed-air valves are closed.

Radar is selected to a position other than off (on ground).

Fuel cross-transfer valve is open.

Taxi light is on with landing gear up.

External power is plugged into aircraft.

Right environmental bleed-air valves are closed.

* Optional equipment

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CHAPTER 5 FUEL SYSTEM CONTENTS Page INTRODUCTION .................................................................................................................. 5-1 DESCRIPTION....................................................................................................................... 5-1 FUEL SYSTEM (SERIES UA/UB) ....................................................................................... 5-2 Main Fuel Tank System .................................................................................................. 5-2 Fuel Tank Vents .............................................................................................................. 5-5 Fuel System Operation .................................................................................................... 5-6 Firewall Shutoff Valves................................................................................................... 5-8 Fuel Transfer Motive-Flow System ................................................................................ 5-8 Cross-Transfer Operation ............................................................................................... 5-8 Fuel Purge System......................................................................................................... 5-10 Fuel Gaging System ...................................................................................................... 5-10 Fuel Drains .................................................................................................................... 5-12 FUEL SYSTEM (SERIES UC/UE) ...................................................................................... 5-13 Fuel Tank System.......................................................................................................... 5-13 Fuel Capacity................................................................................................................. 5-17 Fuel Tank Vents ............................................................................................................ 5-17 Fuel System Operation .................................................................................................. 5-18 Low-Fuel Warning System (Series UC) ....................................................................... 5-19 Low-Fuel Warning System (Series UE)........................................................................ 5-19 Firewall Shutoff Valves................................................................................................. 5-19 Fuel Transfer Motive-Flow System .............................................................................. 5-20

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Cross-Transfer Operation.............................................................................................. 5-20 Fuel Purge System ........................................................................................................ 5-21 Fuel Control Panel ........................................................................................................ 5-22 Fuel Gaging System...................................................................................................... 5-22 Standby Pump Operation .............................................................................................. 5-22 Auxiliary Tank Operation ............................................................................................. 5-23 Visual Fuel Quantity Sensors (Series UE).................................................................... 5-23 Fuel Drains.................................................................................................................... 5-24 Fuel Drain Locations..................................................................................................... 5-24 Fuel Handling Practices (Series UA/UB and UC/UE) ................................................. 5-25 Fuel Grades and Types (Series UA/UB and UC/UE) ................................................... 5-27 Filling the Tanks (Series UA/UB and UC/UE)............................................................. 5-27 Draining the Fuel System (Series UA/UB and UC/UE)............................................... 5-28

5-ii

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ILLUSTRATIONS Figure

Title

Page

5-1

Main Fuel Tank System—Series UA/UB................................................................ 5-2

5-2

Fuel System Schematic Diagram—Series UA/UB Only......................................... 5-3

5-3

Fuel Control Panel—Series UA/UB ........................................................................ 5-4

5-4

Fuel Vent System—Series UA/UB ......................................................................... 5-5

5-5

Fuel Flow Diagram—Series UA/UB....................................................................... 5-6

5-7

FIRE PULL Handles—Series UA/UB .................................................................... 5-8

5-6

Firewall Shutoff Valve—Series UA/UB ................................................................. 5-8

5-8

Fuel Transfer Motive-Flow System—Series UA/UB.............................................. 5-9

5-9

Cross-Transfer System—Series UA/UB ................................................................. 5-9

5-10

Fuel Drain Purge System Schematic—Series UA/UB .......................................... 5-10

5-11

Fuel Probe—Series UA/UB................................................................................... 5-11

5-12

Fuel Gaging System—Series UA/UB ................................................................... 5-11

5-13

Fuel Drains—Series UA/UB ................................................................................. 5-12

5-14

Fuel Tank System—Series UC/UE ....................................................................... 5-13

5-15

Fuel System Diagram—Series UC and After ........................................................ 5-14

5-16

Fuel System Schematic—Series UE...................................................................... 5-15

5-17

Fuel Control Panel—Series UC/UE ...................................................................... 5-16

5-18

Fuel Tank Vent System—Series UC/UE............................................................... 5-16

5-19

Fuel Flow Diagram—Series UC/UE ..................................................................... 5-18

5-20

Firewall Shutoff Valve—Series UC/UE................................................................ 5-19

5-21

FIRE PULL Handles—Series UC/UE................................................................... 5-20

5-22

Cross-Transfer Schematic—Series UC/UE ........................................................... 5-21

5-23

Fuel Drain Purge System Schematic—Series UC/UE........................................... 5-21

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5-24

Fuel Control Panel—Series UE ............................................................................ 5-22

5-25

Fuel Drains—Series UC/UE ................................................................................. 5-24

5-26

Fuel Temperature Graph ....................................................................................... 5-26

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TABLES Table

Title

Page

5-1

Usable Fuel ............................................................................................................ 5-13

5-2

Fuel Drain Locations ............................................................................................. 5-24

FOR TRAINING PURPOSES ONLY

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CHAPTER 5 FUEL SYSTEM

;;; ;;; ;;; ;;;; ;;; ;;;; ;;; ;;;

;;; 6

4

MAIN FUEL

2

8

LBS X 100

0

10

INTRODUCTION A complete understanding of the fuel system is essential to competent and confident operation of the aircraft. Management of aircraft fuel and fuel system components is a major concern.

This section provides the pilot with information needed for safe, efficient fuel management of the UA/UB and UC/UE series of 1900 Airliners.

DESCRIPTION The Fuel System section of the workbook presents a description and discussion of the fuel system. The physical layout of the fuel system and fuel cells are described in this section.

Correct use of boost pumps, transfer pumps, cross-transfer, and firewall shutoff valves are

discussed. Locations and types of fuel drains are also described along with correct procedures for taking and inspecting fuel samples. A list of approved fuels and a discussion of the tank filling sequence are included.

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FUEL SYSTEM (SERIES UA/UB) The Beechcraft 1900 Airliner fuel system (Figure 5-2) simplifies cockpit flight procedures and provides easy servicing access on the ground. The two separate wing fuel systems, one for each engine, are connected by a valve-controlled cross-transfer system.

MAIN FUEL TANK SYSTEM The main fuel system (Figure 5-1) consists of a series of rubber bladder fuel cells and one integral (wet wing) tank in each wing. Total fuel capacity is approximately 216 gallons per side including unusable fuel. Total usable fuel capacity of the main fuel system is 212.5 gallons per side. One fuel filler cap per side is located on the outboard leading edge of each wing near the

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wing tip. The main fuel system in each wing consists of two wing leading edge tanks, two box section tanks, an integral (wet wing) tank, and one center wing tank, all interconnected to gravity-flow into a fuel supply collector tank.

The fuel supply collector tank, located within the center wing fuel tank, is equipped with an electric fuel pump, primary jet pump, and two transfer jet pumps. Fuel feeds the engines directly from the fuel supply collector tank.

The collector tank drain is located under the center wing, adjacent to the fuselage. The inboard leading edge tank drain is under the wing just outboard of the nacelle. The integral (wet wing) fuel tank sump drain is located approximately midway on the underside of the wing, aft of the main spar. The collector tanks are connected by a cross-transfer line which makes fuel available from either wing system. Cross-transfer operation is automatic when cross-transfer is selected.

Figure 5-1 Main Fuel Tank System—Series UA/UB

5-2

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Figure 5-2 Fuel System Schematic Diagram—Series UA/UB Only

BEECH 1900 AIRLINER

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5-3

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Figure 5-3 Fuel Control Panel—Series UA/UB

FUEL CAPACITY The fuel quantity indicating system is a capacitance type that compensates for differences in specific gravity and reads in pounds on a linear scale (Figure 5-3).

Each wing has an independent fuel gaging system, consisting of a fuel quantity (capacitance) probe in the collector tank, one probe in the aft inboard fuel cell, two probes in the integral (wet wing) fuel cell, two probes in

5-4

the inboard leading-edge fuel cell, and two probes in the center wing fuel cell. Electronic circuits process the signals from the capacitance probes to provide an accurate readout on fuel quantity indicators. The fuel quantity gages and the engine fuel flow indicators are read in pounds. At 6.74 pounds per gallon, total usable fuel is 425 gallons, or approximately 2864.5 pounds. Maximum allowable imbalance between left and right fuel systems is 300 pounds, and maximum zero fuel weight of the 1900 Airliner is 14,000 pounds.

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FUEL TANK VENTS The fuel system is vented through a recessed vent coupled to a static vent on the underside of the wing, adjacent to the nacelle (Figure 5-4). One vent is recessed to prevent icing. The second vent, which is heated to prevent icing, also serves as a backup should the other vent become plugged. The wing tanks are cross-vented with one another. The wing tanks are then vented through a float-operated vent valve installed on the forward outboard side of the integral fuel tank. A line just aft of the float-operated vent

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valve extends from the integral fuel tank through a suction relief valve and aft to an air inlet on the underside of the wing.

The line from the float-operated vent valve is routed forward along the leading edge of the wing inboard to the nacelle, and aft through a check valve to the recessed vent just outboard of the nacelle. Another line tees off from the vent line and extends through a flame arrester to a heated ram vent immediately outboard and aft of the recessed vent.

Figure 5-4 Fuel Vent System—Series UA/UB

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5-5

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FUEL SYSTEM OPERATION Fuel flow from each wing main tank system is automatic without pilot action (Figure 5-5). The interconnected wing tanks gravity feed into the center wing tank and then to the collector tank through a line extending from the aft inboard wing tank to the outboard side of the center wing tank. A flapper-type check valve in the end of the gravity feed line prevents any backflow of fuel into the wing tanks.

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The fuel pressure required to operate the engine is provided by an engine-driven fuel pump immediately upstream of the fuel control unit on the accessory case. An engine-driven boost pump (also on the accessory case) provides lubrication for the high-pressure fuel pump and starts the motive flow for operation of the primary jet pump in the collector tank. The primary jet pump assists the engine-driven pumps in removing fuel from the collector tank.

Figure 5-5 Fuel Flow Diagram—Series UA/UB

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The supply line from the collector tank is routed from the aft side of the center wing tank, forward to the engine-driven boost pump through a motor-driven firewall shutoff valve. The valve is installed in the fuel line immediately aft of the engine firewall.

From the firewall shutoff valve, fuel is routed to the main fuel filter, firewall valve, and then to the engine boost pump. A bypass valve in the filter permits fuel flow even if the filter is plugged. A filter drain valve allows any impurities in the fuel system, which may have collected at the fuel filter, to be drained prior to each flight. A pressure switch at the fuel filter senses boost pump fuel pressure. Below one psi of pressure, the switch closes, actuating a red L or R FUEL PRESS warning light in the annunciator panel.

If a fuel pressure warning annunciator illuminates, the pilot can turn on the standby boost pump. Once the standby boost pump is energized, the warning light should be extinguished as fuel pressure increases above one psi. If the FUEL PRESS annunciator remains illuminated, continued engine operation is limited to 10 hours between overhaul or replacement of that engine’s high-pressure fuel pump. If either of the two standby pumps is inoperative, the use of aviation gasoline is prohibited.

Before reaching the high-pressure fuel pump, fuel is routed through the fuel heater, which uses heat from the engine oil to warm the fuel. The high-pressure fuel pump then supplies fuel to the FCU which meters the flow of fuel to the engine fuel nozzles.

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The engine-driven high-pressure fuel pump, rated at 800 psi, is mounted on the accessory case in conjunction with the fuel control unit. The high-pressure pump is protected against fuel contamination by an internal strainer. The engine-driven boost pump, also mounted on a drive pad on the aft accessory section of the engine, develops a maximum pressure of 45 psi.

The engine-driven fuel boost pump is backed up by an electric standby fuel boost pump located in the bottom of each collector tank. The standby boost pump is rated to a maximum pressure of 11 psi. Either the engine-driven boost pump or its backup unit, the standby boost pump, is capable of supplying fuel to the engine-driven highpressure fuel pump at the minimum pressure required by the engine manufacturer.

In addition to serving as a backup unit in the event of an engine-driven boost pump malfunction, the electric standby pump provides the additional pressure required for the cross-transfer of fuel from one side of the aircraft to the other.

Standby boost pump operation is controlled by lever-lock switches on the fuel control panel.

Electrical power for the pumps is supplied from the center bus, and the circuit is protected by 15ampere circuit breakers located below the fuel control panel. The battery switch must be on to provide electrical power to the standby pumps.

If all fuel boost pressure were to fail, fuel would be suction-lifted out of the collector tank by the high-pressure pump. The engine would continue to run, but engine operation would be restricted to 10 hours total time before engine high-pressure pump overhaul or replacement. If the pump is operated on suction lift beyond the 10-hour limit, overhaul or replacement of the pump is necessary.

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FIREWALL SHUTOFF VALVES The 1900 Airliner fuel system incorporates two firewall fuel valves, one for each engine (Figure 5-6). The firewall shutoff valves are motor-driven and require electrical power for opening and closing. The firewall shutoff valves are actuated by two FIRE PULL handles located in the upper center instrument panel (Figure 5-7). When the handle is pulled, its respective firewall shutoff valve closes, and fuel flow to that engine is discontinued (Figure 5-5). Pulling the handle also arms the fire extinguisher on that side.

FUEL TRANSFER MOTIVEFLOW SYSTEM Fuel pressure from the engine-driven boost pump provides the motive flow to operate the primary jet pump (Figure 5-8). If the engine-driven boost pump fails, the standby boost pump provides fuel pressure for motive flow. During start, motive flow is initiated by the engine start and ignition switch which automatically energizes the standby boost pump. The 1900 Airliner is approved for takeoff with one standby boost pump inoperative; however, cross-transfer will not be possible from the side of the inoperative standby pump.

The primary jet pump assists the engine-driven pumps in removing fuel from the collector tank.

The motive-flow fuel supply line is routed along the outboard side of the nacelle, and continues into the primary jet pump installed into the collector tank sump. A check valve in the motiveflow line prevents the fuel system from ingesting air when the boost pump is not operating.

The transfer jet pumps are also activated by motive flow. Their function is to ensure sufficient fuel supply during extreme aircraft attitude conditions.

Figure 5-6 Firewall Shutoff Valve— Series UA/UB

CROSS-TRANSFER OPERATION The two collector tanks are interconnected by a cross-transfer line (Figure 5-9). A cross-transfer valve is externally connected into the line at the forward outboard corner of the left center wing fuel cell. When the valve is in its normally-closed position, each engine draws fuel from its respective fuel tank system. A manually operated

Figure 5-7 FIRE PULL Handles—Series UA/UB

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Figure 5-8 Fuel Transfer Motive-Flow System—Series UA/UB

Figure 5-9 Cross-Transfer System—Series UA/UB

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cross-transfer control switch (labeled TRANSFER FLOW) is mounted on the upper fuel control panel, just above the fuel quantity gages. When the cross-transfer control switch is actuated, power is drawn through a circuit breaker on the lower fuel control panel to the solenoid of the cross-transfer valve. The crosstransfer valve then opens to allow fuel to transfer to either the left wing or right wing fuel system.

The electric standby pump on the transferring side is automatically energized during all crosstransfer operations. Before turning on the transfer flow switch, both standby pump switches should be in the OFF position since a standby pump in operation on the receiving side will not allow cross-transfer to occur. In the event of an inoperative electric boost pump, cross-transfer can only be accomplished from the side of the operative pump.

Cross-transfer can be used to balance fuel loads between left and right fuel systems, and for single-engine operations. Procedures are detailed in the POH Normal Procedures section.

FUEL PURGE SYSTEM This airplane is equipped with a fuel purge system (Figure 5-10). The purge system is designed to assure that any residual fuel in the fuel manifolds is consumed during engine shutdown.

During normal engine operation, compressor discharge air (P3 air) is routed through a filter and two check valves, pressurizing a small air tank on the engine truss mount. Upon engine shutdown the pressure differential between the air tank and fuel manifold causes air to be discharged from the air tank into the fuel manifold system. The air forces all residual fuel out through the nozzles and into the combustion chamber where it is consumed. A momentary surge in N1 may be noticed as fuel is burned. The entire operation is automatic and requires no crew action. During engine starting, fuel manifold pressure closes the fuel manifold valve, allowing P3 air to pressurize the purge tank.

5-10

Figure 5-10 Fuel Drain Purge System Schematic—Series UA/UB

FUEL GAGING SYSTEM The airplane is equipped with a capacitance-type fuel quantity indication system (Figure 5-11). A maximum indication error of 3% may be encount e r e d i n t h e s y s t e m . T h e ga g i n g s y s t e m compensates for changes in fuel density, which result from differences in ambient temperatures.

A Density Variation of Aviation Fuel graph is provided in the Weight and Balance section of the POH to allow more accurate readings for all approved fuels.

The LEFT fuel quantity indicator on the fuel control panel indicates the amount of fuel remaining in the left-wing fuel system tanks. The RIGHT fuel quantity indicator provides the same information for the right-wing fuel systems. Both gages are marked in pounds.

Each side of the airplane has an independent gaging system, consisting of a capacitance probe in the collector tank, one in the aft inboard fuel cell, two in the integral (wet wing) fuel cell, two in the inboard leading edge fuel cell, and two in the center wing fuel cell (Figure 5-12).

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The fuel quantity indicating system is also compensated for specific gravity and reads in pounds on a linear scale. An electronic circuit in the system processes signals from the fuel capacitance probes in the fuel cells for an accurate readout by the fuel quantity indicators. Power to the fuel quantity indicators is supplied from the capacitance probes through a 5 ampere circuit breaker on the fuel system circuit breaker panel.

The fuel quantity probe is simply a variable capacitor comprised of two concentric tubes. The inner tube is profiled by changing the diameter as a function of height so that the capacitance between the inner and outer tube is proportional to the tank volume. The tubes serve as fixed electrodes and the fuel of the tank in the space between the tubes acts as the dielectric of the fuel quantity probe.

Figure 5-11 Fuel Probe—Series UA/UB

The capacitance of the fuel quantity probe varies with respect to the change in the dielectric that results from the ratio of fuel to air in the fuel cell.

As the fuel level between the inner and outer tubes rises, air with a dielectric constant of one is replaced by fuel with a dielectric constant of approximately two, thus increasing the capacitance of the fuel quantity probe. This variation in the volume of fuel contained in the fuel cell produces a capacitance variation that is a linear

Figure 5-12 Fuel Gaging System—Series UA/UB

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5-11

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function of that volume. This function is converted to linear current that actuates the fuel quantity indicator. Fuel density and electrical dielectric constant vary with respect to temperature, fuel type, and fuel batch. The capacitance gaging system is designed to sense and compensate for these variables.

Low-fuel quantity in the center wing tanks is sensed by photoelectric cells which transmit energizing current for the L or R FUEL QTY caution annunciators. Illumination of a FUEL QTY annunciator indicates that enough fuel remains to sustain maximum continuous power on that engine for approximately 30 minutes (computed for sea level fuel requirements). Photoelectric sensors in the collector tanks provide the signals for L or R FUEL FEED caution annunciators. A FUEL FEED annunciator indicates that less than two minutes of fuel remains in that fuel system.

The fuel quantity annunciators and sensors can be functionally tested by pressing the annunciator test switch on the glareshield.

Takeoff is prohibited when the fuel quantity indicator needles are in the yellow arc. Maximum fuel remaining in this range (top of the yellow arc) is 363 pounds.

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FUEL DRAINS During preflight, the fuel sumps on the tanks, pumps and filters should be drained to check for fuel contamination. There are three sump drains in each wing (Figure 5-13).

The collector tank sump drain is located in the center wing adjacent to the fuselage; the inboard leading-edge tank sump drain is on the underside of the wing just outboard of the nacelle; and the integral (wet wing) fuel tank sump drain is located approximately midway on the underside of the wing aft of the main spar. Other drains are the fuel filter drain, in the main landing gear wheelwell, and the center wing tank drain at the wing root forward of the flap.

Since jet fuel and water are of similar densities, water does not settle out of jet fuel as easily as from aviation gasoline. For this reason, the airplane must sit perfectly still, with no fuel being added, for approximately three hours prior to draining the sumps if water is to be removed.

Although water ingestion is not as critical for turbine engines as it is for reciprocating engines, water should still be removed periodically to prevent formations of fungus and contaminationinduced inaccuracies in the fuel gaging system.

When draining flush-mounted drains, do not turn the draining tool. Turning or twisting the tool will unseat the “O-ring” seal and may cause a leak.

Figure 5-13 Fuel Drains—Series UA/UB

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FUEL TANK SYSTEM

The Beechcraft 1900 Airliner fuel system (Figures 5-15 and 5-16) is designed with the pilot in mind; simple to use in normal and emergency conditions with one or more failures. Simple, quick (over-the-wing) refueling is also incorporated to minimize ground turnaround time requirements. In the Series UC/UE and after, the 1900 Airliner differs from previous series in the completely integral wetwing design. Range has been increased approximately 55% without increasing the basic airplane weight. This combination increases the usefulness of the Series UC/UE airliners considerably.

The wet wing fuel system is quite different from the previous series airliners, although pilot operation of the two systems is very similar. The fuel system used in previous series is covered in a separate section of this manual, and of course by a different Pilot’s Operating Handbook. This Section will discuss the Series UC/UE airliner fuel program, including the wet-wing fuel tank arrangement, fuel system components, controls, operation of the system and the associated annunciators.

The wet-wing fuel system consists of two integral fuel tanks in each wing (Figure 5-14). A main tank extends from engine nacelle to wing tip. An auxiliary tank is located between the engine nacelle and the fuselage. The usable fuel in the airliner and the maximum zero fuel weight for UC and UE Series Airliners are shown in Table 5-1. The maximum allowable fuel imbalance between the wings is 200 pounds.

Table 5-1 USABLE FUEL UC

UE

Maximum usable fuel (gallons)

667.2

665.4

Each main tank (gallons)

241.3

240.5

Each auxiliary tank (gallons)

92.3

92.2

14,000

15,165

Maximum zero fuel weight (lbs.)

Figure 5-14 Fuel Tank System—Series UC/UE

Revision 1

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Figure 5-15 Fuel System Diagram—Series UC and After

BEECH 1900 AIRLINER

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Figure 5-16 Fuel System Schematic—Series UE

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5-15

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Figure 5-17 Fuel Control Panel—Series UC/UE

Figure 5-18 Fuel Tank Vent System—Series UC/UE

5-16

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FUEL CAPACITY The main tanks are filled through fill ports located near the wing tips. A collector tank is contained within each main tank immediately outboard of the nacelle. Each collector tank is filled from its main tank by gravity feed and two jet transfer pumps, which maintain the fuel level in the collector tank at normal flight attitudes.

Each auxiliary tank is filled through its own fill port located just inboard of the engine nacelle.

When auxiliary tank fuel is required for a planned flight, the main tanks should be full and the additional fuel to complete the flight placed in the auxiliary tanks. The auxiliary tank fuel should be used first. There is no gravity flow between the main and auxiliary tanks, therefore, each must be filled separately.

FUEL TANK VENTS Under certain conditions, the fuel vent system will allow fuel to flow from the main tank to the auxiliary tank (Figure 5-18). These conditions occur at lower power settings when the auxiliary tank transfer pump is supplying fuel to the collector tank, and the high-pressure pump purge line is directing fuel back to a full main tank.

Since the main tank is already full, any excess fuel flows through the vent system back to the auxiliary tank. This condition can continue until the auxiliary tank is empty. The second condition which will allow fuel to flow from the main tank to the auxiliary tank is thermal expansion. Fuel will not vent outside from the wing unless the auxiliary tank is full.

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The fuel system is vented through a float operated valve near each wing tip. The system contains a flush vent with flame arrester, a heated ram-air vent, to maintain a slight positive tank pressure during flight, and a recessed ram vent.

The recessed ram vent is coupled to the protruding ram vent on the underside of the wing tip.

The recessed vent is naturally ice resistant, while the protruding vent is heated to prevent icing.

The fuel vent heat switch is located on the pilot’s subpanel in the ice protection group.

The check valves in the vent tubes allow the air to flow one way through the vents. The flame arrestors, on the flush vent and incoming line, prevent a flame front produced by a lightning strike or static discharge from traveling up the vent line into the tank system. The vent system also incorporates a pressure-activated relief tube which prevents an overpressure condition in the tank. A valve in the tube opens when the pressure exceeds a set amount. Vent lines connect the main tank and auxiliary tank as we have discussed earlier.

As fuel is used from the main tank, it is gravityfed and also pumped through motive flow to the collector tank. The cross vents to the auxiliary tank then are open and equalize the pressure in all tanks. An anti-siphon valve is installed in each tank filler port to prevent loss of fuel through siphoning in the event of improper securing or loss of the filler cap.

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BEECH 1900 AIRLINER

FUEL SYSTEM OPERATION This fuel system operation is designed to be fully automatic. Power for the aux transfer pumps is supplied through their respective generator busses. All other fuel system functions on the control panel require only battery power.

During normal operation, fuel flow to each engine is provided by the engine-driven fuel pumps (high pressure and boost) which draw fuel from the collector tank in the same wing (Figure 5-19). The collector tank draws from its respective main tank unless fuel is being supplied from the auxiliary tank. Any fuel contained in the auxiliary tanks is to be used prior to using fuel from the main tanks.

The auxiliary tank fuel will be used first. A loss of electrical power or failure of the transfer pump will prevent the use of auxiliary tank fuel. The auxiliary tank will not gravity feed into the main tank fuel system.

The auxiliary tank transfer pump uses fuel for cooling. A thermal cutout switch has been incor-

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porated within the pump should the automatic features fail to turn off the pump when the auxiliary tank fuel has been used. The auxiliary transfer pump can also be turned ON or OFF by placing the AUX pump switch in the desired position to bypass the automatic features.

Engine fuel is supplied from the collector tank through the motive-flow system operated by either the engine-driven boost pump, or the standby pump within the collector tank. The motive-flow system is supported by a series of three transfer jet pumps (main, forward, and aft).

The main jet pump is located within the collector tank. It picks up fuel from the collector tank and sends fuel to the engine. The aft jet pump is also located within the collector tank but it draws fuel from the main tank and fills the collector tank.

The forward jet pump draws fuel from the main tank, forward of the wing spar, and sends it to the collector. There are also three flapper valves and three upper wing stringer cutouts that will allow the collector to be gravity fed from the main tank.

Figure 5-19 Fuel Flow Diagram—Series UC/UE

5-18

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The engine-driven boost pump draws fuel from the collector tank using the jet pump principle.

The components we are concerned with are the motive-flow line, the primary jet pump, two transfer jet pumps, the standby electric boost pump, and the fuel supply line. The transfer jet pump’s function is to keep the collector tank full by transferring fuel from the main fuel tank to the collector tank by means of the venturi effect.

Fuel from the motive-flow line passes through the primary jet pump, which is actually a venturi.

In order to pull fuel from the collector tank, a venturi effect is used. As a mass of fuel is accelerated through a small opening or venturi, it causes a drop in pressure. At this low-pressure point, fuel from the collector tank enters the fuelsupply line through a filter and low-pressure fuel is supplied as needed. The transfer jet pumps draw fuel into the collector tank from the main tank in the same manner.

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LOW-FUEL WARNING SYSTEM (SERIES UE) The low-fuel warning system operation and components are the same as the Series UC system, however, specific quantities have been established for each sensor. The FUEL QUANTITY annunciator units will be activated by the fuel level sensors if the main tank quantity drops below 324 pounds. If the fuel in the collector tank drops below 53 pounds, the COL TANK LOW annunciator will be activated.

FIREWALL SHUTOFF VALVES The 1900 Airliner fuel system incorporates two firewall fuel valves, one for each engine. The firewall shutoff valves are motor-driven and

If the primary engine-driven boost pump fails, fuel can be supplied to the system by the standby electric boost pump. This pump draws fuel directly from the collector tank and passes it to the fuel supply line. As fuel is drawn from the collector tank, it flows through the manual shutoff valve, a fuel filter and through the firewall shut-off valve.

LOW-FUEL WARNING SYSTEM (SERIES UC) The low-fuel warning system provides bilevel warning to the flight crew when the fuel level in each tank reaches a predetermined level, and again when the fuel in the collector tank area is nearly exhausted. When fuel in the main tanks reaches a level allowing approximately 30 minutes of flight time at maximum continuous power, the right, or left, or both FUEL QTY annunciators will be actuated by fuel level sensors mounted on the forward side of the tank. If the fuel in the collector tank area drops to a reserve of two minutes flight time at maximum continuous power, the right, left or both, FUEL FEED annunciators will light up.

Figure 5-20 Firewall Shutoff Valve— Series UC/UE

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Figure 5-21 FIRE PULL Handles—Series UC/UE

require electrical power for opening and closing.

The firewall shutoff valves are actuated by two FIRE PULL handles located in the upper center instrument panel (Figure 5-21). When the handle is pulled, its respective firewall shutoff valve closes, and fuel flow to that engine is discontinued (Figure 5-20). Pulling the handle also arms the fire extinguisher on that side.

FUEL TRANSFER MOTIVEFLOW SYSTEM Fuel pressure from the engine-driven boost pump provides the motive flow to operate the primary jet pump. If the engine-driven boost pump fails, the standby boost pump provides fuel pressure for motive flow. During start, motive flow is initiated by the engine start and ignition switch which automatically energizes the standby boost pump.

The 1900 Airliner is approved for takeoff with one standby boost pump inoperative; however, cross-transfer will not be possible from the side of the inoperative standby pump.

The primary jet pump assists the engine-driven pumps in removing fuel from the collector tank.

The motive-flow fuel supply line is routed along the outboard side of the nacelle, and continues into the primary jet pump installed into the collector tank sump. A check valve in the motiveflow line prevents the fuel system from ingesting air when the boost pump is not operating.

The transfer jet pumps are also activated by motive flow. Their function is to ensure sufficient fuel supply during extreme aircraft attitude conditions.

5-20

CROSS-TRANSFER OPERATION A cross-transfer line connects the collector tanks in each wing (Figure 5-22). A switch-controlled cross-transfer valve in the left wing is externally connected into the line. When the valve is in its normally-closed position, each engine draws fuel from its respective fuel tank system.

A manually operated cross-transfer control switch is mounted on the upper fuel control panel, just above the fuel quantity gages. When the cross-transfer control switch is actuated, the cross-transfer valve opens to allow the standby fuel boost pump to transfer fuel to the opposite collector tank. In addition to the cross-transfer function, the electric boost pump can provide fuel to the engine should the engine-driven boost pump fail. Power for the switches is drawn through the circuit breakers at the bottom of the fuel panel.

During single-engine operation, it may become necessary to supply fuel to the operative engine from the fuel system on the opposite side. The simplified cross-transfer system is placarded for fuel selection with a diagram on the upper fuel control panel. The STANDBY PUMP switches are placed in the OFF position when cross transf e r r i n g . A l eve r- l o c k s w i t c h , p l a c a r d e d TRANSFER FLOW OFF, is moved from the center OFF position to the left or to the right, depending on direction of flow. This opens the cross-transfer valve, energizing the standby pump on the side from which cross transfer is desired. In the event one of the electric boost pumps fail, cross-transfer can only be accomplished from the side of the operative pump.

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Figure 5-22 Cross-Transfer Schematic—Series UC/UE

On the Caution advisory panel, a green FUEL TRANSFER indicator will illuminate to indicate that the cross-transfer is selected (Series UC). In t h e U E S e r i e s , t wo l i g h t s i n t h e C AU TION/ADVISORY panel indicate the condition of the fuel transfer system. The yellow XFR VALVE FAIL annunciator is activated if the transfer valve fails to move to its assigned posit i o n w i t h i n t wo s e c o n d s . A w h i t e F U E L TRANSFER annunciator is activated if the fuel transfer valve is in full open position. To discontinue fuel transfer operations, the transfer flow switch need only be placed in the center OFF position.

starting, fuel manifold pressure closes the fuel manifold valve, allowing P3 air to pressurize the purge tank.

FUEL PURGE SYSTEM This airplane is equipped with a fuel purge system (Figure 5-23). The purge system is designed to assure that any residual fuel in the fuel manifolds is consumed during engine shutdown.

During normal engine operation, compressor discharge air (P3 air) is routed through a filter and two check valves, pressurizing a small air tank on the engine truss mount. Upon engine shutdown the pressure differential between the air tank and fuel manifold causes air to be discharged from the air tank into the fuel manifold system. The air forces all residual fuel out through the nozzles and into the combustion chamber where it is consumed. A momentary surge in N1 may be noticed as fuel is burned. The entire operation is automatic and requires no crew action. During engine

Figure 5-23 Fuel Drain Purge System Schematic—Series UC/UE

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Figure 5-24 Fuel Control Panel—Series UE

FUEL CONTROL PANEL The left and right fuel quantity indicators on the fuel control panel indicate the amount of fuel remaining in their respective main fuel tanks (Figures 5-17 and 5-24). Deflecting the springloaded “FUEL QTY” switch on the fuel control panel to the “AUX” position will cause the indicators to indicate the fuel quantity in the auxiliary tanks. The indicators are marked in pounds. The pilot must hold the spring-loaded fuel quantity selector switch in the “AUX” position to verify the quantity remaining in each auxiliary tank. Once the switch is released, the fuel indicator will return to read only the main tank quantity.

FUEL GAGING SYSTEM The airplane is equipped with a capacitance-type fuel quantity indication system (Figure 5-24). A maximum indication error of 3% may be encount e r e d i n t h e s y s t e m . T h e ga g i n g s y s t e m compensates for changes in fuel density, which result from differences in ambient temperatures.

A Density Variation of Aviation Fuel graph is provided in the Weight and Balance section of

5-22

the POH to allow more accurate readings for all approved fuels.

Fuel quantity probes, which are part of the fuel gaging system, are capacitance-type probes.

These probes measure the density of fuel on board, and are calibrated to read pounds of fuel.

This system is necessary on this airplane because the engines operate on weight flow of fuel rather than gallons. The gages in the cockpit therefore read in pounds of fuel flow and pounds of fuel remaining in the tanks. To provide the pilot with such a readout, it is necessary to use a system which compensates for changes in the specific gravity of the fuel in use. Therefore, each probe is designed to compensate for differences in specific gravity. Each main tank contains six fuel quantity probes, and each auxiliary tank two probes. Information from these probes is relayed to the fuel panel in the cockpit to show fuel remaining in each tank.

STANDBY PUMP OPERATION The electrically driven, standby pumps are activated by individual “On/Off” switches on the pilot’s left side panel. These pumps are activated

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if problems arise with the primary engine-driven boost pumps. The standby pumps are automatically activated during engine start by the START and IGNITION switches. The standby pumps are also used for all cross-transfer operations. In the event one of the electric pumps fail, cross-transfer can only be accomplished from the side of the operative pump.

PILOT TRAINING MANUAL

With fuel in the auxiliary tank and the transfer switch in AUTO, the auxiliary tank transfer pump will run once the 10 psi low-pressure switch is activated. It will continue to run until the boost pressure falls below 10 psi, or the float switch and transfer line low-pressure switches open; in either case the auxiliary tank transfer pump will shut down.

AUXILIARY TANK OPERATION To use fuel from the auxiliary tanks, position the “Aux Pump” switches on the fuel control panel in the “Auto” position. This activates the electric transfer pump in each auxiliary tank and pumps fuel to the collector tank of the same wing. Fuel will continue to be transferred until the auxiliary tank is empty, at which time the pump will automatically shut off. In the event of a transfer system failure, it is permissible to temporarily operate the airplane with fuel in the auxiliary tanks providing fuel imbalance and fuel reserve requirements can be met.

The auxiliary tank transfer pump is controlled by a three-position switch located on the fuel control panel, labeled, “ON-AUTO-OFF”. Normal procedures call for the pump to be left in the AUTO position. In AUTO, there are four additional non-pilot operated control features.

There is a float switch within the auxiliary tank, a five PSI low-pressure switch in the transfer line between the auxiliary and main tank operates LEFT or RIGHT NO-FUEL transfer (L or R NO AUX XFR - UE) annunciator lights), a 10 psi low-pressure switch located after the enginedriven fuel boost pump, and a thermal cut-out switch on the pump.

WARNING Takeoff is prohibited when the fuel quantity indicator needles are in the yellow arc. Maximum fuel remaining in this range (top of the yellow arc) is 363 pounds.

VISUAL FUEL QUANTITY SENSORS (SERIES UE) Two visual fuel quantity gauges are located on the lower surface of each wing to provide the ground crew and pilots an alternate method of fuel gauging in the event of a failure in the capacitance system. The gauges, when not submerged in fuel, are red with a black dot; when they are submerged they are totally black. The outboard probe, when red, indicates less than 1,150 pounds of fuel, the inboard probe, when red, indicates less than 745 pounds of fuel.

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Figure 5-25 Fuel Drains—Series UC/UE

FUEL DRAINS

prevent formations of fungus and contaminationinduced inaccuracies in the fuel gaging system.

During preflight, the fuel sumps on the tanks, pumps and filters should be drained to check for fuel contamination. There are six (five in Series UE) sump drains in each wing (Figure 5-25).

The two (one in Series UE) collector tank sump drains are located below the wing on the outboard side of the nacelle; the two main tank drains are located on the underside of the wing, outboard of the nacelle, one forward and one aft of the main wing spar: the other drains are the fuel filter drain located on the underside of the wing, outboard of the nacelle under a springloaded access panel and the auxiliary tank drain at the wing root forward of the flap.

Since jet fuel and water are of similar densities, water does not settle out of jet fuel as easily as from aviation gasoline. For this reason, the airplane must sit perfectly still, with no fuel being added, for approximately three hours prior to draining the sumps if water is to be removed.

Although water ingestion is not as critical for turbine engines as it is for reciprocating engines, water should still be removed periodically to

5-24

When draining flush-mounted drains, do not turn the draining tool. Turning or twisting the tool will unseat the “O-ring” seal and may cause a leak.

FUEL DRAIN LOCATIONS Table 5-2 FUEL DRAIN LOCATIONS DRAINS Auxiliary tank (1)

LOCATION Underside of wing, inboard of nacelle

Collector tank (2) UC Outboard side of nacelle (1) UE Main tank (2)

Underside of wing, outboard of nacelle

Fuel filter (1)

Underside of wing, outboard of nacelle

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FUEL HANDLING PRACTICES (SERIES UA/UB AND UC/UE) All hydrocarbon fuels contain some dissolved and some suspended water. The quantity of water contained in the fuel depends on its type and temperature. Kerosene, with its higher aromatic content, tends to absorb and suspend more water than aviation gasoline. Along with water, kerosene will suspend rust, lint and other foreign materials longer. Given sufficient time, suspended contaminants will settle to the bottom of the tank.

The settling time for kerosene is five times that of aviation gasoline; therefore, jet fuels require good fuel handling practices to ensure servicing with clean fuel. If recommended ground procedures are carefully followed, solid contaminants will settle, and free water can be reduced to 30 parts per million (ppm), a value considered acceptable by the major airlines.

Dissolved water has been found to be the major potential fuel contaminant. Its effects are multiplied in aircraft that operate primarily in humid regions and in warm climates. Since most suspended matter, including water, can be removed from the fuel by allowing sufficient settling time and by proper filtration, fuel contamination is usually not a major problem.

Dissolved water cannot be filtered from the fuel by micronic-type filters used in the fuel system; however, water in the fuel can be released by lowering fuel temperature, which occurs in flight.

For example, a kerosene fuel may contain 65 ppm (8 ounces per 1,000 gallons) of dissolved water at 80° F. When fuel temperature is lowered to 15° F, only about 25 ppm will remain in solution. The difference of 40 ppm will have been released as super-cooled water droplets which need only a piece of solid contaminant or an impact shock to convert them into ice crystals.

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Tests indicate that released, super-cooled water droplets will not settle during flight. Droplets are pumped freely through the system. If they become ice crystals in the tank, they will not settle since the specific gravity of ice is approximately equal to that of kerosene. 40 ppm of suspended water seems like a very small quantity, but when added to water suspended in the fuel at the time of delivery, that amount is sufficient to ice a filter. Although severe fuel system icing can occur at fuel temperatures from 0 to -20°F, water droplets can freeze at any temperature below 32° F. Water in jet fuel also creates an environment favorable to the growth of a microbiological “sludge” in settlement areas of the fuel cells.

Sludge and other fuel contaminants can cause corrosion of metal parts in the fuel system and clogging of the fuel filters. Although the 1900 Airliner uses integral (wet wing) fuel cells in each wing, and all metal parts (except the standby boost pumps and jet transfer pumps) are mounted above the settlement areas, consistently using contaminated fuels can cause filters to clog and fuel pumps to corrode.

The primary means of fuel contamination control is “good housekeeping” practices by the owner/operator. This applies not only to maintaining a clean fuel supply, but to keeping the aircraft system clean. The following is a list of steps that may be taken to recognize and prevent contamination problems. a. Know your supplier. It is impractical to assume that contaminant-free fuel will always be available. But, it is feasible to exercise caution and be watchful for signs of fuel contamination. b. Be sure, as much as possible, that fuel has been properly stored. Fuel should be filtered as it is pumped to the truck, and again as it is pumped from the truck to the aircraft. c. Perform filter inspections to determine if sludge is present.

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Figure 5-26 Fuel Temperature Graph

d. Maintain good housekeeping by periodically flushing the fuel tankage system. The frequency of flushing will be determined by the climate and the presence of sludge. e. Aviation gas is an emergency fuel. If avgas has been used, observe the requirement for 150 hours maximum operation on aviation gasoline before engine overhaul. The time should be logged in the aircraft engine operation records as gallons of avgas added to the fuel system. f. Use only clean fuel servicing equipment. g. After refueling, allow a settling period of at least three hours, whenever possible; then drain a small amount of fuel from each drain. h. Fuel spills on airplane tires have a deteriorating effect. Be sure to remove spilled fuel from the ramp area immediately to prevent tire damage.

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Even if the fuel does not contain water, or if water has been drained, the possibility of fuel icing still exists at some very low temperatures.

The oil-to-fuel heat exchanger prevents fuel icing during most normal operating conditions; however, in extremely cold temperatures at some cruise altitudes, anti-icing fuel additives must be used.

The Minimum Oil Temperature Required for Operation Without Anti-Icing Additive chart is found in the POH Limitations section. This chart is used as a guide in preflight planning to determine operating temperatures where icing at the fuel control unit could occur (Figure 5-26). Enter the graph with the known or forecast outside air temperature at cruise, and plot vertically to the expected cruise pressure altitude. Since no fuel temperature measurement is available prior to the heat exchanger, fuel temperature must be assumed to be the same as outside air temperature.

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Follow the graph using this example: enter outside air temperature at -30° C and vertically follow the graph to a cruise pressure altitude of 5,000 feet. Next, plot horizontally to determine the minimum oil temperature required to prevent icing. In this example, the minimum oil temperature required is 31° C. The 1900 Airliner maintains a constant oil temperature, although the exact temperature varies from one airplane to another. For most, the oil temperature maintains a constant 50 to 60° C. Compare the minimum oil temperature obtained in the preceding example with the normal oil temperature of the airplane to be used for the flight to determine if anti-icing additive is needed. When required, anti-icing additive conforming to specification MIL-I-27686 should be added during fueling.

FUEL GRADES AND TYPES (SERIES UA/UB AND UC/UE) Jet A, Jet A-1, Jet B, and JP-4 fuels may be mixed in any ratio in the 1900 Airliner fuel system. Aviation gasoline grades 80/87, 100LL, 100/130, and 115/145 are emergency fuels and may be mixed in any ratio with jet fuels.

If the 1900 Airliner is fueled with aviation gasoline, some operational limitations must be observed. If use of aviation gasoline is necessary, operation is limited to 150 hours before engine overhaul, and is prohibited if either of the two standby pumps is inoperative or if flight is conducted above 15,000 feet (18,000 feet in Series UE). When avgas is used, lead deposits form on the turbine wheels causing power degradation; therefore, when operating on avgas, the lowest octane rating available should be used because its lead content is lowest.

Since the aviation gas will probably be mixed with jet fuel already in the tanks, it is easier to record the number of gallons of avgas added than to note hours of operation. If an engine has an average fuel consumption of 55 gallons per hour, each time 55 gallons of aviation gasoline are added, one hour of the 150 hour limitation is being used.

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A chart in the POH/POM Weight and Balance section shows the density of avgas to be considerably less than that of jet fuel. Because it is less dense, aviation gas delivery is much more critical than jet fuel delivery; therefore, operation on avgas is prohibited if above 15,000 feet (18,000 feet in Series UE). Aviation gas feeds well under pressure but cannot suction feed as well, particularly at high altitudes. For this reason, two alternate means of pressure feed must be available. Standby pumps provide alternate pressure feed capability, and both are required to be operational when avgas is used.

The Fuel Brands and Type Designations Chart in the Handling, Servicing and Maintenance section of the POH/POM gives fuel refiner’s brand names, and the corresponding type designations established by the American Petroleum Institute (API) and the American Society of Testing Material (ASTM). Brand names are listed for easy reference and are not specifically recommended

by Beech Aircraft Corporation. Any product conforming to the recommended specification may be used.

FILLING THE TANKS (SERIES UA/UB AND UC/UE) When filling the aircraft fuel tanks, always observe the following: a. Make sure the aircraft and the servicing unit are both grounded to the ground, and that the aircraft is statically grounded to the serving unit. b. The filler caps are located in the main fuel tank on the leading edge of each wing near the wing tip and the auxiliary tank fuel caps are just inboard of each nacelle.

Do not rest fuel nozzle in tank fillers because this may damage the filler neck. c. Allow a three-hour settling period whenever possible, then drain a sufficient amount of fuel from each drain point to remove water or contaminants.

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DRAINING THE FUEL SYSTEM (SERIES UA/UB AND UC/UE) A defueling adapter is located inboard of the standby pump in the center wing tank. The adapter contains a check valve to prevent fuel drainage when the plug is removed. Each wing fuel system may be drained as follows: a. Cut the safety wire and remove the plug.

This will seat the check valve. b. Thread an AN832-12 adapter into the drain, unseating the check valve to start the flow of fuel. Fuel will gravity drain.

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The material normally covered in this chapter is not applicable to this airplane.

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CHAPTER 7 POWERPLANT CONTENTS Page INTRODUCTION .................................................................................................................. 7-1 DESCRIPTION....................................................................................................................... 7-1 1900 AIRLINER POWERPLANT ......................................................................................... 7-3 Engine Stations................................................................................................................ 7-7 Engine Terms .................................................................................................................. 7-7 Engine Airflow ................................................................................................................ 7-8 Compressor Bleed Valve............................................................................................... 7-10 Jet-Flap Intake System .................................................................................................. 7-11 Swing Check Valve (Series UA through UB-40) ......................................................... 7-11 Igniters........................................................................................................................... 7-11 Accessory Section ......................................................................................................... 7-12 Lubrication System........................................................................................................ 7-13 Magnetic Chip Detector (Series UA, UB)..................................................................... 7-16 ENGINE FUEL SYSTEM .................................................................................................... 7-16 Fuel Manifold Purge System ......................................................................................... 7-17 Fuel Control Unit (FCU) ............................................................................................... 7-18 FCU Operation .............................................................................................................. 7-18 Fuel Flow Indicators...................................................................................................... 7-19 Fuel Pressure Indicators ................................................................................................ 7-20 Anti-icing Fuel Additive ............................................................................................... 7-20 Fuel Biocide Additive ................................................................................................... 7-20

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CONTROLS AND INDICATIONS..................................................................................... 7-20 Control Pedestal ............................................................................................................ 7-20 Engine Power Control................................................................................................... 7-20 Power Levers ................................................................................................................ 7-21 Propeller Levers ............................................................................................................ 7-21 Condition Levers........................................................................................................... 7-21 Condition Lever Operation (Series UA, UB, UC)........................................................ 7-21 Condition Lever Operation (Series UE) ....................................................................... 7-22 ITT and Torquemeters .................................................................................................. 7-22 ITT Gage....................................................................................................................... 7-22 Torquemeter (Series UA, UB, UC)............................................................................... 7-23 Torquemeter (Series UE) .............................................................................................. 7-23 Gas Generator (N1) Tachometer (Series UA, UB, UC)................................................ 7-23 Gas Generator (N1) Tachometer (Series UE) ............................................................... 7-23 ENGINE LIMITATIONS..................................................................................................... 7-27 STARTER OPERATING TIME LIMITS............................................................................ 7-29 TREND MONITORING ...................................................................................................... 7-29 DATA COLLECTION ......................................................................................................... 7-30 PROPELLER SYSTEM ....................................................................................................... 7-31 General.......................................................................................................................... 7-31 Blade Angle .................................................................................................................. 7-33 Primary Governor ......................................................................................................... 7-33 Low Pitch Stop.............................................................................................................. 7-38 Beta and Reverse Control ............................................................................................. 7-38 Flight and Ground Low Pitch Stops ............................................................................. 7-41

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Overspeed Governor...................................................................................................... 7-44 Fuel Topping Governor ................................................................................................. 7-45 Power Levers................................................................................................................. 7-45 Propeller Control Levers ............................................................................................... 7-46 Propeller Feathering ...................................................................................................... 7-46 Before Taxi and Before Takeoff Checks....................................................................... 7-46 Autofeather System (Series UA, UB, UC).................................................................... 7-47 Autofeather System (Series UE) ................................................................................... 7-49 Propeller Synchrophaser Systems ................................................................................. 7-49

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ILLUSTRATIONS Figure

Title

Page

7-1

PT6A-65B Powerplant Installation.......................................................................... 7-2

7-2

Engine Installation ................................................................................................... 7-3

7-3

Engine Cutaway....................................................................................................... 7-5

7-4

Engine Gas Flow...................................................................................................... 7-6

7-5

Free-Turbine Reverse Flow Principle...................................................................... 7-7

7-6

Engine Modular Concept ......................................................................................... 7-8

7-7

Jet-Flap, Compressor Bleed Valve, Swing Check Valve ...................................... 7-10

7-8

Engine Start and Ignition Switches........................................................................ 7-12

7-9

Typical PT6A Engine ............................................................................................ 7-13

7-10

Engine Lubrication Diagram ................................................................................. 7-14

7-11

Engine Oil Dipstick ............................................................................................... 7-15

7-12

Magnetic Chip Detector......................................................................................... 7-16

7-13

Simplified Fuel System Diagram........................................................................... 7-17

7-14

Simplified Fuel Control System ............................................................................ 7-19

7-15

Fuel Flow Gages .................................................................................................... 7-19

7-16

Fuel Pressure Annunciator..................................................................................... 7-20

7-17

Control Pedestal..................................................................................................... 7-21

7-18

Control Levers ....................................................................................................... 7-22

7-19

Engine Instrument Markings (Series UA, UB, and UC) ....................................... 7-24

7-20

Engine Instruments (Series UE-1 through UE-92) ................................................ 7-25

7-21

Engine Instruments (Series UE-93 and After)....................................................... 7-26

7-22

Engine Limits Chart—PT6A-65B ......................................................................... 7-27

7-23

Engine Limits Chart—PT6A-67D ......................................................................... 7-28

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7-24

Overtemperature Limits (Starting) ........................................................................ 7-28

7-25

View through Exhaust Duct .................................................................................. 7-29

7-26

In-Flight Engine Data Log .................................................................................... 7-30

7-27

Hartzell Propeller .................................................................................................. 7-31

7-28

Propeller Tiedown Boot Installed ......................................................................... 7-31

7-29

Propeller Blade Angle Diagram ............................................................................ 7-32

7-30

Primary Governor Diagram................................................................................... 7-34

7-31

Propeller Onspeed Diagram .................................................................................. 7-35

7-32

Propeller Overspeed Diagram ............................................................................... 7-36

7-33

Propeller Underspeed Diagram ............................................................................. 7-37

7-34

Beta and Reverse Control...................................................................................... 7-39

7-35

Beta Range and Reverse Diagram......................................................................... 7-40

7-36

Propeller Postioning Diagram ............................................................................... 7-42

7-37

Overspeed Governor Diagram .............................................................................. 7-44

7-38

Power Levers......................................................................................................... 7-45

7-39

Propeller Control Levers ....................................................................................... 7-46

7-40

Autofeather Test Diagram—Series UA, UB, UC ................................................. 7-47

7-41

Autofeather System Diagram (Armed)—Series UA, UB, UC.............................. 7-48

7-42

Autofeather System Diagram (Armed, Left Engine Failure)— Series UA, UB, UC ............................................................................................... 7-48

7-43

Autofeather System Diagram—Series UE............................................................ 7-50

7-44

Propeller Synchrophaser ....................................................................................... 7-51

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TABLES Table

Title

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PT6A-65B Specifications (Sea Level Static ICAO Standard Atmosphere Conditions) .................................... 7-4

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PT6A-67D Specifications (Sea Level Static ICAO Standard Atmosphere Conditions) .................................... 7-4

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PILOT TRAINING MANUAL

CHAPTER 7 POWERPLANT #1 DC GEN

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INTRODUCTION In-depth knowledge of powerplants is essential to the pilot’s ability to operate the engines.

Performance within the normal parameters of powerplant and propeller systems extends engine life and ensures safety. This chapter describes basic engine components, limitations, and system checks. In-depth knowledge of the propeller system is essential to proper operation of the engine power system. Operating within safe parameters of the powerplant and propeller systems extends engine life and ensures safety. This chapter also describes the propeller system and its operational limits and preflight checks.

DESCRIPTION This chapter describes and discusses the Pratt & Whitney PT6A-65B and -67D turboprop engines.

The purpose of this chapter is to provide pilots with sufficient engine operating details to further understand normal, abnormal, and emergency procedures.

of propeller controls, principles of operation, reversing systems, and propeller feathering are detailed. Descriptions include primary and overspeed governors, autofeather system, and synchrophaser. Propeller system checks, as outlined in the POH/AFM Before Taxi and Before Takeoff (runup) checklists, are discussed.

This chapter also presents a description and discussion of the propeller system. Location and use

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Figure 7-1 PT6A-65B Powerplant Installation

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1900 AIRLINER POWERPLANT Beech designers chose the Pratt and Whitney PT6A-65B (UA, UB, UC) and -67D (UE) powerplants for the 1900 Airliner (Figure 7-1). The PT6A-65B reverse flow, free-turbine, turboprop engine (Figure 7-2) is flat-rated to 1,100 shaft horsepower. The PT6A-67D is flat-rated to 1,279 shaft horsepower.

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The engines are equipped with composite fourblade, full-feathering, reversing, constant-speed propellers mounted on the output shaft of the engine reduction gearbox. Engine oil supply and single-action, engine-driven governors control propeller pitch and speed. When the engines are shut down, propellers automatically feather, and will unfeather when engines are started as engine oil is pumped into the propeller dome. Reference to the right or to the left side of the aircraft, propellers, or engines always assumes the pilot is looking from the rear of the aircraft forward (Figure 7-3).

Figure 7-2 Engine Installation

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Turboprop engine power is measured in equival e n t s h a f t h o r s e p ow e r ( E S H P ) a n d s h a f t horsepower (SHP). SHP is determined by propeller rpm and torque applied to turn the propeller shaft. Hot exhaust gases leaving the engine also develop some kinetic energy similar to a turbojet engine. Jet thrust (approximately 10% of total

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engine power), combined with SHP, is equal to ESHP. Turboprop engine specifications usually show ESHP, SHP, and limiting ambient temperatures. The engine specification tables show engine ratings and temperatures (Tables 7-1 and 7-2).

Table 7-1 PT6A-65B SPECIFICATIONS (SEA LEVEL STATIC ICAO STANDARD ATMOSPHERE CONDITIONS) ESHP

SHP

PROPELLER RPM (A)

JET THRUST (LBS)

SPECIFIC FUEL CONSUMPTION (LB./ESHP/HR.)

Takeoff

1174

1100

1700

181 x 2

0.536

Max. Continuous

1174

1100

1700

181

0.536

Max. Climb +

1069

1000*

1700

171

0.550

Normal Cruise +

1069

1000*

1700

171

0.550

OPERATING CONDITION

* Available to 21.1˚C (70˚F) + For information only; not certification ratings (a) Corresponding Speed: Power Turbine - 29,920 rpm

Table 7-2 PT6A-67D SPECIFICATIONS (SEA LEVEL STATIC ICAO STANDARD ATMOSPHERE CONDITIONS) OPERATING CONDITION Takeoff

ESHP

SHP

PROPELLER RPM (A)

JET THRUST (LBS)

SPECIFIC FUEL CONSUMPTION (LB./ESHP/HR.)

1353

1279*

1700

186 x 2

0.530

Max. Continuous

1285

1214**

1700

178

0.539

Max. Climb +

1172

1106***

1700

165

0.557

Normal Cruise +

1172

1106***

1700

165

0.557

* ** *** + (a)

Available to 48˚C Available to 46.5˚C Available to 45˚C For information only; not certification ratings Corresponding Speed: Power Turbine - 29,920 rpm

Engine Type............................................................................................................... Free Turbine Type of Combustion Chamber...........................................................................................Annular Compression Ratio ..................................................................................................................12:1 Compressor Shaft Rotation (looking forward) ..................................................Counterclockwise Propeller Shaft Rotation (looking forward)...................................................................Clockwise Propeller Shaft Gear Ratio....................................................................................................17.6:1 Oil Consumption, Maximum Average......................(0.316 lb./hr. UE) 0.2 lb./hr. (0.0907 kg./hr.)

Gas Generator Speed 100% N1 ................................................................................... 37,468 rpm Max. Continuous Gas Generator Speed 104% N1 ...................................................... 39,000 rpm

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Figure 7-3 Engine Cutaway

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Figure 7-4 Engine Gas Flow

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ENGINE STATIONS

NF—Power turbine rpm (not indicated on engine instruments).

To identify locations in the engine, it is common practice to establish engine station numbers at various points (Figure 7-3). To refer to pressure or temperature at a specific point in engine airflow path, the appropriate station number is used, such as P3 for Station 3 pressure or T5 for gas temperature at Station 5. For instance, airflow temperature measured between the compressor and first-stage power turbine at Engine Station 5, is called interstage turbine temperature (ITT) or T5. Bleed air, located after the centrifugal compressor stage and prior to entering the combustion chamber, is commonly referred to as P3 or bleed air. Bleed air is used for cabin heat, pressurization, and the pneumatic system.

ENGINE TERMS To properly understand the operation of the PT6A series engines, you should know the following basic definitions. These terms should be memorized since they are used often when describing PT6A engines.

N1 or Ng—Gas generator rpm in percent of turbine speed.

Np—Propeller rpm.

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P2.5—Air pressure between Engine Stations 2 and 3. Also referred to as axial stage air or compressor interstage air.

P3—Air pressure at Engine Station 3, the source of bleed air used for some aircraft systems.

ITT or T5—Interstage Turbine Temperature in degrees centigrade at Engine Station 5.

Free-Turbine Reverse-Flow Principle The “free-turbine” design of the PT6A series engines refers to turbine sections which rotate freely, having no physical connection between them (Figure 7-5). The compressor turbine drives the engine compressor and accessories. Dual power turbines drive the power section and propeller through the planetary reduction gearbox.

Compressor and power turbines are mounted on separate shafts and are driven in opposite directions by gas flow across them. The term “reverse flow” refers to airflow through the engine. Inlet air enters the compressor at the aft end of the engine, moves forward through the combustion section and the turbines, and is exhausted at the front of the engine.

Figure 7-5 Free-Turbine Reverse Flow Principle

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Engine Modular Free-Turbine Design The modular concept is an important feature of the PT6A engines. As a modular free-turbine design, the engine is basically divided into two sections—a gas generator module and a power module (Figure 7-6). The gas generator module includes the compressor and the combustion section. Its function is to draw air into the engine, add energy to it in the form of burning fuel, and produce the gases necessary to drive the compressor and power turbines. The power module converts the gas flow from the gas generator into mechanical action to drive the propeller. This is done through an integral planetary gearbox, which converts the high-speed, low torque of the power turbine to low-speed, high torque required at the propeller. The reduction ratio from power turbine shaft rpm to propeller rpm is approximately 17.6:1.

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occurring at mid-TBO, involves splitting the engine between the compressor and power turbines. Since it is not necessary to remove the engine from the airplane to accomplish the HSI, inspection is both simple and fast. Because of modular design, the gas generator section or the combustion section can be completely replaced independently of each other. This feature permits easy maintenance, modular overhaul, and onwing HSI.

ENGINE AIRFLOW Another important feature of the PT6A engines is reverse-flow design. Inlet air enters the rear of the engine through an annular plenum chamber, formed by the compressor inlet case, where it is directed forward to the compressor (see Figure 7-4). The compressor consists of four axial stages and a single centrifugal stage assembled as a single unit on a common shaft.

The modular engine requires minimum maintenance. A hot section inspection (HSI), usually

Figure 7-6 Engine Modular Concept

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Rows of stator vanes between each stage of compression diffuse the air, raise its static pressure, and direct it to the next stage of compression.

After the final compression stage, compressed air passes through diffuser tubes which turn it 90Ëš in direction, and convert its velocity back to static pressure. Now diffused, the air passes through straightening vanes to the annulus surrounding the combustion chamber liner.

The flow of air changes direction 180Ëš as it enters and mixes with fuel in the combustion chamber.

The combustion chamber liner contains perforations of varying size that allow entry of compressor delivery air. Approximately 25% of the air mixes with fuel to support combustion.

The remaining 75% is used to center the flame in the combustion can and for internal engine cooling. The fuel/air mixture is ignited and the resultant expanding gases are directed to the turbines. The unique location of the combustion chamber liner, utilizing flow reversal, eliminates the need for a long shaft between the compressor and the compressor turbine, thus reducing the engine’s overall length and weight.

For smoother engine starts, the PT6A-65B fuel is introduced into the combustion chamber liner in two stages through 14 simplex fuel nozzles. The nozzles are supplied by a dual fuel manifold consisting of primary and secondary transfer tubes and adapters. The seven primary nozzles inject fuel into the combustion chamber during initial start, and the remaining secondary nozzles are activated as N1 increases to approximately 36%.

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In the PT6A-67D engine, fuel is introduced through 14 duplex nozzles. The nozzles are supplied by a dual fuel manifold consisting of primary and secondary transfer tubes and adapters. The 14 primary nozzles inject fuel into the combustion chamber during initial start, and the 14 secondary nozzles are activated as N 1 increases to approximately 40 to 45%. The fuel/air mixture is ignited by two spark igniters which protrude into the liner.

After combustion, expanding gases reverse direction in the exit duct zone, and pass through compressor turbine inlet guide vanes to the s i n g l e - s t a g e c o m p r e s s o r d r iv e t u r b i n e .

Expanding gases are then directed forward through power turbine inlet guide vanes to drive the power turbine section. The guide vanes ensure that expanding gases impinge on the turbine blades at the correct angle with minimum energy loss. Approximately 60% of the energy from combustion gases is extracted by the compressor turbine, and the balance is used by the power turbines. Exhaust gas from the power turbines is then directed to the atmosphere through abifurcated exhaust plenum to twin opposed-exhaust ports.

Compressor and power turbines are located in the approximate center of the engine, with their respective shafts extending in opposite directions. This feature simplifies installation and inspection procedures.

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COMPRESSOR BLEED VALVE At low N1 rpm, compressor axial stages produce more compressed air than the centrifugal stage can use. A compressor bleed valve compensates for excess air flow at low rpm by bleeding axial stage air (P2.5) to reduce back pressure on axial stages (Figure 7-7). Pressure relief helps prevent axial stage compressor stall.

The compressor bleed valve is a pneumatic piston that references pressure differential between axial and centrifugal stages. Looking forward, the valve is located at the 3 o’clock position on the gas generator case. The valve helps prevent

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compressor stalls and provides smoother compressor operation in the low N1 rpm range.

At low N1 rpm, the compressor bleed valve is open. As power is increased beyond approximately 72% N 1 , the valve begins to close progressively. At takeoff and cruise, above approximately 90% N1 rpm, the bleed valve is closed. If the compressor bleed valve were to stick closed at low N1 speeds, compressor stall could result from an attempt to accelerate the engine to higher power. If the valve were to stick open at high N1 speeds, power output would be considerably reduced. With the valve open, and

Figure 7-7 Jet-Flap, Compressor Bleed Valve, Swing Check Valve

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at a given N1 rpm, ITT will increase slightly and torque will decrease.

NOTE On UB-40 and after, the following changes have been made:

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SWING CHECK VALVE (SERIES UA THROUGH UB-40) A swing check valve is located on the left side of the compressor at the 9 o’clock position. It is a plate valve, hinged at the upper edge, capable of pivoting approximately 45˚. The valve relieves excess P2.5 pressure that is trapped in the compressor case when the compressor bleed valve closes.

●

Improved pre-swirl design supersedes jet flap.

●

Swing check valve is eliminated.

IGNITERS

●

Rolling diaphragm is eliminated.

Two spark-type igniters in the combustion chamber provide positive ignition during engine start.

Although the engine is equipped with two igniters, it needs only one for start. The system is designed so that if one igniter malfunctions, the remaining igniter will continue to operate. Igniters are activated by the IGNITION AND ENGINE START switch, and are turned off after engine start when combustion becomes selfsustaining.

JET-FLAP INTAKE SYSTEM A unique feature of the PT6A-65B engine is its efficient utilization of P2.5 air. In the 65B, air is ported into a “jet flap” system. A jet flap slot, which secures the accessory section to the engine compressor section, is machined into one side of each hollow strut. A jet flap intake system (Figure 7-7) functions as a variable inlet guide vane without variable geometry. Compressor interstage air (P 2.5 ) enters and exits through the narrow slot, passing into the intake to provide a swirl effect on inlet air entering the compressor.

This “pre-swirl” effect improves low-speed compressor characteristics and eliminates the requirement for an additional compressor bleed valve.

Spark ignition is effective for quick engine starting throughout a wide temperature range. The system consists of an airframe-mounted ignition exciter, two individual high-tension cable assemblies, and two spark igniters. It is energized from the aircraft nominal 28-volt DC supply, and will operate in the 9 to 30-volt range. The igniter system can produce up to 3,000 volts.

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Figure 7-8 Engine Start and Ignition Switches

Engine start switches on the pilot’s left subpanel (Figure 7-8) are placarded IGNITION AND ENGINE START. The three switch positions are ON, OFF, and STARTER ONLY. The leverlocked ON position activates the starter and both igniters. STARTER ONLY is a momentary holddown position of the spring-loaded center (OFF) position. STARTER ONLY provides a motoringonly function sometimes used to clear the engine of unburned fuel. Igniters are not energized in this position.

The ignition system features automatic capability. When the ENG AUTO IGNITION switches are in the lever-locked ARM position, igniters will automatically operate if engine torque falls below 500 to 550 foot-pounds, and IGNITION ON annunciators will illuminate. The auto-ignition system should be armed in turbulence, precipitation, and icing conditions. In

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the PT6A-67D engine, actuation will occur if the torque falls below 700 to 750 foot-pounds.

ACCESSORY SECTION All engine-driven accessories, except the propeller tachometer and propeller governors, are mounted on the accessory gearbox at the rear of the engine (Figure 7-9). Accessories are driven by the compressor shaft (N1) through a coupling shaft. One lubricating oil pressure pump and two scavenge oil pumps are mounted inside the accessory gearbox. Two additional oil scavenge pumps are externally mounted. The starter/generator, high-pressure fuel pump, N1 tachometer generator, and other optional accessories are mounted on pads on the rear of the accessory drive case. Each mounting pad has its own specific gear ratio.

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Figure 7-9 Typical PT6A Engine

LUBRICATION SYSTEM The PT6A engine lubrication system functions primarily to cool and lubricate engine bearings and bushings (Figure 7-10). It also provides oil to the propeller governor and propeller reversing control system. The main oil tank houses a geartype engine-driven pressure pump, an oil pressure regulator, a cold pressure relief valve, and an oil filter. The engine oil tank, an integral part of the compressor inlet case, is located in front of the accessory gearbox.

As oil is pumped from the tank, it passes through pressure- and temperature-sensing bulbs mounted on the rear accessory case. At gas generator speeds above 72% N1, normal oil pressure is between 90 and 135 psi. Oil is then delivered through an external oil transfer line below the engine to bearing compartments and to the nose case. Gear-type scavenge pumps return the oil through external oil transfer lines and through an external oil cooler below the engine.

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Figure 7-10 Engine Lubrication Diagram

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The oil cooler is thermostatically controlled to maintain desired oil temperature. When scavenge oil temperature reaches 71° C, a thermostatically controlled diverter valve opens to route oil through the cooler. Another externally mounted unit, the oil-to-fuel heat exchanger, uses hot engine oil to heat fuel before it enters the engine fuel system.

Total oil system capacity is 3.9 U.S. gallons, including the 2.3 gallon oil tank. Maximum oil consumption is 1 quart every 10 hours; however, normal oil consumption may be as little as 1 quart per 50 hours. Most PT6A engines normally seek an oil level of one to two quarts low. When adding oil between oil changes, do not overfill, and do not mix types or brands of oil due to the possibility of chemical incompatibility.

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A placard inside the engine cover shows the brand and type of oil used in that particular engine. Although the preflight checklist calls for checking oil level, the best time to check oil quantity is shortly after shutdown, since oil levels are most accurately indicated at that time. Oil level checks during preflight may require motoring the engine to obtain an accurate level indication.

The oil tank is provided with a filler neck and integral quantity dipstick housing. The cap and dipstick are secured to the filler neck, which passes through the gearbox housing and accessory diaphragm into the tank. Dipstick markings indicate the number of U.S. quarts of oil less than full (Figure 7-11).

Figure 7-11 Engine Oil Dipstick

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MAGNETIC CHIP DETECTOR (SERIES UA, UB) A magnetic chip detector (Figure 7-12) is installed in the bottom of each engine nose gearbox to indicate the presence of ferrous particles in the lubrication system. The detector activates a yellow light on the annunciator panel, L CHIP DETECT or R CHIP DETECT, to alert the pilot to possible oil contamination.

Illumination of the CHIP DETECT annunciator is not in itself cause for an engine to be shut down. Engine parameters should be monitored for abnormal indications. If parameters are

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abnormal a precautionary shutdown may be made at the pilot’s discretion.

ENGINE FUEL SYSTEM The engine fuel control system for PT6A-65B and -67D engines consists of the basic components shown in the block diagram (Figure 7-13). They are the primary low-pressure boost pump, oil-to-fuel heat exchanger, high-pressure fuel pump, fuel control unit, fuel cutoff valve, flow divider, and dual fuel manifold with 14 simplex nozzles, 14 duplex nozzles in the -67D. The low-pressure boost pump is engine driven, and operates when the gas generator shaft (N1) is turning. It provides sufficient fuel head pressure (approximately 45 psi maximum) for proper cooling and lubrication of the high-pressure pump. The oil-to-fuel heat exchanger regulates fuel temperature at the fuel pump inlet to prevent icing at the pump filter. This is done automatically and requires no action by the pilot.

After fuel passes through the oil-to-fuel heat exchanger, it flows into the high-pressure, engine-driven fuel pump and into the fuel control unit (FCU). Prior to entering the FCU, a fuel purge line constantly directs a small amount of fuel back to wing fuel tanks to clear vapors and bubbles from the fuel control system.

The high-pressure fuel pump is an engine-driven, gear-type pump that can supply fuel at 850 psi maximum pressure to the fuel side of the FCU. Its primary purpose is to supply sufficient pressure to fuel nozzles for adequate spray pattern during all modes of engine operation. Flow rates and pressures will vary with changes in gas generator (N1) rpm.

Figure 7-12 Magnetic Chip Detector

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The fuel cutoff valve is internal in the FCU. The valve is controlled by the condition lever, and is either open or closed; it has no intermediate position. When the fuel cutoff valve is open, fuel flows to the minimum pressurizing valve, which blocks fuel flow during start until fuel pressure is sufficient for proper spray pattern in the combustion chamber. As high-pressure fuel pump output

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Figure 7-13 Simplified Fuel System Diagram

increases to at least 100 psi, the minimum pressurizing valve opens, delivering fuel to the flow divider. If the high-pressure pump fails, the valve will close and combustion will cease.

During start, fuel flows initially through the flow divider to seven primary fuel spray nozzles in the combustion chamber. As the engine accelerates through approximately 36% N1 , fuel pressure increases sufficiently to supply secondary fuel nozzles. All 14 nozzles then deliver atomized fuel to the combustion chamber. The progressive sequence of primary and secondary fuel nozzle operation provides cooler starts. Increased acceleration in N 1 speed may be noticed when secondary fuel nozzles activate.

FUEL MANIFOLD PURGE SYSTEM The fuel manifold purge system is designed to eliminate residual fuel which remains in the flow divider and fuel manifold when the fuel cutoff valve is closed during engine shutdown. The system consists of a P3 accumulator purge tank with P3 air input at one end and P3 discharge to the flow divider at the other end.

D u r i n g n o r m a l e n g i n e o p e r a t i o n , P3 a i r constantly pressurizes the purge tank. As long as the engine is running, fuel pressure keeps the flow divider purge port closed. As fuel pressure drops to zero during engine shutdown, P 3 air escapes through a check valve into the flow

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divider, pushing residual fuel into the combustion chamber where it is burned. As a result, the pilot may notice a one- or two-second delay in initial engine spooldown after the condition levers are moved into fuel cutoff.

FUEL CONTROL UNIT (FCU) The fuel control unit (Figure 7-14) meters proper fuel amount for all modes of engine operation.

Flow rates are calibrated for starting, acceleration, and maximum power. The FCU compares gas generator speed (N1) with power lever setting and regulates fuel to engine fuel nozzles. The FCU also senses compressor discharge pressure and compares it to N1 rpm to establish acceleration and deceleration fuel flow limits. A minimum flow stop, set to approximately 90 pounds per hour per engine, guarantees sufficient fuel flow at all operating altitudes to sustain engine operation at minimum power.

FCU OPERATION The fuel control unit (Figure 7-14) is mounted on the rear flange of the fuel pump. A splined coupling between the pump and the FCU transmits a speed signal to the governing section in the FCU. The FCU determines the amount of fuel scheduled to the combustion chamber by controlling gas generator speed. Engine power output is directly dependent upon gas generator speed.

Compressor discharge pressure (P3), sensed by the fuel control unit, is used to establish acceleration fuel flow limits. This fuel limiting function prevents overtemperature conditions in the engine during start and acceleration.

The fuel control unit receives input from the condition lever, the power lever, the N1 flyweight governor, and a pneumatic bellows. FCU operation is complex, but it will be simplified and described briefly in this section. For more detail refer to the Pratt & Whitney Maintenance Manual for the PT6A-65B or PT6A-67D engines.

Power control levers position a 3-D cam in the FCU that, through a cam follower and lever, determines fuel flow corresponding to selected N1 speed. The condition lever selects LOW IDLE

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to HIGH IDLE N1 speeds when the power lever is at idle. The power lever selects speeds between idle and maximum.

Engine speed is controlled by the gas generator (N1 ) governor, which contains two flyweights mounted on a ballhead driven by the engine. The flyweight governor is the feedback element of the speed-select system. It controls “on-speed� condition by positioning the 3-D cam as required in response to speed variations in the gas generator.

As N1 speed increases or decreases, resulting flyweight action changes the 3-D cam setting, which changes fuel flow valve setting to maintain selected N1 speed.

The cam follower and arm transmit 3-D cam motion to the fuel metering valve. As the 3-D cam moves upward, fuel flow to the engine is increased and N1 speed increases. Downward movement of the 3-D cam decreases fuel flow and N1 speed. N1 speed is therefore maintained continuously by the N1 governor in response to variations in gas generator speed.

In an overspeed condition, increasing pressure by the governor flyweights moves the 3-D cam downward, resulting in decreased fuel flow through the fuel metering valve. Balance occurs when N1 speed is reduced to selected speed, and the cam is stationary at the new speed position.

In an underspeed condition, decreasing pressure by the governor flyweights moves the 3-D cam upward, resulting in increased fuel flow through the fuel metering valve until the system reaches equilibrium.

Compressor discharge pressure (P3 air) also affects fuel metering valve position during acceleration or deceleration. Increase in P3 causes the fuel metering valve to increase fuel flow in response to increased P3 pressure until N1 speed is stabilized.

A decrease in P3 causes the fuel metering valve to decrease fuel flow until N1 speed is stabilized at the lower selected valve. In the event of power turbine overspeed, a decrease in P3 air pressure at the fuel metering valve allows the fuel control unit to reduce fuel flow to the gas generator. Overspeed protection is discussed in greater detail in the Propeller System section at the end of this chapter.

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FUEL FLOW INDICATORS Fuel flow information is sensed by a transmitter in the engine fuel supply line between the fuel cutoff valve and the flow divider. Flow rate is indicated on the fuel flow gage on the instrument panel (Figure 7-15). Indications are in pounds-per-hour units times 100; therefore, when the needle indicates 2 on the dial, fuel flow is 200 pounds per hour. Fuel flow gages are DC-powered.

Figure 7-15 Fuel Flow Gages

Figure 7-14 Simplified Fuel Control System

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FUEL PRESSURE INDICATORS If the engine-driven, high-pressure fuel pump fails, combustion will cease since high-pressure fuel is required to open the minimum pressurizing valve.

If the primary engine-driven boost pump fails, a FUEL PRESS red light in the warning annunciator panel will illuminate (Figure 7-16), and the master warning lights will flash. The FUEL PRESS light illuminates when outlet pressure at the engine-driven boost pump decreases below one psi (10 psi Series UC & after). Switching on the standby fuel boost pump should increase fuel pressure to more than 11 ±2 psi, extinguishing the warning light. Engine operation with the FUEL PRESS light on is limited to 10 hours between overhaul or replacement of the enginedriven, high-pressure fuel pump.

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graph is based on known or forecast operating conditions, and shows operating temperatures where icing at the FCU could occur. Refer to the 1900 Airliner Maintenance Manual for procedures to be followed when blending anti-icing additive with airplane fuel. Use anti-icing additive conforming to Specification MIL-I-27686.

FUEL BIOCIDE ADDITIVE Fuel biocide-fungicide “BIOBOR JF” in concentrations of 135 or 270 parts per million may be used in the fuel. BIOBOR JF may be used as the only fuel additive, or it may be used with MIL-I27686 anti-icing additive. Used together, additives have no detrimental effect on fuel system components. Refer to 1900 Airliner Maintenance Manual for concentrations and detailed procedures for adding BIOBOR JF to airplane fuel.

CONTROLS AND INDICATIONS CONTROL PEDESTAL The control pedestal extends between the pilot and copilot (Figure 7-17). The three sets of powerplant control levers, from left to right, are power levers, propeller rpm and feather levers, and condition levers.

Figure 7-16 Fuel Pressure Annunciator

ANTI-ICING FUEL ADDITIVE Anti-icing fuel additives are not normally required since engine oil heats fuel before it enters the FCU; however, if oil temperature versus OAT indicates ice formation could occur during takeoff or in flight, anti-icing additive should be mixed with the fuel to ensure safe operation. For preflight planning purposes, use the Minimum Oil Temperature Required for Operation without Anti-Icing Additive graph.

This graph can be found in the POH or in the Ice and Rain Protection section of this manual. The

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ENGINE POWER CONTROL Engine power is controlled by power levers which set N 1 speed, and by propeller levers which adjust propeller speed (Figure 7-18). The propeller maintains set speed by varying blade angle as requested by the propeller governor in response to changes in torque. Torque changes result from power lever input to the N1 governor.

When power lever position calls for more torque, N1 governor settings prevent bleed-off of internal fuel pressures and of P3 air in the FCU. The governor then signals the fuel metering valve to allow more fuel flow into the spray nozzles to meet requested power conditions.

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detent. All ground operation should be conducted with the propeller levers in this position.

Although the 1900 Airliner is equipped with an automatic propeller feathering system, the propeller can be manually feathered by pulling the propeller lever back past the friction detent into the red and white striped section of the quadrant.

To unfeather, push the lever back into the governing range. Regardless of propeller lever position, the propellers will move toward feather when oil pressure is lost as the engines are shut down. For further details, see the Propeller System section at the end of this chapter.

CONDITION LEVERS Figure 7-17 Control Pedestal

POWER LEVERS Power levers control engine power from idle to takeoff power by operation of the N1 governor in the fuel control unit (Figure 7-18). Increasing N1 rpm results in increased engine power. The power levers control power in three regions: FLIGHT, BETA, and REVERSE. The bottom of the flight range is called IDLE. When power levers are lifted over the IDLE detent and pulled back into the beta range, they control propeller blade angle only. The beta range is normally used for taxi. The bottom of the beta range is called GROUND FINE. When the levers are lifted over the GROUND FINE detent into the REVERSE range, they control propeller blade angle and engine power to provide reverse thrust.

PROPELLER LEVERS The propeller levers are conventional in setting required rpm for takeoff and cruise (Figure 7-18).

T h e n o r m a l g ove r n i n g r a n g e i s 1 4 0 0 t o 1700 rpm.

However, in the 1900D (UE Series), the minimum governing range is 1200 rpm, and ground taxi position is marked on the power quadrant just before the propeller levers reach the feather

The condition levers have three positions: FUEL CUTOFF, LO IDLE, and HI IDLE (Figure 7-18).

In the FUEL CUTOFF position, all fuel flow to the engines is cut off.

At LO IDLE, engine gas generator speed (N1) is a minimum of 58%; N1 at HI IDLE is 70%. Condition levers can be set between these two values for any speed between 58% and 70% N1. For the -67D engines these figures are 65% for LOW IDLE and 71% for HI IDLE.

CONDITION LEVER OPERATION (SERIES UA, UB, UC) Engines are controlled from the cockpit by using the propeller, power, and condition levers. Both the power and condition levers are connected to the N 1 governing section of the FCU. Either lever will reset the FCU to maintain a new N1 rpm. During start, power levers are at IDLE. Once condition levers are moved to LO IDLE, the fuel cutoff valve opens and the N1 governor is set at LO IDLE. The condition levers are continuously variable from LO IDLE at 58% to HI IDLE at 70% N1. When condition levers are set to LO IDLE, the power levers will select N1 rpm from 58% to 104%, the maximum for takeoff; however, if condition levers are at HI IDLE, the power levers can only select N1 rpm from 70% to 104%.

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Moving either the power levers or condition levers changes only N1 rpm. As the power or condition levers are advanced, ITT, torque and fuel flow increase; however, these increases are byproducts of the N1 speed maintained by the FCU. With power levers in a fixed position, N1 remains constant even in climb or descent, but ITT, torque and fuel flow will vary with altitude, ambient air temperature, and propeller setting.

ITT AND TORQUEMETERS The two primary operating parameters which limit available engine power are temperature and torque. Engine torque and ITT are affected by ambient temperature and altitude. During operations requiring maximum engine performance at cold temperature or low altitude, torque limits power. When ambient temperature is hot or when operating at high altitude, ITT limits power.

Whichever condition reaches its limit first, determines the amount of power available.

ITT GAGE

Figure 7-18 Control Levers

CONDITION LEVER OPERATION (SERIES UE) In the PT6A-67D equipped airplanes, the LO IDLE setting is 65%. HI IDLE is 71% with the maximum power setting being 104%.

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The ITT gage monitors interstage turbine temperature at Station 5 (Figures 7-19, 7-20, and 7-21). ITT is a prime limiting indicator of power available under varying ambient temperature and altitude conditions. The normal operating range, indicated by the green arc on the gage, is 400 to 810° C (400 to 780° C—UE). Maximum starting-only temperature of 1000°C, indicated by the solid white line (red diamond on Series UE and after) on the instrument, is limited to five seconds. Maximum continuous temperature (red line) is 810° C (800° C—UE), and maximum cruise temperature is 750° C. Engines can be damaged if limiting temperatures indicated on the ITT gage are exceeded. Temperature limitations for all operating conditions are listed in the POH or AFM, and should be committed to memory. ITT gages are self-generating and do not require electrical power. Starting with UE-93, this gage will be powered by 28 VDC plus the electrical signal from the thermocouples installed in the engine.

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TORQUEMETER (SERIES UA, UB, UC) The torquemeter, marked with the green arc from zero foot-pounds to 3,400 foot-pounds, constantly measures rotational force applied to the propeller shaft (Figures 7-19, 7-20, and 7-21).

Maximum permissible sustained torque is 3,400 foot-pounds, indicated by the red radial on the instrument. Torque is measured by a hydromechanical torquemeter in the first-stage reduction gearcase. Rotational force on the first-stage ring gear compresses oil in the torquemeter chamber.

The difference between torquemeter chamber pressure and reduction gear internal pressure accurately indicates the torque produced at the propeller shaft. The torquemeter transmitter measures this torque and sends an AC electrical signal to the torquemeter on the cockpit instrument panel. Recommended cruise torques vary with altitude and temperature.

TORQUEMETER (SERIES UE) The normal operating range for the PT6A-67D engine is from 0 to 3,750 foot-pounds (green arc) with the maximum permissible torque set at 3,950 foot-pounds and with a time limit of 5 minutes (yellow arc). The torque reading is displayed on 26-VAC powered gages. Starting with UE-93, this gage will be changed to a 28-VDC stepper motor gage.

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GAS GENERATOR (N1) TACHOMETER (SERIES UA, UB, UC) The gas generator (N1 ) tachometer measures rotational speed of the compressor shaft, in percent rpm, based on 37,468 rpm at 100% (Figures 7-19, 7-20, and 7-21). The outer scale of the indicator is measured in units of ten, and the smaller inner dial is read in single units. The N1 indicator is self-generating, requiring no electrical power for operation. The tachometer generator, located on the engine accessory section, is geared directly to the N1 compressor shaft. The tachometer generator sends an electrical signal to the N1 indicator on the cockpit instrument panel, indicating the percentage of N1 rpm. Maximum allowable continuous gas generator speed is 39,000 rpm, read as 104% on the N1 indicator.

GAS GENERATOR (N1) TACHOMETER (SERIES UE) The indicator for the 1900D (UE Series) combines an analog and digital read-out. This gage is powered by 28 VDC plus a signal from the N1 tachometer generator located in the engine accessory section.

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Figure 7-19 Engine Instrument Markings (Series UA, UB, and UC)

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Figure 7-20 Engine Instruments (Series UE-1 through UE-92)

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Figure 7-21 Engine Instruments (Series UE-93 and After)

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ENGINE LIMITATIONS Airplane and engine limits are described in the Limitations section of the POH or AFM. These limitations have been approved by the Federal Aviation Administration, and must be observed when operating the 1900 Airliner. The Engine Operating Limits Chart provides important limitations for all operating conditions. The Power Plant Instrument Markings Chart lists minimum, normal, and maximum limits (Figures 7-22 and 7-23).

NOTE Engine limits charts are reproduced from applicable POH or AFM.

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During engine start, temperature is the most critical limit (Figure 7-24). The ITT starting limit of 1,000° C is limited to five seconds. During any start, if the indicator needle approaches the limit, start should be aborted before the needle passes the solid white line. For this reason, it is helpful during starts to keep the condition lever out of the LO IDLE detent so that the lever can be quickly pulled back to FUEL CUTOFF. Monitor oil pressure and oil temperature gages to further determine engine condition. During start, oil pressure should come up quickly to the minimum red line at 60 psi, but should not exceed the maximum 200 psi. During normal operation oil temperature and pressure gages should be in the normal operating range, indicated by a green arc from 90 to 135 psi. Fluctuations of ±10 psi are acceptable, but pressures between 60 and 90 psi are undesirable. Low oil pressures (60 to 90 psi) should be tolerated only for completion of the flight, and then only at a reduced power setting.

Oil pressure below 60 psi is unsafe. Below 60 psi, the engine should be shut down, or a landing should be made as soon as possible using the minimum power required to sustain flight.

Normal operating oil temperatures are limited to 0 to +110° C. A minimum of -40°C is required for engine start, and temperature limits at idle are -40 to +110° C. However, temperatures between +99 and +110° C are limited to a maximum of 10 minutes.

During ground operations, ITT must remain below 700° C (750° C—Series UE). Engine temperatures can be controlled by regulating N1 rpm and generator load. When condition levers are at LO IDLE, high ITT can be corrected by first reducing N1 loads, such as generators and air conditioning, then by advancing condition levers to increase N1 rpm. HI IDLE will reduce ITT since increased compressor speed increases the amount of cooling airflow available to the engine. Once ITT has been reduced below the idle temperature limit, N1 loads may be restored as desired. During normal flight operations, ITT should never be allowed to exceed the maximum continuous limit (red line).

Figure 7-22 Engine Limits Chart—PT6A-65B

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Figure 7-23 Engine Limits Chart—PT6A-67D

Figure 7-24 Overtemperature Limits (Starting)

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During the climb, available torque will decrease and ITT may increase slightly. Recommended cruise climb and normal cruise ITT limit is shown by a white triangle on the indicator.

Torque and N1 limitations are the same for maxim u m c r u i s e a n d f o r t a ke o ff ; h ow ev e r, recommended cruise torque values will vary with altitude and temperature.

Transient limits provide buffers for surges during engine acceleration. Torque and ITT have an allowable excursion duration of 20 seconds. A momentary peak of 5,000 foot-pounds and 8 7 0 ° C i s a l l ow e d f o r t o r q u e a n d I T T respectively.

The overtemperature chart (Figure 7-24) shows specific actions required if ITT limits are exceeded during start. If engine temperature limits in Area A have been exceeded, determine and correct the cause of the overtemperature. Then the engine must be visually inspected through the exhaust duct (Figure 7-25), and any action taken must be recorded in the engine log book.

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engine must be returned for overhaul. Exceeding ITT limits in Area C for more than a few seconds may cause extensive engine damage.

The POH lists generator limitations for operation at various N1 rpms. To remain within limits it may be necessary to reduce accessory loads or to increase N1 rpm. LOW IDLE is sufficient for generator loads of 50% and lower with air conditioning off. With air conditioning on, or with loads higher than 50%, N1 rpm must be increased as specified in the limitations.

STARTER OPERATING TIME LIMITS Engine starters are time-limited during the starting cycle to prevent the possibility of starter damage due to overheating. The starter is limited to 30 seconds ON and 3 minutes OFF for cooling before the next sequence of 30 seconds ON. After the second cycle of 30 seconds ON, the starter must remain OFF for 30 minutes.

Starting with Series U C-143 and after and those airplanes that have been modified, the starter duty limitation has been changed to 20 seconds ON, 30 seconds OFF; 20 seconds ON, 60 seconds OFF; 20 seconds ON, 5 minutes OFF. For continuous motoring without engine starting, the limit is set at 20 seconds ON, 5 minutes OFF.

TREND MONITORING

Figure 7-25 View through Exhaust Duct

Overtemperature in Area B will require a hot section inspection. During a hot section inspection, combustion chamber and turbine areas and components are examined and replaced as necessary.

If an overtemperature occurs in Area C, the

During normal operations, gas turbine engines are capable of producing rated power for extended periods of time. Engine operating parameters, such as output torque, interstage turbine temperature compressor speed, and fuel flow for individual engines are predictable for specific ambient conditions. On PT6A engines, these predictable characteristics may be taken advantage of by establishing and recording individual engine performance parameters.

Parameters can then be compared periodically to predicted values for day-to-day confirmation of engine efficiency.

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The engine condition trend monitoring system recommended by Pratt & Whitney is a process of periodically recording engine instrument readings such as torque, interstage turbine temperature, compressor speed, and fuel flow.

Readings are corrected for altitude, outside air temperature, and airspeed, if applicable, then compared to a set of typical engine characteristics to determine deviations.

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DATA COLLECTION The trend monitoring procedure specifies that flight data be recorded on each flying day, every five flight hours, or other predetermined flight period. Select a flight with long established cruise, preferably at a representative altitude and airspeed. With engine power established and stabilized for a minimum of five minutes, record the requested data on a form similar to the in-flight engine data log shown (Figure 7-26).

Figure 7-26 In-Flight Engine Data Log

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PROPELLER SYSTEM GENERAL Each engine is equipped with a composite, fourblade, counterweighted, full-feathering, variablepitch, constant-speed, reversing propeller mounted on the output shaft of the reduction gearbox (Figure 7-27). Since the engines are free turbines, with no mechanical connection between compressor and power turbines, the propeller can rotate freely on the power shaft when the engine is shut down. Propeller tiedown boots (Figure 7-28) are provided to prevent windmilling at zero oil pressure when the airplane is parked.

Figure 7-28 Propeller Tiedown Boot Installed

Figure 7-27 Hartzell Propeller

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Figure 7-29 Propeller Blade Angle Diagram

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Propeller pitch (Figure 7-29) and speed are controlled by engine oil pressure supplied to the propeller dome through engine-driven propeller governors. A governor oil pump boosts oil pressure delivered by the engine oil system to a pressure high enough to control movement of the propeller blades. When oil pressure is present in the propeller dome, propeller pitch (blade angle) is controlled normally by the propeller governor or by the beta valve, depending upon the propeller’s mode of operation. As oil pressure increases, the propeller moves toward low pitch (high rpm). Loss of oil pressure will cause centrifugal counterweights and feathering springs to move propeller blades toward high pitch (low rpm) and, eventually, into the feathered position.

As oil pressure decreases during engine shutdown, the propeller automatically moves toward feather.

The minimum low pitch propeller position is determined by a mechanically actuated hydraulic stop, referred to as the primary low pitch stop.

The power levers control beta and reverse blade angles by adjusting the low pitch stop position in beta and reverse ranges.

Two governors (a primary governor and an overspeed governor) control propeller rpm. The primary governor controls the propeller through its normal governing range. The propeller control lever selects propeller rpm by adjusting the primary governor condition. Should the primary governor malfunction, the overspeed governor prevents propeller speed from exceeding 1,768 rpm (1,802 rpm in the 1900D [UE Series]). The fuel topping governor acts as a backup governor, limiting propeller speed to 106% of that selected by the propeller lever. In the reverse range, the fuel topping governor is reset, limiting propeller rpm to approximately 96% of the primary governor setting. The fuel topping governor limits propeller rpm by reducing fuel flow to the engine.

The propeller rpm is displayed in the cockpit on a gage that receives its input from the propeller tachometer located on the right side of the engine crankcase. On the 1900D (UE Series) model, this tachometer signal is fed to a gage that is powered by 28 VDC.

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BLADE ANGLE Blade angle is the angle between the chord of the propeller and the propeller’s plane of rotation.

Because of the normal twist of the propeller, blade angle is different near the hub than it is near the tip. Blade angle for the 1900 Airliner is measured at the chord, 42 inches from the propeller’s hub. This position is referred to as the “42-inch station.” All blade angles specified in this section are approximate values.

PRIMARY GOVERNOR The primary governor modulates oil pressure in the propeller dome to change blade angle to maintain a constant propeller speed. As oil pressure in the dome changes, propeller blade angles change to maintain the propeller speed the operator has selected. The primary governor can maintain any selected propeller speed from approximately 1,400 to 1,700 rpm (1,200 to 1,700 rpm for UE).

For example, suppose an airplane is in normal cruising flight with the propeller turning at 1,550 rpm. If the pilot begins a descent without changing power, the airspeed will increase. This decreases the angle of attack of the propeller blades, causing less drag on the propeller, thus beginning to increase its rpm. If this propeller has variable pitch capabilities and is equipped with a governor set at 1,550 rpm, the governor will sense this “overspeed” condition (Figure 7-30) and increase blade angle to a higher pitch. The higher pitch increases the blade’s angle of attack, slowing it back to 1,550 rpm, or “onspeed.”

Likewise, if the airplane changes from cruise to climb airspeeds without a power change, the propeller rpm tends to decrease. The governor responds to this “underspeed” condition by decreasing blade angle to a lower pitch, and the rpm returns to its original value. Thus the governor gives “constant speed” characteristics to the variable pitch propeller.

Power changes, as well as airspeed changes, cause the propeller to momentarily experience overspeed or underspeed conditions, but again

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the governor reacts to maintain the onspeed condition. Due to the smooth action of the governor, the pilot will notice few, if any, of these minor adjustments.

There are times, however, when the primary governor is incapable of maintaining selected rpm.

For example, imagine an airplane approaching to land with its governor set at 1,700 rpm. As power and airspeed are both reduced, underspeed conditions exist which cause the governor to decrease blade angle as it attempts to restore the onspeed condition. If blade angle were allowed to decrease to its full reverse limit, aircraft control would be dramatically reduced. In such a situation the propeller, acting as a large disc, would create excessive drag, disturbing airflow around tail surfaces. A rapid nosedown pitch change would result.

To prevent these undesirable flight characteristics, a device is provided to stop the governor from selecting blade angles that are too low for safety. As blade angle is decreased by the governor, eventually the low pitch stop is reached.

Blade angle then becomes fixed, preventing its continued movement toward a lower pitch. At the low pitch stop, the governor is prevented from restoring the onspeed condition, and propeller rpm decreases below the selected governor rpm setting. Once the low pitch stop is reached, blade angle cannot decrease further until the pilot selects beta or reverse.

Primary Governor Operation The propeller levers adjust the primary propeller governor between 1,400 and 1,700 rpm (1,200 and 1,700 rpm for UE). The primary propeller governor, mounted at the top of the engine reduction gearbox, has two functions: it selects constant propeller rpm, and it can also feather the propeller. The primary propeller governor adjusts propeller rpm by controlling the oil supply to the propeller mechanism.

Figure 7-30 Primary Governor Diagram

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An integral part of the primary propeller governor is the governor pump. This pump is driven by the N P shaft and raises engine oil pressure to approximately 375 psi (750 psi for UE) for use

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by the propeller governing system. A transfer gland surrounds the NP shaft, allowing oil to continuously enter and exit the propeller dome area.

Thus, the transfer gland constantly supplies fresh, warm oil to the propeller pitch control mechanism. Propeller control is a balancing act of opposing forces. Although the feathering springs and centrifugal counterweights constantly exert force on the propeller to drive the blade angle toward high pitch and feather, oil pressure constantly attempts to maintain low pitch and high rpm.

For propeller speed reference, the primary propeller governor uses a set of rotating flyweights that are geared to the propeller shaft. The flyweights compare governor set speed (set with propeller levers) to actual propeller rotational speed. These flyweights are connected to a free-

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floating pilot valve. The position of the pilot valve determines how much oil pressure will be supplied to the propeller pitch control mechanism. The slower the flyweights are turning in relation to the desired reference speed, the lower the position of the pilot valve. If the propeller and flyweights turn faster than governor set speed, the resulting centrifugal force causes the pilot valve to rise inside the governor, decreasing the oil pressure. Here are a few examples.

If a propeller rpm of 1,550 is selected and the propeller is actually turning at 1,550, the flyweights are in their center, or “onspeed” condition (Figure 7-31). The pilot valve is also in the center position. This “onspeed” configuration provides constant oil pressure to the propeller pitch mechanism, which therefore maintains constant pitch and constant rpm.

Figure 7-31 Propeller Onspeed Diagram

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If the airplane begins a descent, or if engine power is increased without any change to the propeller levers, airspeed will tend to increase, causing the propeller to turn faster (Figure 7-32).

The flyweights will, in turn, rotate faster. The additional centrifugal force will cause the pilot valve to rise. Notice that oil can now escape through the pilot valve. The decreased oil pressure results in a higher pitch and in a reduction of propeller rpm, returning the propeller to its original rpm setting. As propeller speed decreases, the flyweights slow, and the pilot valve returns to the equilibrium position to maintain selected propeller rpm.

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If the airplane enters a climb, or if engine power is decreased without any change in propeller controls, airspeed will decrease, and the propeller will tend to slow (Figure 7-33). The flyweights in the propeller governor will slow down as propeller rpm decreases, and the pilot valve will lower.

When pilot valve position moves down, more oil pressure is supplied to the propeller pitch mechanism. Higher oil pressure results in lower propeller pitch, causing an increase in propeller rpm. As the propeller increases to its original rpm setting, the flyweights speed up, and the pilot valve returns to its equilibrium or “onspeed� position.

Figure 7-32 Propeller Overspeed Diagram

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The flyweights and pilot valve continuously make small adjustments which change propeller blade angles, maintaining constant propeller rpm.

The cockpit propeller lever adjusts where the equilibrium or “onspeed” condition will occur.

The pilot can select any constant speed from 1,400 to 1,700 rpm. Normally 1,700 rpm is used for takeoff and for initial climb. Cruise climb, maximum range power, and recommended cruise power are set at 1,550 rpm.

The propeller primary governor cannot maintain selected propeller rpm if power and airspeed are reduced below the governing range. For example, with the progressive reduction of power and airspeed during final approach, the propeller and rotating flyweights will tend to go to the under-

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speed condition. In underspeed the pilot valve drops, increasing oil pressure to the dome, and decreasing propeller pitch as power and airspeed are reduced. Since a reversible propeller is capable of continued decrease into negative or reverse blade angles, a low pitch stop is necessary to prevent the blade angle from decreasing beyond a predetermined value of 13° (12.7° for UE). When the propeller governor becomes incapable of maintaining the onspeed condition, the propeller rpm will fall below the selected governor rpm setting, indicating that the low pitch stop has been reached. In the event of a governor control linkage failure, an external spring on top of the governor will move the governor set speed to 1,700 rpm.

Figure 7-33 Propeller Underspeed Diagram

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LOW PITCH STOP WHENEVER PROPELLER RPM IS BELOW SELECTED GOVERNOR RPM, THE PROPELLER BLADE ANGLE IS AT THE LOW PITCH STOP (assuming the propeller is not feathered or in the process of being feathered).

NOTE The preceding statement assumes that momentary periods of underspeed are not being considered. Rather, the propeller rpm is below and staying below the selected governor rpm. For example, if the propeller control is set at 1,550 rpm but the propeller is turning at less than 1,550 rpm, the blade angle is at the low pitch stop.

On many airplanes, the low pitch stop is simply at the low pitch limit of travel, determined by the propeller’s construction. But with a reversing propeller, extreme travel in the low pitch direction is past 0°, into reverse or negative blade angles. Consequently, the 1900 Airliner’s propeller system has been designed to allow the low pitch stop to be repositioned when reversing is desired.

The low pitch stop is created by mechanical linkage which senses blade angle. At the low pitch stop the linkage causes a valve to close, stopping the flow of oil into the propeller dome. Since more oil causes low pitch and reversing, blocking off oil flow creates a low pitch stop. The low pitch stop valve, commonly referred to as the “beta” valve, is quite positive in its mechanical operation. Furthermore, the valve is springloaded to provide redundancy in the event of mechanical loss of beta valve control.

Low Pitch Stop Operation Low pitch propeller position is determined by a mechanically monitored hydraulic stop. The propeller servo piston is connected by four springloaded sliding rods to the slip ring mounted behind the propeller. A carbon brush block riding on the slip ring transfers the movement of the slip

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PILOT TRAINING MANUAL

ring through the propeller reversing lever to the beta valve of the governor. The initial forward motion of the beta valve blocks off the flow of oil to the propeller. Further motion forward dumps the oil from the propeller into the reduction gearbox sump. A mechanical stop limits the forward motion of the beta valve. Rearward movement of the beta valve does not affect normal propeller control. When the propeller is rotating at a speed lower than that selected on the governor, the governor pump provides oil pressure to the servo piston and decreases the pitch of the propeller blades until the feedback of motion from the slip ring pulls the beta valve into a position blocking the supply of oil to the propeller, thus preventing further pitch changes.

BETA AND REVERSE CONTROL The position of the low pitch stop is controlled from the cockpit by the power lever. Whenever the power lever is at IDLE or above, this stop is set at 13° (12.7° for UE) blade angle in flight and 7° (+4.7° for UE) blade angle on the ground.

Bringing the power lever aft of IDLE progressively repositions the low pitch stop to lower blade angles (Figure 7-34).

The geometry of the power lever linkage through the cam box is such that power lever increments from idle to full forward thrust have no effect on the beta valve’s position. When the power lever is moved from idle into the reverse range, it repositions the beta valve to direct governor pressure to the propeller piston, decreasing blade angle through zero into a negative range. The travel of the propeller servo piston is fed back to the beta valve to null its position and, in effect, to provide infinite negative blade angles all the way to MAXIMUM REVERSE. The opposite will occur when the power lever is moved from full reverse to any forward position up to idle, thus providing the pilot with manual blade angle control for ground handling.

The region between GROUND IDLE and GROUND FINE is referred to as the “beta for taxi,” or simply “BETA” range. In this range, the engine’s compressor speed (N1) remains at the

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Figure 7-34 Beta and Reverse Control

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Figure 7-35 Beta Range and Reverse Diagram

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value selected when the power lever was at IDLE (58 to 70% [65 to 7l% for UE] based on condition lever position).

To enter the BETA range, the power lever must be lifted beyond the IDLE stop and then moved aft. The aft stop of BETA range is called GROUND FINE. With aft movement of the power levers, blade angle moves progressively from GROUND IDLE to GROUND FINE. The region between GROUND FINE and MAXIMUM IDLE is referred to as “beta plus power” or simply REVERSE range. In this range, N1 progressively increases to a maximum value of 85 ±3% while the blade angle decreases. To enter the REVERSE range, the power lever must be lifted beyond the GROUND FINE stop and moved aft. With aft movement of the power levers, the blade angle will progressively decrease from GROUND FINE to MAXIMUM REVERSE at REVERSE.

Beta and Reverse Control Operation When propeller blade angle reaches approximately 18 ±1°, the four flanges extending from the dome make contact with four beta nuts (Figure 7-35). As propeller pitch angle continues to decrease, each flange on the propeller dome pushes each beta nut and attached polished rod forward. As the rod moves forward, it pulls the feedback ring forward. In turn a beta valve inside the governor is pulled into the oil cutoff position.

The linkage is set to cut off oil supply to the dome when blade angle reaches 13° (12.7° for UE). This provides the governor with a hydraulic low pitch stop of 13° (12.7° for UE) for in-flight operations.

If the low pitch stop were fixed at 13° (12.7° for UE), the propeller could not enter the beta and reverse ranges; however, the low pitch stop can be reset to allow the propeller to operate in the beta and reverse ranges while the aircraft is on the ground.

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When the power levers are lifted up and over the idle detent into the beta range, they are pulling back on the top of the reverse lever. As the reverse lever moves back, the beta valve is pushed back, reestablishing oil flow to the propeller dome. This allows propeller blade angle to go below the low pitch stop. As the propeller blades go below the low pitch stop, the propeller dome and feedback ring continue forward, eventually pulling the beta valve back into the oil cutoff position.

In summary, the position of the low pitch stop is controlled by the power levers. When the power levers are set at idle or above, the stop is set at 13° (12.7° for UE). When the power levers are moved aft of idle, the low pitch stop is repositioned to blade angles of less than 13° (12.7° for UE).

FLIGHT AND GROUND LOW PITCH STOPS Both a flight low pitch stop and a ground low pitch stop are utilized in this propeller system. In flight, the minimum blade angle is approximately 13° (12.7° for UE) with the power levers at idle.

When the wheels touch the ground upon landing, the propeller low pitch blade angle automatically changes from 13° to +7° (+12.7 to +4.9° for UE) through the action of a ground idle solenoid. The ground idle solenoid acts on the top of the reverse lever in the same manner as the power levers when selecting beta and reverse (Figure 7-36).

The right landing gear safety switch controls the ground low pitch stop. During the landing flare at idle power, the propeller blade angle will be approximately 13° (12.7° for UE). Upon touchdown the landing gear safety switch causes the propeller blade angle to immediately decrease to approximately +7° (+4.9° for UE). The propeller blade angle remains at +7° (+4.9° for UE) until the power levers are retarded behind the idle stop. As a backup for the landing gear safety switch, switches installed in the power quadrant ensure that the ground idle solenoid is activated by the time the power levers enter the beta range.

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Figure 7-36 Propeller Postioning Diagram (Sheet 1 of 2)

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Both flight and ground low pitch stops are set by the beta valve. Upon touchdown, the reverse lever position is electrically reset, causing the beta valve to open slightly. As the beta valve opens, more oil flows to the propeller dome, and blade pitch is reduced to the ground idle low pitch stop position. The electric circuits of the idle low pitch stops are protected by the PROP GOV TEST circuit breaker on the copilot’s subpanel (Series UA, UB, UC).

Keep in mind that although the ground low pitch stop is set at less than 13°, the actual blade angle is only affected when the propeller is at the low pitch stop. It follows, then, that as long as the propeller rpm is still on the selected governor setting, bringing the power lever aft of IDLE will not necessarily cause the propeller to reverse.

Reverse thrust can only occur when propeller rpm is below selected governor rpm (at the low pitch stop).

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In a condition of constant underspeed, the primary governor cannot prematurely interrupt oil flow to the propeller dome. Once the low pitch stop is reached, propeller blade angle is controlled by the beta valve. Since the beta valve controls oil flow to the dome for beta and reverse, it is mechanically impossible to bring the power levers into reverse unless the propellers are on the low pitch stops.

Attempting to pull the power levers into reverse with the propellers feathered causes damage to the power levers’ reversing linkage. Attempting to pull the power levers into reverse with the engines shut down will damage the reversing system.

Figure 7-36 Propeller Positioning Diagram (Sheet 2 of 2)

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OVERSPEED GOVERNOR Since the PT6’s propeller is driven by a free turbine (independent of the engine’s compressor), overspeed can rapidly occur if the primary governor fails. The overspeed governor provides protection against excessive propeller speed in the event of primary governor malfunction.

The hydraulic overspeed governor (Figure 7-37) is located on the left side of the propeller reduction gearbox. The operating point of the overspeed governor is 4% (6% for UE) greater than the primary governor’s maximum speed.

Since maximum propeller speed selected on the primary governor is 1,700 rpm, then overspeed

PILOT TRAINING MANUAL

governor set speed is 1,768 rpm (1,802 rpm for UE). If a propeller’s speed reached 1,768 rpm (1,802 rpm for UE), the overspeed governor would begin increasing blade angle to a higher pitch to prevent the rpm from continuing its rise.

From a pilot’s point of view, a propeller tachometer stabilized at approximately 1,768 rpm (1,802 rpm for UE) would indicate failure of the primary governor and proper operation of the overspeed governor. For pretakeoff check purposes, the set point of the overspeed governor is rescheduled using the prop governor test switch on the pilot’s left subpanel. During testing, propeller speed should not exceed approximately 1,520 to 1,610 rpm (1,535 to 1,595 rpm for UE).

Figure 7-37 Overspeed Governor Diagram

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FUEL TOPPING GOVERNOR The primary propeller governor contains a fuel topping governor which prevents power turbine overspeed if a propeller malfunctions. An overspeed could occur, for example, if a propeller blade were to stick in a fixed position during normal primary governor operation. In addition, during reverse thrust operation, the fuel topping governor is set below the speed selected by the primary governor to permit indirect control of propeller speed by the FCU (fuel control unit) servo system.

The speed at which fuel topping governor operation occurs is determined by the speed selected with the propeller levers, and by the position of the reset lever. In the flight range the reset lever is set to regulate power turbine (N 2 ) speed at approximately 6% higher than the propeller lever setting. In the ground range the reset lever is set to 4% lower than the propeller lever position.

If propeller speed (NP) exceeds levels sensed by the fuel topping governor, fuel flow to the N1 section will be reduced, and engine power will decrease. When this occurs, propeller rpm (NP) normally remains constant, but it may decrease if propeller blades are frozen in a fixed position.

POWER LEVERS The power levers (Figure 7-38) are located on the power lever quadrant (first two levers on the left side) on the center pedestal. They are mechanically interconnected through a cam box to the fuel control unit, reverse lever, beta valve and follow-up mechanism, and the propeller (NP ) governor. The power lever quadrant permits movement of the power lever in the forward thrust (alpha) range from idle to maximum thrust, and in the beta/reverse range from idle to maximum reverse. Detents in the power lever quadrant at the IDLE and GROUND FINE positions prevent inadvertent movement of the lever into the beta/reverse range. To select beta or reverse, the pilot must lift the power levers up and over the detents.

Figure 7-38 Power Levers

In the forward thrust (alpha) range the power levers establish gas generator rpm (N G ) by selecting a gas generator governor speed which results in a fuel flow that will maintain the selected N1 rpm.

In the beta range, the power levers control the beta valve to reduce propeller blade angle, thus reducing residual propeller thrust. N1 rpm is not affected in the beta range.

In the reverse range, the power lever: (1) selects a blade angle proportionate to the aft travel of the lever, (2) selects a fuel flow that will sustain the selected reverse power, and (3) resets the fuel topping governor (NP) from its normal 106% to a range between 93 and 97%. Therefore, rpm in reverse is a function of the primary propeller governor acting through the FCU to limit fuel flow and control propeller rpm in relation to power lever position.

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PROPELLER CONTROL LEVERS Propeller rpm, within the primary governor range of 1,400 to 1,700 rpm (1,200 to 1,700 rpm for UE), is set by the position of the propeller control levers (Figure 7-39). These levers, one for each engine, are located between the power levers and the fuel cutoff levers on the center pedestal quadrant. The full-forward position sets the primary governor at 1,700 rpm. In the full-aft position, forward of the feathering detent, the primary governor is set at 1,400 rpm (1,200 rpm for UE).

Intermediate propeller rpm positions can be set by moving the propeller levers to select desired rpm as indicated on the propeller tachometer. The tachometers are read directly in revolutions per minute. In the Series UE the propeller rpm gages are 28-VDC-powered gages.

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A detent at the low rpm position prevents inadvertent movement of the propeller lever into feather. The feather position is indicated by red and white stripes at the bottom of the propeller lever slots in the power quadrant. The taxi position for the propeller levers is indicated by a white line just prior to the feather detente on Series UE.

PROPELLER FEATHERING The propeller can be manually feathered by moving the propeller lever full aft, past the detent, into feather. This action locks the governor’s pilot valve in the full-up position, opens the feather valve, and all oil quickly drains from the propeller pitch mechanism. As oil is dumped from propeller servo chambers, the counterweights and springs drive propeller blades to the feathered position. Since the propeller shaft and the N1 shaft are not connected, the propeller can be feathered with the engine running; however, to avoid excessive torque loads on the propeller gearbox, the engine should be at idle power when the propellers are manually feathered. If an engine fails with the autofeather system inoperative, the propeller will maintain onspeed rpm unless it is feathered manually.

BEFORE TAXI AND BEFORE TAKEOFF CHECKS The Before Takeoff (Runup) checklist in the POH/AFM “Normal Procedures” section contains preflight checks for the propeller system.

To minimize propeller blade erosion and pitting of airplane surfaces, the following system checks should be accomplished on a clean, hard surface which is free of sand and gravel.

Figure 7-39 Propeller Control Levers

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AUTOFEATHER SYSTEM (SERIES UA, UB, UC) The autofeather system (Figure 7-40) provides a means of dumping oil from the propeller dome automatically. Thus, the feathering spring and counterweights start the blade feathering action immediately in the event of an engine failure.

Although the system is armed by a switch on the subpanel, placarded “AUTOFEATHER” with ARM, OFF, and TEST positions, the arming phase is completed only when both power levers are advanced above 85 to 90% N1 and engine torque is above 525 foot-pounds. When armed, both right and left indicator lights on the caution/advisory annunciator panel will be illuminated (Figure 7-41). The green annunciator

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lights are placarded “L AUTOFEATHER” and “R AUTOFEATHER.” The autofeather system will be inoperative as long as either power lever is retarded below 85 to 90% N1.

The autofeather system should be armed during takeoff, climb, approach, and landing, and should be turned off in cruise. When the system is armed, if torquemeter oil pressure on either engine drops below 250 foot-pounds, oil is dumped from the dome, the feathering spring starts the blades toward feather, and the autofeather system of the other engine is disarmed. When the autofeather portion of the operative engine is disarmed, the annunciator indicator light for that engine will be extinguished (Figure 7-42). Autofeather is required to be armed and operable for flight.

Figure 7-40 Autofeather Test Diagram—Series UA, UB, UC

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Figure 7-41 Autofeather System Diagram (Armed)—Series UA, UB, UC

Figure 7-42 Autofeather System Diagram (Armed, Left Engine Failure)—Series UA, UB, UC

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AUTOFEATHER SYSTEM (SERIES UE)

PROPELLER SYNCHROPHASER SYSTEMS

The autofeather system (Figure 7-43) provides a means of dumping oil from the propeller dome automatically. Thus, the feathering spring and counterweights start the blade feathering action immediately in the event of an engine failure.

Although the system is armed by a switch on the subpanel, placarded “AUTOFEATHER” with ARM, OFF, and TEST positions, the arming phase is completed only when both power levers are advanced above 90% N1 and engine torque is above 1,000 foot-pounds. When armed, the green L and R AUTOFEATHER and L and R AFX annunciators are illuminated. The AFX DISA B L E a n n u n c i a t o r i s ex t i n g u i s h e d . T h e autofeather system will be inoperative on either engine as long as its power lever is retarded below 90% N1.

The propeller synchrophaser automatically matches the speed of the two propellers to synchronize rpm. In addition, propeller blades on one engine are maintained at a predetermined relative position with the blades of the other so that only one propeller blade at a time will rotate past the fuselage. The synchrophaser system reduces propeller beat and cabin noise.

The autofeather system is required to be armed and operable for flight during takeoff, climb, approach, and landing, and should be turned off in cruise. When the system is armed, if torquemeter oil pressure on either engine drops below 350 foot-pounds, oil is dumped from the dome, the feathering spring starts the blades toward feather, and the autofeather system of the other engine is disarmed. When the autofeather portion of the operative engine is disarmed, the AUTOFEATHER and AFX annunciator lights for that engine will be extinguished, and the AFX DISABLE annunciator will be illuminated.

After the autofeather system has been tested, the propeller manual feathering system should be checked. Be sure to verify that engine power is at idle, then bring both propeller levers into the feather detent. The propellers may be allowed to completely feather with the compressor operating at low idle without engine damage; however, ground operations while the propellers are feathered should be kept to a minimum. Extensive ground operations in feather may overheat the fuselage and can damage nose-mounted avionics when hot exhaust gases are not being blown aft by the propeller’s air blast.

Synchrophaser Operation The synchrophaser system (Figure 7-44) is an electronic system certificated for use during all flight operations, including takeoff and landing.

It is not a designated master-slave system. It always matches the rpm of the slower propeller to the rpm of the faster propeller, and constantly keeps the propellers in phase to reduce cabin noise to a minimum.

The synchrophaser has a limited range of authori t y. T h e m a x i m u m p o s s i b l e i n c r e a s e i s approximately 25 rpm. In no case will the rpm fall below that selected by the propeller control lever. Normal governor operation is unchanged, but the synchrophaser continuously monitors propeller rpm and resets either governor as required.

Propeller rpm and relative blade position are sensed by a magnetic pick-up mounted adjacent to each propeller spinner bulkhead. The magnetic pick-up transmits electrical pulses, once per revolution, to a control box installed forward of the pedestal.

The control box converts any pulse rate differences between the propellers into correction commands, and transmits the commands to coils mounted close to the flyweights of each primary governor. As coil voltages vary, the governor speed settings are biased until the propeller rpm’s match exactly. A toggle switch, installed adjacent to the synchroscope, turns the system on. When the synchrophaser is off, propeller governors operate at the manual speed settings selected by the pilot. To operate the synchrophaser system,

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Figure 7-43 Autofeather System Diagram—Series UE

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synchronize the propellers manually to establish a maximum of 25 rpm difference between them, then turn the synchrophaser on. The system can be on at all times unless a malfunction is indicated.

Propeller rpm can be reset when the synchrophaser system is turned on. To do so, simply adjust both propeller controls simultaneously. If the synchrophaser is on, but is unable to adjust the propeller rpm’s to match, the system has reached the end of its operating range. Increasing the setting of the slow propeller, or reducing the setting of the fast propeller, will bring the speeds within synchrophaser range. If preferred, turn the synchrophaser switch off, resynchronize manually, and turn the synchrophaser on.

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Propeller Synchroscope A propeller synchroscope, located to the left of the oil pressure/temperature gages, indicates the status of propeller synchronization. The face of the synchroscope has a black-and-white cross pattern which can spin either left or right and turns toward the faster propeller. Therefore, if the right propeller rpm is greater than the left, the face turns clockwise (to the right). When the left propeller rpm is greater than the right, the face turns counterclockwise (to the left). No rotation of the synchroscope indicates that both propellers are synchronized.

Figure 7-44 Propeller Synchrophaser

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CHAPTER 8 FIRE PROTECTION CONTENTS Page INTRODUCTION .................................................................................................................. 8-1 FIRE DETECTION SYSTEM (SERIES UA, UB, UC, UE).................................................. 8-1 Fire Detection and Extinguisher Test (Series UA, UB) .................................................. 8-3 Fire Detection Test (Series UC) ...................................................................................... 8-3 Fire Detection Test (Series UE) ...................................................................................... 8-5 FIRE EXTINGUISHING SYSTEM (SERIES UA, UB, UC, UE)......................................... 8-5 Fire Extinguisher Test (Series UC) ................................................................................. 8-6 Fire Extinguisher Test (Series UE) ................................................................................. 8-6

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ILLUSTRATIONS Figure

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Engine Fire Detection System Schematic................................................................ 8-2

8-2

Fire Extinguisher Control Switches and Firewall Fuel Shutoff Valve “T” Handles ............................................................... 8-3

8-3

Engine Fire Extinguishing System Schematic......................................................... 8-4

8-4

Fire Extinguisher Control Switches......................................................................... 8-5

8-5

Engine Fire Extinguisher Test Switches.................................................................. 8-6

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CHAPTER 8 FIRE PROTECTION

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FIRE WARN

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FIRE PULL

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INTRODUCTION

Fire detection and protection systems provide fire warning and fire extinguishing capability of a fire in either engine.

FIRE DETECTION SYSTEM (SERIES UA, UB, UC, UE) The engine fire detection system provides immediate warning in the event of fire in either engine compartment (Figure 8-1). The system incorporates a temperature sensing cable looped continuously around each engine. The cables from left and right nacelles are interconnected to a single control amplifier, mounted on the forward pressure bulkhead behind the left subpanel.

Should an engine fire occur, the amplifier activates a red warning light in the firewall fuel shutoff valve handle placarded FIRE PULL. Two fire extinguisher control switches placarded L or R ENGINE FIRE PUSH TO EXT are in the

glareshield above the firewall valve “T� handles (Figure 8-2).

The sensing cable forms a continuous loop over and around areas in which critical fire situations could develop. The cable functions as a temperature-sensitive resistor (thermistor) which senses temperature increase as a drop in resistance.

When resistance drops below a preset value, a signal is sent to a control amplifier which then illuminates the fire warning annunciators. Since the warning system functions as a temperature averaging device, and variances in temperatures cause differences in resistance at different points

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along the cable, base resistance must be determined. This means that the temperatures along the entire length of the cable are compared and averaged to determine the base resistance at which a fire will be sensed. In addition, the resistance value that triggers fire warning indications is different on any given day relative to outside air temperature; therefore, base resistance automatically adjusts to compensate for differences in outside ambient air temperatures.

In the unlikely event of an engine fire, cable temperature rises, thereby decreasing its resistance.

When the resistance drops below its absolute set

PILOT TRAINING MANUAL

point value of approximately 100 ohms, it will trigger the amplifier, illuminating the “T� handle lights.

NOTE (Series UE only) In the UE Series, the loop continuity is monitored and if a break is detected, the Fire Loop annunciator and Master Caution Flasher will illuminate, warning the pilot that the detector circuit may not be functioning properly in the event of a fire.

Figure 8-1 Engine Fire Detection System Schematic

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Length of the continuous loop is a critical variable in the system. In the event of a single cable break, the system remains operational. If the total length of the loop is changed, the resistance at which fire is sensed will also change. If less than the full length of the cable is heated to a specific temperature, the resistance value will be greater than if the full length is heated to the same temperature. In such a situation the resistance value would be too high to activate the fire warning.

The control amplifier is set to an alarm level high enough to prevent false fire warnings. The single amplifier, which interconnects left and right fire warning systems, discriminates for short circuits in control circuitry. Ample margin exists between the fire alarm trip setting and the short circuit discriminator.

A time delay in the amplifier prevents false fire warnings because only a steady resistance signal will be sensed as an actual alarm. Intermittent signals due to shorts in control circuitry or cable connectors will not trip the fire warning annuncia t o r. T h e FAU LT t e s t f u n c t i o n h a s b e e n incorporated to test the control amplifier for open circuit conditions.

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FIRE DETECTION TEST (SERIES UC) The airplane is equipped with two toggle test switches placarded ENG FIRE TEST-DETECT, one for the LEFT system and one for the RIGHT system. The switches are three-position switches spring-loaded to the center. The switch positions are placarded LOOP-OFF-AMP. When either toggle switch is placed in the LOOP position, the integrity of the appropriate firezone cable is tested. A good test is indicated by the red lights in the appropriate FIRE PULL “T” handle being illuminated. When either toggle switch is placed in the AMP position, the integrity of the circuitry within the control amplifier is tested. A good test is indicated by the red lights in the appropriate FIRE PULL “T” handle being illuminated as in the loop test.

FIRE DETECTION AND EXTINGUISHER TEST (SERIES UA, UB) The UA/UB Series airliners were equipped with a fire detection and extinguisher test system combined as a single knob. This system has been replaced on many airliners by a retrofit kit of the same configuration as the UC-1 and after series.

For those airliners with the original system, refer to applicable POM/AFM.

Figure 8-2 Fire Extinguisher Control Switches and Firewall Fuel Shutoff Valve “T” Handles

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Figure 8-3 Engine Fire Extinguishing System Schematic

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FIRE DETECTION TEST (SERIES UE) The airplane is equipped with two toggle test switches placarded ENGINE FIRE TESTDETECT, one for the LEFT system and one for the RIGHT system. The switches have three positions and are spring-loaded to the center position. When either toggle switch is placed in the LOOP position, the integrity of the appropriate fire zone cable is tested. A good test is indicated by illumination of the appropriate yellow FIRE LOOP annunciator. When either toggle switch is placed in the AMP position, the integrity of the circuitry within the control amplifier is tested. A good test is indicated by the red light in the appropriate FIRE PULL “T” handle being illuminated.

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FIRE EXTINGUISHING SYSTEM (SERIES UA, UB, UC, UE) Figure 8-3 shows the fire extinguishing system. A fire extinguisher supply cylinder is mounted aft of the main spar in each wheelwell. Each cylind e r i s c h a rg e d w i t h 2 . 1 0 p o u n d s o f bromotrifluoromethane (Halon) pressurized to 360 psi at 700 F. Each cylinder supply line branches into two spray tubes which diffuse the extinguishing agent into the engine nacelle when activated. One nozzle discharges into the engine exhaust area, and the other discharges into the engine accessory area. Once the system is activated, the entire supply of extinguishing agent is discharged.

The fire extinguisher control switches are located on the glareshield on either side of the warning annunciator panel (Figure 8-4). Their power is derived from the hot battery bus through microswitches mounted on the firewall fuel shutoff valve. The push-to-activate switches incorporate three indicator lights.

The red lens, placarded L or R ENG FIRE PUSH TO EXT, indicates the fire “T” handle has been pulled closed, thus arming the fire extinguisher pushbutton circuitry. The red light in the extinguisher pushbutton will illuminate only if a fire exists and the “T” handle is pulled, or if the rotary test switch is in FIRE DET TEST with the “T” handle pulled. A green lens, placarded OK, is provided for test functions only. The amber lens, placarded D, monitors the status of the cylinder charge and the condition of the pyrotechnic cartridge, which must be intact before the bottle can be discharged. As long as the cartridge is intact and the cylinder has not been discharged, the amber light will remain extinguished. When the cartridge has been fired, the “D” light will remain illuminated until the cartridge is replaced.

Figure 8-4 Fire Extinguisher Control Switches

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FIRE EXTINGUISHER TEST (SERIES UC)

FIRE EXTINGUISHER TEST (SERIES UE)

The airplanes are equipped with two toggle test switches placarded ENG FIRE TEST-EXT TEST, one for the LEFT system and one for the RIGHT system installed in the copilot’s inboard subpanel. These switches test the circuitry of the fire extinguisher pyrotechnic cartridges. The switches are moved to the EXT TEST position while verifying the illumination of the appropriate yellow D light and the appropriate green OK light in each fire extinguisher activation switch on the glareshield. The toggle switches are spring-loaded and will return automatically to the center OFF position.

The airplanes are equipped with two toggle test switches placarded ENG FIRE TEST-EXT TEST A and TEST B, one for the LEFT system and one for the RIGHT system installed in the copilot’s inboard subpanel. These switches are for the purpose of testing the dual circuitry of the fire extinguisher pyrotechnic cartridges. The switches are moved to the EXT TEST A position while verifying the illumination of the appropriate yellow D light and the appropriate green OK light in each fire extinguisher activation switch on the glareshield. The switches are then moved to the EXT TEST B position while verifying the illumination of the appropriate green OK light in each fire extinguisher activation switch on the glareshield. The toggle switches are springloaded and will return automatically to the center OFF position.

A cylinder gage, calibrated in psi, provides a means of determining the charge level of each supply cylinder (Series UC and UE). These gages must be checked at the cylinder in each wheelwell during preflight.

Figure 8-5 Engine Fire Extinguisher Test Switches

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CHAPTER 9 PNEUMATICS CONTENTS Page INTRODUCTION .................................................................................................................. 9-1 DESCRIPTION....................................................................................................................... 9-1 ENGINE BLEED AIR PNEUMATIC SYSTEM ................................................................... 9-3 Pneumatic Air Source...................................................................................................... 9-3 Vacuum Air Source ......................................................................................................... 9-3 Bleed Air Control ............................................................................................................ 9-4 ENGINE BLEED AIR WARNING SYSTEM ....................................................................... 9-4 Brake Deice Overheat Warning ...................................................................................... 9-4 HYDRAULIC FILL CAN PRESSURE ................................................................................. 9-6

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ILLUSTRATIONS Figure

Title

Page

9-1

Pneumatic System Diagram..................................................................................... 9-2

9-2

Pneumatic Pressure Gage ........................................................................................ 9-3

9-3

Gyro Suction Gage .................................................................................................. 9-4

9-4

Bleed Air Valve Controls ........................................................................................ 9-4

9-5

Bleed Air Warning System Diagram ....................................................................... 9-5

9-6

Bleed Air Warning Annunciators ............................................................................ 9-5

9-7

Hydraulic Fill Can ................................................................................................... 9-6

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CHAPTER 9 PNEUMATICS

BLEED AIR CO

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R AIR

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LV VA E

INTRODUCTION Many small but important tasks are accomplished by the aircraft pneumatic and vacuum systems. This section identifies these systems, their controls, and proper utilization.

DESCRIPTION This chapter describes the sources of pneumatic and vacuum air.

Acceptable gage indications and normal, abnormal, and emergency procedures are discussed.

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Figure 9-1 Pneumatic System Diagram

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ENGINE BLEED AIR PNEUMATIC SYSTEM The pneumatic system in the Beechcraft 1900 Airliner provides support for several aircraft system operations. These operations include surface deice, brake deice, brake deice overheat warning, and bleed air warning. Pneumatic pressure is also used to create vacuum for air-driven gyros, pressurization control and deflation of the deice boots.

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Bleed air is extracted after the final stage of the engine compressor at a maximum temperature of approximately 800° F. As bleed air enters the pneumatic manifold, it is cooled to approximately 70° F above ambient temperature due to heat transfer in the pneumatic plumbing.

Ordinarily, the pressure regulator valve, which is mounted under the center aisle floor forward of the main spar, will provide 18 ± 1 psi with the engine running at 70% to 80% N1. System pressure is monitored on the pneumatic pressure gage (Figure 9-2).

High pressure bleed air from each engine compressor is routed through a pneumatic/instrument air valve and regulated at 18 psi to supply pressure for the pneumatic system and vacuum source (Figure 9-1). Vacuum for the flight instruments and for deice boot deflation is supplied by routing 18 psi pneumatic air through a bleed air ejector (venturi). One engine can supply sufficient bleed air for all pneumatic and vacuum systems.

A suction gage on the copilot’s subpanel indicates instrument vacuum pressure in inches of mercury. To the right of the suction gage is a pneumatic pressure gage, calibrated in pounds per square inch.

Figure 9-2 Pneumatic Pressure Gage

PNEUMATIC AIR SOURCE Bleed air at 90 to 150 psi pressure is obtained from both engines, and flows through the pneumatic/instrument air valve and pneumatic lines to a common tee located in the fuselage. Check valves installed on each side of the tee prevent loss of bleed air during single engine operation.

Downstream from the check valves, bleed air passes through an 18 psi regulator. The regulated bleed air then flows through a manifold that supplies pneumatic pressure for the surface deicers, the bleed air failure and brake deice overheat warning systems, the flight hourmeter, and the hydraulic powerpack reservoir. Regulated bleed air also provides flow and pressure for the vacuum ejector. Should the pneumatic pressure regulator fail, an overpressure relief valve will open at 21 psi.

VACUUM AIR SOURCE Vacuum is obtained from the bleed air vacuum ejector. Vacuum is supplied at a regulated 4.3 to 5.9 inches Hg through the vacuum regulator valve. The vacuum regulator is protected by a foam filter, and is mounted in the nose compartment on the right side of the pressure bulkhead.

The instrument vacuum line is routed through a suction relief valve designed to admit into the system the amount of air required to maintain sufficient vacuum for proper operation of the instruments. A gyro suction gage (Figure 9-3), which is calibrated in inches of mercury, is located on the copilot’s right subpanel. With one engine running at 70% to 80% N1, the vacuum gage should indicate approximately 5.9 (+0/–.2) inches Hg.

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ENGINE BLEED AIR WARNING SYSTEM The engine bleed air warning system provides visual indication of a rupture in the bleed air lines. If a leak is detected by the warning system, the affected engine’s bleed air valves can be immediately shut down before heat of the escaping bleed air damages the skin and structure adjacent to the break in the line (Figure 9-5).

Figure 9-3 Gyro Suction Gage

BLEED AIR CONTROL Bleed air entering the cabin is controlled by the two BLEED AIR VALVE switches (Figure 9-4) marked OPEN, ENVIR OFF, and INST & ENVIR OFF. When the switches are in the OPEN position, the environmental valves and the pneumatic/instrument air valves are open. When the switches are in the ENVIR OFF position, the environmental valves close and the pneumatic/instrument air valves remain open. In the INST & ENVIR OFF position, all bleed air valves are closed.

Bleed air lines are routed from the engines to the cabin, in close proximity to the ethylene vinyl acetate (EVA) tubing of the bleed air warning system. The tubing is pressurized by air which is tapped off the pneumatic manifold. Two pressure switches are mounted under the center aisle floorboards at the ends of the EVA tubing.

Excessive heat caused by a ruptured bleed air line melts the EVA tubing. When the tubing melts the pressure drops below that required to keep the pressure switch actuated and the switch closes. A circuit is then completed, illuminating the appropriate BL AIR FAIL light in the warning annunciator panel (Figure 9-6).

When the indication of bleed air failure becomes evident, all bleed air for that side must be turned off by placing the respective bleed air valve switch in the INSTR & ENVIR OFF position.

With the switch in this position, all environmental and pneumatic bleed air shutoff valves close, stopping bleed air flow at the engine firewall; however, the BL AIR FAIL light will remain illuminated since the EVA tubing is unable to hold the 18 psi required to deactivate the pressure switch. Before the annunciator can be extinguished, the leak in the EVA tubing must be repaired.

BRAKE DEICE OVERHEAT WARNING Figure 9-4 Bleed Air Valve Controls

9-4

The brake deice overheat warning system is similar to the bleed air failure warning system. The EVA tubing and pressure switches are routed near the brake deice plumbing inside the wheelwell. Normal operation of the brake deice

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Figure 9-5 Bleed Air Warning System Diagram

Figure 9-6 Bleed Air Warning Annunciators

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system will not activate the brake deice overheat warning system.

To further prevent overheating in the wheelwell, a timer automatically deactivates brake deice 10 to 12 minutes after gear retraction. However, if the timer fails, allowing brake deice to operate for extended periods, or if a bleed air leak occurs in the brake deice system, the tubing will melt.

This activates the pressure switch, illuminating t h e a p p r o p r i a t e B R K D I OV H T c a u t i o n annunciator.

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HYDRAULIC FILL CAN PRESSURE Pneumatic air pressure is routed to the landing gear hydraulic powerpack fill can (Figure 9-7). It provides positive pressure to the reservoir and pressurizes the hydraulic fluid to provide positive feed to the hydraulic pump. When the engines are shut down, the pneumatic pressure in the hydraulic fill can be bled off by a pushbutton relief valve. Before checking fluid level in the hydraulic reservoir, residual pressure must be bled off with the relief valve to prevent loss of fluid when the cap is opened.

Figure 9-7 Hydraulic Fill Can

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CHAPTER 10 ICE AND RAIN PROTECTION CONTENTS Page INTRODUCTION ................................................................................................................ 10-1 DESCRIPTION..................................................................................................................... 10-1 ICE PROTECTION SYSTEMS ........................................................................................... 10-3 Description and Operation............................................................................................. 10-3 Precautions during Icing Conditions ............................................................................. 10-3 Surface Deice Systems .................................................................................................. 10-6 Wing Ice Lights ............................................................................................................. 10-9 Stall Warning Heat ........................................................................................................ 10-9 Engine Inertial Separators ........................................................................................... 10-11 Ice Vane Controls ........................................................................................................ 10-11 Engine Air Inlet Lip Heat ............................................................................................ 10-12 Engine Autoignition System ....................................................................................... 10-13 Fuel System Anti-ice ................................................................................................... 10-14 Propeller Electric Deice System.................................................................................. 10-17 Windshield Anti-ice .................................................................................................... 10-18 Windshield Wipers ...................................................................................................... 10-22 Brake Deice System .................................................................................................... 10-23 Pitot-Static Mast Heat ................................................................................................. 10-24 Alternate Static-Air System......................................................................................... 10-25 Alternate Static Heat ................................................................................................... 10-25

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ILLUSTRATIONS Figure

Title

Page

10-1

Anti-icing and Deicing Components ..................................................................... 10-2

10-2

Ice and Rain Protection Required Equipment ....................................................... 10-2

10-3

Ice and Rain Protection Controls ........................................................................... 10-4

10-4

Surface Deice Boot Installation ............................................................................. 10-6

10-5

Surface Deice System Diagram (Tailets—Series UE Only) ................................. 10-8

10-6

Surface Deice Controls .......................................................................................... 10-8

10-7

Wing Ice Lights ..................................................................................................... 10-9

10-8

Stall Warning Vane and Heat Control ................................................................... 10-9

10-9

Inertial Separators in RETRACT Position .......................................................... 10-10

10-10

Inertial Separators in EXTEND Position............................................................. 10-10

10-11

Ice Vane Controls ................................................................................................ 10-11

10-12

Caution/Advisory Annunciators .......................................................................... 10-12

10-13

Engine Air Inlet Lip Heat .................................................................................... 10-13

10-14

Engine Autoignition System................................................................................ 10-13

10-15

Fuel System Anti-ice ........................................................................................... 10-14

10-16

Oil-to-Fuel Heat Exchanger................................................................................. 10-15

10-17

Propeller Electric Deice System .......................................................................... 10-16

10-18

Propeller Deice Boots .......................................................................................... 10-17

10-19

PROP AMPS Indicator ........................................................................................ 10-18

10-20

Windshield Anti-ice Switches ............................................................................. 10-18

10-21

Windshield Anti-ice Diagram.............................................................................. 10-19

10-22

Windshield Anti-ice Diagram—Normal Heat ..................................................... 10-20

10-23

Windshield Anti-ice Diagram—High Heat ......................................................... 10-21

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10-24

Windshield Wipers.............................................................................................. 10-22

10-25

Brake Deicer........................................................................................................ 10-23

10-26

Pitot Masts and Heat Controls............................................................................. 10-24

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CHAPTER 10 ICE AND RAIN PROTECTION

INTRODUCTION Flight into known icing conditions requires pilot understanding of meteorological conditions which are conducive to icing. The pilot must be familiar with all aircraft anti-ice and deice systems which prevent excessive ice formation from interfering with the safety of flight. This section identifies these systems, their controls, and best usage.

DESCRIPTION This chapter presents a description and discussion of the airplane ice and rain protection systems. All anti-ice and deice systems are described showing location, controls, and procedures for use. The purpose of this section is to acquaint the pilot with the systems for flight into

icing or heavy rain conditions. Systems controls and use are described, and procedures to be followed in the event of system malfunctions are discussed. Information concerning preflight deicing and defrosting are included.

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Figure 10-1 Anti-icing and Deicing Components

Figure 10-2 Ice and Rain Protection Required Equipment

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ICE PROTECTION SYSTEMS The Beech 1900 Airliner has been approved by the Federal Aviation Administration for flight in known icing when the required equipment is installed and operational (Figure 10-1). The POH/AFM Limitations section contains the Required Equipment for Various Conditions of Flight List (Figure 10-2). Ice protection controls are primarily grouped together on the pilot’s subpanel. The windshield wiper control is located on the overhead panel (Figure 10-3).

DESCRIPTION AND OPERATION The airplane is equipped with a variety of ice protection systems for operation in inclement weather conditions. Only one, the surface deice, is a deicing system designed to be used AFTER ice has accumulated. All other ice protection systems are to be used as anti-icing systems to PREVENT the formation of ice on aircraft surfaces. The following is a list of ice protection systems provided for the 1900 Airliner: ●

Inertial separators (ice vanes)

●

Engine intake lip heat

●

Pitot heat

●

Alternate static

●

Propeller deice

●

Windshield anti-ice

●

Stall warning heat

●

Fuel vent heat

●

Brake deice

●

Surface deice (leading-edge boots)

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Ice protection for the engine is provided by an inertial separation system which is driven by an electrical actuator. Should the main electrical actuator motor fail, a standby actuator motor can be used. The leading-edge lip of the engine air inlet is continuously heated by engine exhaust air. The propellers are protected from icing by electrothermal boots on each blade.

Electrical heating elements embedded in the windshield provide adequate protection against the formation of windshield ice. Warm air from the cabin heating system is used for defogging.

Heavy-duty windshield wipers for both the pilot and copilot provide increased visibility for approaches and taxi operations in rain.

A heating element in the pitot-static mast prevents the pitot opening from becoming clogged with ice. The heating element is powered by the airplane electrical system through a 15-ampere circuit breaker switch.

Brake deice is a standard installation which prevents ice and slush buildup on the main wheel brakes.

Pneumatic deicing boots on the wings, stabilons, and horizontal stabilizers remove ice after it has formed on the leading edges of these surfaces.

On Series UE, the tailets also have boots on the leading edges. Regulated bleed-air pressure inflates the boots, and vacuum pressure deflates the boots. The selector switch that controls the system permits automatic single-cycle or manual operation.

PRECAUTIONS DURING ICING CONDITIONS During winter, a careful preflight inspection is required before operating in cold weather or in potential icing conditions. In addition to the normal exterior inspection, special attention should be paid to areas where frost and ice may accumulate.

Even a thin layer of frost can cause great harm. It is not the thickness of the frost that matters, it is the texture. A slightly irregular surface can sub-

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Figure 10-3 Ice and Rain Protection Controls

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stantially decrease proper airflow over wings and stabilizers. Never underestimate the damaging effects of frost. All frost should be removed from leading edges of the wings, stabilons, stabilizers and propellers before the airplane is flown.

Pitot-static masts should always be covered while the airplane is parked for extended periods. Once the covers are removed, make sure both masts are free of ice or water. Flight instrument indications may be faulty if masts are clogged.

Control surfaces, hinges, windshield, pitot masts, fuel tank caps and vents should also be free of frost before flight. Deicing fluid should be used when needed.

Before taxiing, the brake deicers may be turned on to help rid the brake mechanisms of accumulated ice. If the brake deicers are used, the condition lever should be placed in the HIGH IDLE position.

Fuel drains should be tested for free flow. Water in the fuel system has a tendency to condense more readily during winter months, and if left unchecked, large amounts of moisture may accumulate in the fuel system. Moisture does not always settle at the bottom of the tank. Occasionally a thin layer of fuel gets trapped under a large mass of water, and water may not drain if only a small sample is taken.

Keep flaps retracted to avoid throwing snow or slush into flap mechanisms, and to minimize the possibility of damage to flap surfaces. When taxiing in extremely icy conditions, make sure tires are rolling, not just sliding on the icy surface.

Engine autoignition should not be used for extended taxiing or ground holding. This precaution prolongs the service life of the igniters.

Although anti-icing additives are rarely needed in the 1900 Airliner fuel system, it is important to add correct amounts of additive when necessary.

Higher concentrations of anti-icer does not ensure lower fuel freezing temperatures, and too great a concentration can damage the fuel system. Fuel system ice protection will be discussed later in this section.

Snow, slush or standing water on the runway degrades airplane performance during takeoffs and landings. For takeoff, more runway is needed to achieve necessary takeoff speed, and landing roll is longer because of reduced braking effectiveness. Brake deicers should be activated well before landing and left on to dry the brakes after takeoff when conditions are slushy or snowy.

The brakes and tires should be checked before taxiing the airplane. If an anti-ice solution is needed to free the brakes, be sure the solution does not contain oil-based lubricants. If tires are frozen to the ground, use undiluted defrosting fluid or a ground heater to melt the ice, then move the airplane as soon as the tires are free.

Heat applied to tires should not exceed 160Ëš F or 71Ëš C.

The 1900 Airliner is equipped with both deicing and anti-icing equipment. However, only the surface deice is a true deicing system. That is, surface deice is intended to eliminate ice which has already accumulated. The remaining ice protection systems are considered to be anti-ice systems, and should be used to prevent ice formation. Accumulated ice on even the bestequipped airplane will degrade its performance and alter time and fuel calculations. A minimum speed of 160 KIAS should be maintained at all times while flying in icing conditions to prevent ice formation on the underside of the wing, since these surfaces have no ice protection.

In addition to preventing unnecessary reduction gearbox wear, using propeller tie-downs is effective as an ice preventive when the airplane is parked during cold weather. When the propeller is properly secured, moisture is channeled down the blades, past the propeller hub, and off the lower blade. The propeller hubs should also be inspected for ice and snow accumulation.

Due to distortion of the wing airfoil, stalling airspeeds should be expected to increase as ice accumulates on the airplane. For the same reason, stall warning devices are not accurate and should not be relied upon. Always maintain a comfortable margin of airspeed above the normal

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stall airspeed when ice is on the airplane. It may be necessary to reduce airspeed should ice accumulate on the windshield.

Engine ice vanes should be extended any time potential icing conditions are encountered.

Because of ram-air effect, the engine icing will occur at ambient air temperatures slightly above freezing. Even small pieces of ice can damage compressor blades.

Engine anti-ice should be used: ●

Before visible moisture is encountered at OAT +5°C and below

●

At night when freedom from visible moisture is not assured and the OAT is +5° C or below

Before entering icing conditions, fuel vent heat, pitot heat, prop deice, windshield heat, and stall warning heat should all be ON.

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Since engine bleed air is used for pneumatic boot operation and for brake deice, the bleed-air warning system should be monitored during icing flight. If either BLEED AIR FAIL annunciator illuminates in flight, the bleed-air switch on the affected engine must be moved to the INST & ENVIR OFF position. Since pneumatic air is used to pressurize the bleed-air warning system, BLEED AIR FAIL lights may illuminate momentarily during simultaneous wing boot and brake deice operation at low N1 speeds. If the annunciators are immediately extinguished after increasing N1 speed, they can be disregarded.

SURFACE DEICE SYSTEMS The wings, stabilons, and horizontal stabilizer, also tailets, (UE only) are deiced in flight with a system of inflatable rubber boots attached to leading edges of these surfaces (Figure 10-4).

Figure 10-4 Surface Deice Boot Installation

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After ice has accumulated, pneumatic pressure can be cycled through a deice distributor valve to inflate the boots. After the inflation cycle, vacuum is cycled through the distributor valve to deflate the boots, and hold them tightly against the leading-edge surfaces. Ice is removed by alternately inflating and deflating the deice boots (Figure 10-5).

When the switch is held in the MANUAL position, all the boots are inflated simultaneously, remaining inflated until the switch is released.

After the switch is released, vacuum is supplied to the boots until the switch is again pressed to manual. The manual position bypasses the electronic timer, providing an alternate means of inflating the boots should the timer fail.

To ensure normal pneumatic deicing operation, a check valve is incorporated into the bleed-air line from each engine, preventing loss of bleed-air pressure through the compressor of an inoperative engine.

Electrical power is required for boot inflation in both single-cycle and in manual. If electrical power is lost, the boots cannot be inflated, but vacuum is always available for deflation since the distributor valve is deenergized to the vacuum side of the cycle. A single circuit breaker, located on the copilot’s circuit-breaker panel, supplies electrical power for the surface deice system.

A three-position switch on the pilot’s subpanel, placarded SURFACE DEICE – SINGLE – MANUAL, controls surface deice operation (Figure 10-6). The switch is spring-loaded to the center position. Momentary activation of the switch to the SINGLE position, starts the inflation cycle. During single-cycle operation, an electronic timer controls the deice distributor valve to provide sequential inflation of the boots.

First, the outboard wing boots are inflated for approximately six seconds. During the next sixsecond cycle, all other boots are inflated, including the center wing, horizontal stabilizer, and stabilon boots. When all boots have inflated and deflated, after approximately 12 seconds, the cycle is complete. The deice distributor valve then returns to the deflate position, distributing constant vacuum to all surface deice boots.

For most effective deicing, at least 1 to 1-1/2 inches of ice should be allowed to form before attempting ice removal. Very thin ice may crack and cling to the boots instead of shedding when the boots are inflated. Subsequent cyclings will then have a tendency to build up a shell of ice outside the contour of the leading edges of the boots, making ice removal efforts ineffective.

CAUTION The deicer boots must not be cycled below –40° F. Exceeding this limit can result in permanent damage to the boots.

The 1900D (Series UE) has an INBD WG DEICE, an OUTBD WG DEICE and a TAIL DEICE annunciator which will illuminate during the period the selected boots are inflated. There are no annunciators on other 1900 models.

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Figure 10-5 Surface Deice System Diagram (Tailets—Series UE Only)

Figure 10-6 Surface Deice Controls

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WING ICE LIGHTS During night flight, the wing ice lights can be used as necessary to check for wing ice accumulation. The wing ice lights should not be used for prolonged periods while the airplane is on the ground.

The wing ice lights are on the outboard side of each nacelle. A circuit-breaker control switch is located on the pilot’s right subpanel in the lighting control panel above the ice protection control switches (UA, UB, UC) and on the overhead lighting panel (Series UE) (Figure 10-7).

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transducer vane and the metal plate surrounding the vane (Figure 10-8) are provided with an antiicing system which uses electrical heating elements to prevent ice formation.

The stall warning heat system supplies electric heat to the metal plate. It is activated by a twoposition circuit-breaker switch located on the pilot’s right subpanel. The stall warning vane and plate are heated whenever the battery switch and the stall warning heat switch are selected ON. A safety switch on the left landing gear limits current flow to the vane and metal plate to prevent overheating while the airplane is on the ground. In flight, after the left strut extends, the full 24 to 28 volts is applied to the stall warning heat components.

The UE model also incorporates an amber STALL HEAT annunciator that will illuminate whenever insufficient current is being applied to the vane. The light will also be illuminated whenever the switch is in the OFF position.

Figure 10-7 Wing Ice Lights

STALL WARNING HEAT Ice buildup changes wing contour, and the disrupted airflow may prevent the stall warning system from accurately indicating an imminent stall. In addition, the stall speed increases whenever ice accumulates on any airplane.

However, the FAA requires ice protection for the stall warning system for flight into known or forecast icing conditions. The stall warning lift

Figure 10-8 Stall Warning Vane and Heat Control

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Figure 10-9 Inertial Separators in RETRACT Position

Figure 10-10 Inertial Separators in EXTEND Position

.

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ENGINE INERTIAL SEPARATORS

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ICE VANE CONTROLS

The inertial separator system is installed in each engine to prevent foreign objects such as ice, dust, and gravel from entering the engine inlet plenum. During normal flight operations when icing conditions are not present, the ice vanes and bypass doors are retracted (Figure 10-9). At temperatures above +5° C the ice vane and door should be in the RETRACT position since ice formation is unlikely at these temperatures.

The ice vanes should be extended for all ground operations and during flight operations when visible moisture is present and the temperature is +5° C or below. Since air temperature decreases as it moves through the inlet toward the engine air intake screen, moisture can enter the engine as water or water vapor and freeze when it reaches the engine intake screens. As the ice continues to build on the intake screens, it may break off into small pieces which enter the compressor section and may cause severe damage to compressor blades. Therefore, the ice vanes should be extended any time the outside air temperature reaches +5˚ C and moisture is present or suspected.

When the ice vanes are in the EXTEND position (Figure 10-10), a sudden turn is introduced into the air inlet creating a venturi effect. At the same time the bypass door in the lower cowling at the aft end of the air duct opens. As the mixture of air and ice particles or water droplets enters the inlet, it is accelerated by the Venturi effect. Due to their greater mass, and therefore greater momentum, the heavier ice particles accelerate past the screen area and are discharged overboard through the bypass door. The airstream, however, makes the sudden turn free of ice particles and enters the engine through the inlet screen.

The ice vane and bypass doors are extended or retracted simultaneously by electric actuators.

The actuators are controlled by switches labeled ENGINE ANTI-ICE which are located on the pilot’s left subpanel (Figure 10-11).

Figure 10-11 Ice Vane Controls

When the ice vanes are extended, two green advisory annunciators will be illuminated. Since inlet airflow is restricted by the vanes, torque will decrease proportionate to power setting, and ITT will be increased very slightly. When the ice vanes are retracted, the annunciators will be extinguished, torque will increase and ITT will decrease.

The ice vane control switches are placarded ENGINE ANTI-ICE LEFT – RIGHT – ON – OFF. Dual actuator motors provide redundancy for operation of the electric actuators. Additional selector switches for left and right engine ice vanes allows the system to be driven by either the main or the standby actuator motor. The switch is placarded ACTUATORS – STANDBY – MAIN. The ice vanes cannot be extended to intermediate positions. They are either extended or retracted.

Ice vane position and status of the inertial separator system is indicated by illumination of L and R ENG ANTI-ICE (green) or L and R ENG ICE FAIL annunciators (Figure 10-12).

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When the ice vanes have been extended fully, the L and R ENG ANTI-ICE annunciators will be illuminated. If the vanes have not reached the fully extended position within 30 seconds after the system has been turned on, the L or R ENG ICE FAIL annunciator will be illuminated. If the ice vanes have not reached the fully retracted position within 30 seconds after selecting ice vanes off, the L or R ENG ICE FAIL annunciator will be illuminated. If the annunciator is illuminated, the standby actuator should be selected.

The ice fail annunciator circuit compares ice vane control switch position to the actuator position microswitches. The L and/or R ENG ICE FAIL annunciators will be illuminated if the control switch position and microswitch position do not agree. In addition, if the power source for the actuator system selected (either MAIN or STBY)

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is inoperative, the engine ice fail annunciator will be illuminated immediately. Immediate illumination of the ICE FAIL annunciator indicates loss of electrical power, whereas delayed illumination indicates an inoperative actuator.

ENGINE AIR INLET LIP HEAT The lip around each air inlet is continuously heated by hot exhaust gases (Figure 10-13). A scoop inside the engine’s left side exhaust stack deflects the hot exhaust gases downward into the hollow lip tube that encircles the engine air inlet.

The gases are expelled with the engine exhaust through a line in the right side exhaust stack.

Heat flows through the inlet whenever the engine is running.

Figure 10-12 Caution/Advisory Annunciators

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w h e n fl y i n g i n t o t u r bu l e n c e o r h e av y precipitation. If the system is armed, the igniters will be automatically activated when torque falls below 550 foot-pounds (750 foot-pounds on Series UE).

Figure 10-13 Engine Air Inlet Lip Heat

ENGINE AUTOIGNITION SYSTEM The engine autoignition system provides automatic continuous ignition to prevent engine p ow e r l o s s d u e t o c o m bu s t i o n f a i l u r e .

Autoignition is armed during icing flight, or

Control switches for the autoignition system are located on the pilot’s left subpanel, above the ice vane switches and to the left of the control column (Figure 10-14). The system is activated by moving the switches up to the ARM position.

Each switch must be lifted over a detent before it can be moved into, or out of, the ARM position.

This lever-lock feature prevents inadvertent operation of autoignition.

When autoignition is armed, as torque falls below 550 foot-pounds (750 foot-pounds on Series UE), the green IGNITION ON annunciator on the caution/advisory panel will be illuminated, indicating that the igniters are energized. During ground operation, the system should be turned off to prolong the life of the igniter units.

Figure 10-14 Engine Autoignition System

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FUEL SYSTEM ANTI-ICE Several anti-ice systems protect fuel flow through the fuel lines to the engine (Figure 10-15). Moisture in fuel can freeze and fuel can thicken during flight in extremely cold temperatures.

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Ice formation in the fuel vent system is prevented by electrically-heated vents in each wing. Fuel vent heat is controlled with toggle switches for left and right fuel vents. These switches, located on the ice protection panel, should be turned on whenever ice is anticipated or encountered.

Fuel temperature in the fuel control unit is maintained by an oil-to-fuel heat exchanger, mounted on the engine’s accessory section (Figure 10-16).

An engine oil line within the heat exchanger is located next to the fuel line. Heat transfer occurs through conduction between these two lines before fuel is delivered to the fuel control unit.

The heat exchanger melts ice particles, and prevents the fuel from thickening in extremely cold temperatures. The heat exchangers operate automatically whenever the engines are running.

Refer to the POH/AFM Limitations Section for fuel temperature limitations.

Figure 10-15 Fuel System Anti-ice

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Figure 10-16 Oil-to-Fuel Heat Exchanger

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Figure 10-17 Propeller Electric Deice System

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PROPELLER ELECTRIC DEICE SYSTEM The propeller electric deicer system consists of electrically heated deice boots, slip rings and brush block assemblies, a timer for automatic operation, and a dual-scale ammeter (Figure 10-17). Two switches on the pilot’s right subpanel control automatic or manual deicing operations, and two circuit breakers for auto and manual control circuit protection are located on the copilot’s circuit-breaker panel.

Although propeller deice is capable of removing ice from the propeller after it has accumulated, the system is normally used as an anti-icing system and it should be turned on before entering icing conditions. The heated boots reduce ice adhesion on propeller blades (Figure 10-18). The ice is then removed by the centrifugal effect of the propellers and the blast of the airstream.

Figure 10-18 Propeller Deice Boots

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When the AUTO switch is turned on, electrical current flows from the timer through the brush block assemblies and to the slip rings on the back of the propeller spinner, where it is distributed to the individual propeller deicing boots. Since the automatic control switch is a circuit-breaker switch, excessive current flow will cycle the switch off and remove power to the deicer timer.

When propeller deice is operated in AUTO, power to heating elements in the deicer boots is cycled in 90-second phases. The first 90-second phase heats all the boots on one propeller, and the next 90-second phase heats all the boots on the other propeller. The full cycle may begin with either propeller. The automatic timer completes one full cycle every three minutes, and propeller deice functions continuously until the AUTO switch is turned off.

A manual system is provided as a backup to the automatic system. The MANUAL control switch, located to the right of the AUTO control switch (Figure 10-6), controls the manual override relays. The manual switch is spring-loaded to the center (off) position, and must be held in place until the ice has been dislodged from the propell e r s u r fa c e . W h e n t h e s w i t c h i s h e l d i n MANUAL, electrical current bypasses the automatic timer, simultaneously heating all propeller deice boots.

Both automatic and manual operations can be monitored on the propeller deice ammeter (Figure 10-19). The right side of the dual-scale ammeter provides an indication of current flow to the right propeller and the left side of the ammeter indicates current for the left propeller. Normal indications are 26 to 32 amperes (32 to 38—Series UE) per side. During automatic deicing, as heating cycles move from one propeller to the other, ammeter deflection will move from one side of the scale to the other. In manual, both scales will show simultaneous deflections, indicating both left and right systems are energized.

The electrical system load meters will indicate approximately 0.10 increase in load in manual and 0.05 increase in automatic.

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Propeller deice must not be operated when the propellers are static to avoid damage to the brush blocks and slip rings.

Figure 10-19 PROP AMPS Indicator

WINDSHIELD ANTI-ICE The pilot’s and copilot’s windshields are each heated independently. The windshield heat control switch can be selected to a HI intensity heat level or a lower NORMAL intensity heat level.

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embedded in each windshield and a temperature controller in each windshield heat circuit. The temperature controller maintains windshield temperatures at 90 to 110° F. The control switches for windshield heat (Figure 10-20) are placarded WSHLD ANTI-ICE – NORMAL – OFF – HI – PILOT COPILOT. When NORMAL is selected, the low heat relay is energized and the entire windshield is heated (Figure 10-22). The low heat relay is activated by the temperature controller, as necessary, to maintain the preset temperature. When the switches are in the HI position, the high heat relay is energized, and a smaller portion (about 2/3) of the windshield area is heated (Figure 10-23).

Although temperature of the windshield is maintained to the same preset value in HI, the windshield is heated faster in the HI position.

Each switch must be lifted over a detent before it can be moved into high heat. The lever-lock feature prevents inadvertent selection of HI heat when moving the switches from NORMAL to the OFF (center) position.

The pilot’s and copilot’s windshield heat systems are controlled and powered differently. On UC-1 and after, the pilot’s system utilizes a remote control circuit breaker (RCCB) for control and one for power. The RCCBs are controlled by 1/2 amp circuit breakers on RH circuit breaker panel.

The windshields are composed of three layers.

The thick glass inner panel is the supporting structural layer. The middle layer is a polyvinyl sheet which contains the gold-filament fine wire heating grids. The outer layer is protective glass bonded to the first two layers. The outside of the windshield is treated with a static discharge film called a NESA (non-electrostatic application) coating. This transparent material (usually stannous oxide) has high electrical resistance which promotes a constant bleed-off of static electricity on the windshields.

Windshield heating elements (Figure 10-21) are connected through terminal blocks in the corner of the glass to the control switches in the ice protection group of the pilot’s right subpanel.

Windshield temperature is controlled automatically by a temperature-sensing element

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Figure 10-20 Windshield Anti-ice Switches

The power circuit of both HI and NORMAL heat is protected by 50 ampere current limiters located in the power distribution panel. Windshield heater control circuits are protected with a 5ampere circuit breaker on the copilot’s circuit breaker panel.

Use of windshield heat causes erratic operation of the magnetic compass. It may also be necessary to reduce airspeed in order to keep the windshield ice-free in sustained icing conditions.

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Figure 10-21 Windshield Anti-ice Diagram

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Figure 10-22 Windshield Anti-ice Diagram—Normal Heat

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Figure 10-23 Windshield Anti-ice Diagram—High Heat

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WINDSHIELD WIPERS Separate windshield wipers are mounted on the pilot’s and copilot’s windshields. The dual wipers are driven by a common mechanism operated by a single electric motor, all located forward of the instrument panel.

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The windshield wiper control is located on the overhead light control panel (Figure 10-24). It provides the wiper mechanism with SLOW, FAST, and PARK operations. The wipers may be used either on the ground or in flight, as required.

To protect the NESA coating, the wipers must not be operated on a dry windshield. The windshield wiper circuit breaker is on the copilot’s circuit breaker panel in the WEATHER group.

Figure 10-24 Windshield Wipers

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BRAKE DEICE SYSTEM Brake deice is installed on the main wheels to prevent ice and slush from building up between the wheels and freezing the brakes.

Heated air for brake deicing is supplied by P3 bleed air from the compressor of each engine (Figure 10-25). Engine bleed air is routed to a solenoid-operated shutoff valve in each main gear wheel well, then to a distributor manifold attached to the lower gear assembly. The bleed air is directed to the brakes through orifices around the circumference of each ring of the distributor manifold.

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The brake deice system is controlled by an ON – OFF toggle switch mounted on the pilot’s right subpanel. When the switch is ON, power from the airplane electrical system is supplied through a 5-ampere circuit breaker in the copilot’s circuitbreaker panel to a control module. The control module supplies electrical current to the solenoid shutoff valves on the aft side of the firewall in each wheel well. Both DC power and bleed air must be available before the solenoid will open to allow the hot bleed air to enter the distributor manifold.

A switch which is part of the solenoid shutoff valve provides a signal that illuminates the L or R

Figure 10-25 Brake Deicer

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BRK DEICE ON (green) advisory annunciators when brake deice is selected on.

If the airplane is flown with the control switch for the brake deice system switched ON, a circuit is completed through the left main gear switch to a timer in the control module when the main landing gear is retracted. The timing circuit will cycle the deice system off after a maximum of 12 minutes of operation, closing the solenoid valve in the wheel well to shut off the flow of bleed air to protect brakes and adjacent components in the wheel well from overheating. The brake deice cannot be reactivated until the landing gear is extended and the brake deice switch is cycled to OFF and then ON. Additional protection from damage due to overheating of the brake deice system is provided by the brake deice overtemperature warning system.

In this system, low-pressure lines are routed from the pneumatic air manifold into each wheel well, where the lines terminate in heat-sensitive polyvinyl tubing plugged at the end. If overheating of the brake deice system should occur, the tubing will melt, thus relieving the pressure in the warning system lines. As pressure is relieved, a pressure switch, which is tapped off from the EVA tubing in each wheel well, will activate L BK DI OVHT and or R BK DI OVHT lights in the annunciator panel. When a brake deice overheat annunciator is illuminated, the brake deice system must be turned off.

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PITOT-STATIC MAST HEAT Two pitot-static masts, located on the nose of the airplane, contain heating elements to protect against ice accumulation (Figure 10-26). The masts are electrically heated. Each heating element is controlled by a two-position, circuit breaker switch placarded PITOT – LEFT RIGHT, located next to the stall warning heat switch. The down position is OFF, and the up position is ON. The pitot heat system should not be operated on the ground, except for testing or for short intervals to remove snow or ice from the mast. Pitot heat should be turned on for takeoff and can be left on in flight during icing conditions, or whenever icing conditions are expected. If, during flight at altitude, the airspeed indicator shows a gradual reduction in airspeed, icing of the pitotstatic mast may be suspected. If turning on the pitot heat restores airspeed, leave the pitot heat on. For many pilots, it is standard practice to keep the pitot heat on during all flights at higher altitudes.

The UE Series incorporate an amber L and R PITOT HEAT light that will illuminate any time insufficient current draw is sensed. The lights would also be illuminated any time the switch is OFF. The other 1900 models do not have a failure indication.

Figure 10-26 Pitot Masts and Heat Controls

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ALTERNATE STATIC-AIR SYSTEM The pilot’s or copilot’s alternate static-air source is used any time the normal static-air source is obstructed. The alternate static-air sources are button shaped and are located on each side of the lower fuselage below the cockpit. Alternate static heat is applied by moving the circuit-breaker control switch on the ice protection control panel to the ALT STATIC position.

When the airplane has been exposed to moisture and/or icing conditions (especially on the ground), the possibility of obstructed pitot-static masts should be considered. Partial obstructions will result in the rate-of-climb indication being sluggish during a climb or descent, inaccurate airspeed indications, and incorrect altimeter indications. A suspected obstruction is verified by switching to the alternate system and noting a sudden sustained change in the rate of climb.

This may be accompanied by abnormal indicated airspeed and altitude changes beyond normal calibration differences.

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ALTERNATE STATIC HEAT The alternate static-air system is protected from icing by the alternate static heat system. Alternate static heat is controlled by a 5-ampere circuit breaker switch on the ice protection panel.

The switch controls heat for both alternate staticair ports.

The alternate static system is the primary static source for the differential pressure gage and pneumatic pressure gage in Series UA, UB, and UC. In Series UE, the differential pressure gauge and differential pressure switch are connected to the alternate static source, therefore alternate static heat should be used under the same conditions that require use of pitot-static heat.

Whenever an obstruction exists in the normal static-air system the pilot’s and/or copilot’s alternate static-air source can be selected. When using the alternate system, the POH/AFM should be consulted for the corrections to airspeed and altimeter indications. In general, whenever the alternate system is selected, the aircraft is actually lower and slower than indicated by the aircraft’s flight instruments.

Be certain the static-air selector switches are in the NORMAL position when the alternate system is not needed.

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CHAPTER 11 AIR CONDITIONING CONTENTS Page INTRODUCTION ................................................................................................................ 11-1 GENERAL ............................................................................................................................ 11-1 ENVIRONMENTAL SYSTEM DESCRIPTION ................................................................ 11-3 ENVIRONMENTAL SYSTEM CONTROLS ..................................................................... 11-3 ENVIRONMENTAL BLEED AIR AND CONTROL......................................................... 11-5 CABIN AIR DISTRIBUTION ............................................................................................. 11-5 UNPRESSURIZED VENTILATION................................................................................... 11-7 HEATING (SERIES UA, UB, UC) ...................................................................................... 11-8 HEATING (SERIES UE)...................................................................................................... 11-8 COOLING (SERIES UA, UB, UC)...................................................................................... 11-8 COOLING (SERIES UE) ..................................................................................................... 11-8 COOLING (ALL) ............................................................................................................... 11-10 AIR-CYCLE MACHINE SYSTEM ................................................................................... 11-10 VAPOR-CYCLE SYSTEM ................................................................................................ 11-11 VCS (Series UA, UB, UC) .......................................................................................... 11-11 VCS (Series UE) ......................................................................................................... 11-11 VCS (All) .................................................................................................................... 11-11 TEMPERATURE CONTROL............................................................................................ 11-11 ENVIRONMENTAL CONTROLS .................................................................................... 11-12 Automatic Mode Control............................................................................................. 11-13 Manual Mode Control ................................................................................................. 11-14

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MANUAL COOL Mode Control (Series UE)............................................................ 11-14 ENVIRONMENTAL SYSTEM PROTECTION AND ANNUNCIATORS..................... 11-14 TEST FUNCTIONS ........................................................................................................... 11-15

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ILLUSTRATIONS Figure

Title

Page

11-1

Environmental System Schematic ......................................................................... 11-2

11-2

Ram-Air Scoop ...................................................................................................... 11-3

11-3

Bleed-Air System Schematic ................................................................................. 11-4

11-4

Environmental Annunciators ................................................................................. 11-4

11-5

BLEED AIR VALVES Switches .......................................................................... 11-5

11-6

Cabin “Eyeball” Outlets ........................................................................................ 11-5

11-7

Environmental Air Distribution ............................................................................. 11-6

11-8

Cockpit “Eyeball” Outlets ..................................................................................... 11-7

11-9

Cabin Floor Outlets................................................................................................ 11-7

11-10

Ram-Air Door and Solenoid Valve ....................................................................... 11-7

11-11

VENT AIR Control................................................................................................ 11-7

11-12

Precooler and Valves ............................................................................................. 11-8

11-13

Air-Cycle Machine Schematic............................................................................... 11-9

11-14

Vapor-Cycle Cooling System .............................................................................. 11-10

11-15

Environmental Group Switches and Knobs......................................................... 11-12

11-16

BLOWERS Switch .............................................................................................. 11-12

11-17

Air Control Knobs ............................................................................................... 11-13

11-18

MODE CONTROL Selector Switch ................................................................... 11-13

11-19

Cabin Temperature Rheostat ............................................................................... 11-13

11-20

MAN TEMP Switch ............................................................................................ 11-14

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CHAPTER 11 AIR CONDITIONING

INTRODUCTION Passenger comfort and safety is of prime importance. The following chapter teaches operators of the 1900 Airliner to use the aircraft environmental systems effectively and within the limitations of the system.

GENERAL This chapter describes air conditioning, bleed-air heating, and fresh air systems. Each component discussion includes general description, princi-

ples of operation, system controls, and emergency procedures.

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Figure 11-1 Environmental System Schematic

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ENVIRONMENTAL SYSTEM DESCRIPTION “Environmental System” refers to the devices which control the pressure vessel’s environment.

Along with ensuring air circulation, the system controls temperature by using the heating and cooling devices as needed.

Pressurization and air conditioning systems operate in conjunction with each other and as separate systems. Together they maintain the desired cabin pressure and cabin air temperature. The cabin is pressurized, heated, or cooled through common ducts.

The Beechcraft 1900 Airliner environmental system (Figure 11-1) uses engine bleed air not only for cabin pressurization, but also for cabin heating. In addition, bleed air provides the motive force to operate the air-cycle machine (ACM), the primary source of cabin cooling. A vaporcycle system, driven by the right engine, augments air-cycle machine output when additional cooling is required.

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Various modes of environmental temperature control are provided for by the environmental system control circuitry. The system can be cont r o l l e d m a n u a l l y b y t h e fli g h t c r ew, o r automatically by the cabin temperature controller. The evaporator blowers can be operated automatically, or they may be selected independently of temperature control.

Environmental bleed air is distributed and recirculated through two distinctly separate ducting systems, providing unpressurized ventilation or pressurized air which is temperature regulated.

Floor vents deliver bleed air into the cabin. Eyeball outlets in the cabin mid-sidewall provide cool vapor-cycle air or recirculated cabin air to the passengers. On the flight deck, bleed air is ducted through vents under the pilots’ seats, and cool vapor-cycle air is dusted through eyeball outlets in the overhead panel. When operating unpressurized, ventilation can be obtained on demand through a ram-air scoop (Figure 11-2) on the right side of the nose. Unpressurized air is then delivered through the floor vents and recirculated through the eyeball outlets.

ENVIRONMENTAL SYSTEM CONTROLS A system of valves, regulators, and temperature and pressure sensors controls all physical aspects of the bleed air flowing into the cabin. Two outflow valves, modulated by the pressurization controller and mounted on the aft pressure bulkhead, provide a controlled exit for bleed air. The pressurization system maintains a specific and constant pressure differential between the cabin environment and the outside air.

Figure 11-2 Ram-Air Scoop

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Figure 11-3 Bleed-Air System Schematic

Figure 11-4 Environmental Annunciators

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ENVIRONMENTAL BLEED AIR AND CONTROL Bleed air from the P3 stage of the engines is precooled to 450° ±25° F by the precooler heat exchanger mounted immediately aft of the engine oil cooler (Figure 11-3). Bleed air is then regulated to 38 ±2 psi by the environmental regulator/shutoff valve. A precooler bypass and a precooler through-valve modulate the amount of bleed air passing through the precooler. The two valves oppose each other in operation when the through-valve is opening, the bypass valve is closing, and vice versa.

When the BLEED AIR VALVES switch is OPEN, all environmental bleed-air valves (precooler through-valves, precooler bypass valves, and shutoff valve regulators) are energized through the normally closed contacts of the bleed-air valve deactivate relays. If bleed-air pressure is adequate, a temperature controller assumes control of the precooler valves and pneumatically controls the valves to maintain the specified bleed-air temperature prior to reaching the environmental system.

Bleed air entering the cabin is controlled by two switches on the copilot’s subpanel (Figure 11-5) placarded BLEED AIR VALVES OPEN – ENVIR OFF – INSTR & ENVIR OFF. When the switches are OPEN, the bleed-air regulator/shutoff valves and the pneumatic/instrument air valves are both open. When switches are placed in ENVIR OFF, the bleed-air regulator/shutoff valves are closed and the pneumatic/instrument air valves are open. In ENVIR OFF, no bleed air is allowed to enter the environmental system for pressurization or temperature control. When switches are in the INSTR & ENVIR OFF position, all environmental bleed-air valves and pneumatic/instrument air valves are closed. This position eliminates all pressurization, cabin temp e r a t u r e c o n t r o l , a n d a i r- d r iv e n fl i g h t instruments.

Figure 11-5 BLEED AIR VALVES Switches

CABIN AIR DISTRIBUTION There are two separate environmental dusting systems for delivering air to the cabin (Figure 11-7).

Engine bleed air is cooled and delivered to the cabin through outlets in the lower cabin sidewalls, near the floor. The vapor-cycle cooling system recirculates and further cools cabin air as required and distributes the air through adjustable eyeball outlets (Figure 11-6). Air ducted to each eyeball outlet can be directionally controlled by moving the outlet in the socket. Air volume is regulated by twisting the outlet to open or close the damper.

Figure 11-6 Cabin “Eyeball” Outlets

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Figure 11-7 Environmental Air Distribution

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Figure 11-8 Cockpit “Eyeball” Outlets

Two evaporator blowers recirculate cooled cabin air. Evaporator coils are mounted on the inlet side of the evaporator blowers to facilitate the exchange of heat between cabin air and the cooling freon when the vapor-cycle system is operating. The evaporator blowers recirculate cabin air through the eyeball outlets.

Figure 11-9 Cabin Floor Outlets

UNPRESSURIZED VENTILATION Ventilation is available during the unpressurized mode. With one or both bleed-air valves open, air enters the cabin through the floor outlets (Figure 11-9). However, for the cabin to remain unpressurized, the pressurization control switch must be in the DUMP position. Air volume through the floor outlets is regulated by using the CABIN AIR control knob located on the copilot’s subpanel.

Figure 11-10 Ram-Air Door and Solenoid Valve

The second source of fresh air, available only when unpressurized, is ambient air which is supplied through a ram-air door (Figure 11-10).

During pressurized operations, the door is held closed by a solenoid and by cabin pressure. During unpressurized operation, a manually controlled valve located in the nose ram-air duct can be opened to allow air to enter the airplane; however, before the valve can be opened, the pressurization switch must be placed in DUMP. The vent control knob, placarded VENT AIR – PULL ON (Figure 11-11), is located under the copilot’s left subpanel. When the control is pulled out, and the pressurization switch is in dump, ambient air flows into the cabin through the floor outlets.

Figure 11-11 VENT AIR Control

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HEATING (SERIES UA, UB, UC) Engine bleed air, precooled by modulation of the precooler valves (Figure 11-12) to approximately 450° F, enters the heat distribution ducts through two electrically operated rotary valves (the ACM bypass valve and the ejector bypass valve) located adjacent to the ACM.

The two valves operate sequentially. When heating is required, the ACM bypass valve opens first, reducing air-cycle machine output. When more heating is required, the ACM bypass valve opens fully, contacting a valve-limit switch that sends operating current to the ejector bypass valve.

When the ACM bypass valve is thus fully opened, with more heating required, the ejector bypass valve will begin to open. When maximum heating is needed, both bypass valves are completely open.

HEATING (SERIES UE) Engine bleed air, precooled by modulation of the precooler valves (Figure 11-12) to approximately 450° F, enters the heat distribution ducts through

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a simple electrically operated rotary valve adjacent to the ACM. When heating is required, the ACM bypass valve opens reducing air-cycle machine output. When more heating is required, the ACM bypass valve opens fully, allowing maximum heat to be supplied to the cabin.

COOLING (SERIES UA, UB, UC) All cabin cooling is provided by the air-cycle machine (Figure 11-13) and, when required, by the vapor-cycle cooling system (Figure 11-14).

When a cool cabin is required, the ejector bypass valve begins closing first. When the ejector bypass valve is fully closed, a limit switch completes the circuit for operation of the ACM bypass valve. When the ACM bypass valve is fully closed, maximum cooling is available from the ACM.

COOLING (SERIES UE) All cabin cooling is provided by the air-cycle machine (Figure 11-13) and, when required, by

Figure 11-12 Precooler and Valves

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Figure 11-13 Air-Cycle Machine Schematic

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Figure 11-14 Vapor-Cycle Cooling System

the vapor-cycle cooling system (Figure 11-14).

When cabin cooling is required, the ACM bypass valves begin to close. When the ACM bypass valve is fully closed, maximum cooling is available from the ACM.

COOLING (ALL) Vapor-cycle system cooling is turned on by a limit switch on the ACM bypass valve. Therefore, the vapor-cycle system is activated only when the ACM bypass valve is fully closed, and maximum cooling is required.

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AIR-CYCLE MACHINE SYSTEM The air-cycle machine uses engine bleed air to drive a compressor which compresses the air, increasing its pressure and making it hotter. The excess heat of compression is removed through the use of heat exchangers. This cooler, highpressure air is then released through an ejector, and as pressure drops rapidly, the air is cooled.

The sequence in the air-cycle machine is as follows: 1. Bleed air enters the air-cycle machine through the first-stage heat exchanger where excess heat is removed.

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2. Bleed air enters the air-cycle machine compressor where it is compressed to a higher pressure and a hotter temperature.

3. Compressed bleed air passes through the second-stage heat exchanger where the excess heat of compression is removed.

4. Bleed air passes through the air-cycle machine expansion turbine, cooling the air and providing the motive force necessary to drive the ACM compressor.

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VCS (SERIES UA, UB, UC) As commands are again received for warmer air, the ACM bypass valve is the first to begin to open. If heating continues to be required, the ejector bypass valve will also begin to open. A limit switch at the open position of the ACM bypass valve provides the electrical signal which terminates vapor-cycle operation.

VCS (SERIES UE)

5. Bleed air exits the air-cycle machine through the air-cycle machine ejector, where the pressure is released and the bleed air allowed to expand, producing cold air for circulation through the cabin.

In the UE Series, the ejector bypass valve has been removed and control of the vapor-cycle system is handled by a limit switch in the ACM bypass valve. Once activated, the vapor-cycle w i l l o p e r a t e u n t i l t h e AC M va l ve o p e n s completely.

VAPOR-CYCLE SYSTEM

VCS (ALL)

When the air-cycle machine is at maximum cooling, a switch on the ACM bypass valve transmits a signal to the freon air conditioner. When the ACM bypass valve is fully closed, the signal is transmitted to the temperature controlling circuitry. The heat/cool command relay then energizes the freon compressor clutch to initiate vapor-cycle system cooling. Once the vaporcycle system is activated, it will remain in operation until the ACM reaches the full open position.

An outside air temperature sensor and overpressure and underpressure switches are installed to protect the vapor-cycle system. When outside ambient air temperatures are 45 ¹5° F or below, the freon air-conditioning system will not operate. If system pressures exceed maximum or minimum safe limits, overpressure and underpressure switches deactivate the freon air conditioner.

An air-conditioning compressor, driven by the right engine, compresses gaseous freon into a h i g h - t e m p e r a t u r e , h i g h - p r e s s u r e ga s . A condensing coil and blower assembly in the right center wing then removes excess heat, allowing the gas to condense into a liquid state. This highpressure, low-temperature liquid then passes through an expansion valve where the pressure is relieved, and into an evaporator where the liquid freon again becomes a gas. Both the expansion valve and the evaporators cause the freon to cool.

Cabin air is circulated over the evaporator coil where heat is transferred from the cabin air to the gaseous refrigerant. The low-pressure, lowtemperature freon then returns to the compressor, and the entire cycle is repeated.

TEMPERATURE CONTROL Temperature control in the 1900 Airliner can be either automatic or manual. When operating automatically, a cabin temperature controller registers cabin temperatures and signals control of the ACM (all series) and ejector bypass valves (Series UA, UB, and UC).

The solid-state automatic temperature controller is located overhead in the center passenger compartment. In the automatic mode, the cabin temperature controller issues commands to the air-cycle machine (all series) and ejector bypass valves (Series UA, UB, and UC) to control the amount of ACM cooling provided to the cabin.

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The bypass valves can also be modulated manually with a toggle switch on the environmental control panel.

ENVIRONMENTAL CONTROLS The cabin environmental controls are located on the copilot’s left subpanel (Figure 11-15). The environmental system’s controls include bleedair valve switches, a blower control switch, a manual temperature control switch, a cabin-temperature level control, and the environmental mode control switch.

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at night, all cockpit ceiling and cabin eyeball outlets should be closed for maximum cabin heating.

Vent blowers are controlled by a three-position switch (Figure 11-16) placarded BLOWERS – HI – LO – AUTO (HI – OFF – AUTO – UE).

When the blower control switch is in AUTO, the blower is controlled by the mode control selector switch. When the mode control switch is OFF, and the blower switch is in AUTO, the blowers will not operate. When the switch is in AUTO and mode control is in any other position (i.e., MANual or AUTOmatic) (MAN COOL - UE), the vent blowers will operate continuously at low speed.

Four additional manual controls (Figure 11-17) are used to distribute airflow and to regulate the volume of air delivered to cockpit and cabin compartments. These are labeled PILOT AIR, DEFROST AIR, CABIN AIR, and COPILOT AIR. When all control knobs are in the full out position, maximum airflow is distributed to the cockpit. If all controls are pushed in, maximum airflow is provided to the cabin.

When flying in very warm weather, cockpit ceiling outlets and all cabin eyeball outlets should be fully open for maximum cooling. For cold weather flights, particularly high altitude or

Figure 11-16 BLOWERS Switch

Figure 11-15 Environmental Group Switches and Knobs

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Figure 11-17 Air Control Knobs

When the BLOWERS switch is operated in LO or HI, the vent blowers are controlled independently of any other environmental system component, including the mode control switch.

(Note below)

AUTOMATIC MODE CONTROL When the environmental mode control selector switch (Figure 11-18) is in the AUTO position, air-cycle machine and vapor-cycle systems operate automatically, and the cabin temperature

Figure 11-18 MODE CONTROL Selector Switch

controller modulates bypass valve positions to control temperature.

A rheostat knob, placarded AUTO TEMP INCR – CABIN TEMP (Figure 11-19) determines temperature levels when the automatic mode is selected. To select warmer cabin temperatures, rotate the control clockwise, and for cooler cabin temperatures, rotate the knob counterclockwise.

When the rheostat is set, the desired temperature will be maintained automatically.

Figure 11-19 Cabin Temperature Rheostat

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MANUAL MODE CONTROL To select manual control of cabin temperature, the environmental mode control selector must be in MAN. When in the manual mode, cabin temperature is controlled by a spring-loaded toggle switch placarded MAN TEMP – INCR – DECR (Figure 11-20). Moving this switch to the INCR or DECR position manually controls the circuit which provides power to modulate ACM bypass and ejector bypass valves. Approximately 60 seconds (30 seconds – Series UE) is required for the valve(s) to complete their movement toward the full hot or full cold position. Just as in the automatic mode, only one valve moves at a time, varying the amount of conditioned bleed air routed through the heat exchangers and into the cabin. But the pilot, rather than the automatic temperature controller, must now determine whether to increase or decrease cabin temperature.

Figure 11-20 MAN TEMP Switch

When the mode control selector is in MAN, the vapor-cycle air conditioner system will operate, provided the ACM bypass valve is fully closed and OAT is above 45 ±5° F. Therefore, if maximum cooling is required, be sure to allow time for the bypass valve(s) to close completely.

MANUAL COOL MODE CONTROL (SERIES UE) When the mode control selector is placed in the MANUAL COOL position, the vapor-cycle air-

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conditioner system will operate without regard to the position of the ACM bypass valve provided the OAT is above 45 ±5° F.

ENVIRONMENTAL SYSTEM PROTECTION AND ANNUNCIATORS Hot bleed air flows through a heat exchanger and two valves on each engine to precool the air to 450° F before it is ducted into the air-cycle machine or bypassed into the cabin. A bleed-air shutoff valve/regulator for each engine, downstream of the precooler valves, provides master control of bleed-air flow.

Protective mechanisms terminate bleed-air flow if any failure allows bleed-air temperature or pressure to increase without control. Should specific temperature or pressure limits be exceeded, both bypass valves and the regulator/shutoff valve on the affected side will automatically close.

A small surge tank attached to the bleed-air line dampens any surges in bleed-air pressure, and provides a stable reference source for the overpressure limit switch. If the overpressure limit is exceeded due to a malfunction of the shutoff valve regulator, the overpressure limit switch closes, and all three environmental bleed-air valves automatically close. When the overpressure switch closes, a signal is simultaneously transmitted to the annunciator system (See Figure 11-4), illuminating the L or R ENVIR FAIL light and the green (white – Series UE) L or R ENVIR OFF light.

A temperature sensor in the bleed-air line monitors bleed-air temperature and transmits that information to the bleed-air overtemperature detector. If a failure in the precooler system a l l ow s b l e e d - a i r t e m p e r a t u r e t o ex c e e d 500 ±20° F, the detector transmits energizing current to the bleed-air valve deactivate relay, automatically closing all three environmental bleed-air valves. At the same time, a signal is

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generated, illuminating the L or R ENVIR FAIL annunciator and the L or R ENVIR OFF light.

Anytime the bleed-air deactivate relay is energized, opening the bleed-air valve power circuits, a latching circuit holds the relay in that mode. If an overtemperature or overpressure has tripped the bleed-air deactivate relay, it can be reset by placing the bleed-air valve switch in the ENVIR OFF position. The bleed-air valves can then be reopened by moving the switch back to OPEN.

TEST FUNCTIONS The overtemperature circuits can be functionally tested in the T TEST position of the mode control rotary switch on the copilot’s subpanel (See Figure 11-18).

This test simulates an overtemperature condition, causing the overtemperature sensing circuits to shut down the environmental system. The L and R ENVIR FAIL and L and R ENVIR OFF annunciators (See Figure 11-4) will be illuminated immediately after T TEST is selected if the system is operating correctly. The mode control switch should again be returned to AUTO and the bleed-air valve switches to ENVIR OFF then OPEN to reset the system.

Complete procedures for T TEST are outlined in B e f o r e Ta k e o ff ( R u n u p ) c h e c k s i n t h e POH/AFM.

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CHAPTER 12 PRESSURIZATION CONTENTS Page INTRODUCTION ................................................................................................................ 12-1 DESCRIPTION..................................................................................................................... 12-1 PRESSURIZATION SYSTEM ............................................................................................ 12-3 AIR DELIVERY SYSTEM .................................................................................................. 12-4 Cabin Differential Pressure Hi Warning (Series UE) ................................................... 12-7 Cabin Pressure Control.................................................................................................. 12-7 Preflight Check............................................................................................................ 12-10 Climb and Cruise......................................................................................................... 12-10 Descent ........................................................................................................................ 12-10 Malfunctions and Troubleshooting ............................................................................. 12-16

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ILLUSTRATIONS Figure

Title

Page

12-1

Pressurization and Air Conditioning Distribution System .................................... 12-2

12-2

Cabin Altitude for Various Airplane Altitudes Graph—Series UE....................... 12-3

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Bleed-Air Valve Switches ..................................................................................... 12-4

12-4

Cabin-Air Outflow Valve ...................................................................................... 12-5

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Pressurization Control Schematic.......................................................................... 12-6

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Cabin Pressure Switch ........................................................................................... 12-7

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Cabin Altimeter ..................................................................................................... 12-7

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Cabin Climb Indicator ........................................................................................... 12-7

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Pressurization Controller ....................................................................................... 12-8

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Pressurization Control Electrical Schematic ......................................................... 12-8

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Pressurization System Circuit Breakers................................................................. 12-9

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Troubleshooting Information Chart ..................................................................... 12-18

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CHAPTER 12 PRESSURIZATION

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INTRODUCTION Pressurization is desirable in an airplane because it decreases or eliminates the need for supplementary oxygen. In this chapter the pilot learns how the pressurization system operates, how it is controlled, and how to handle system malfunctions.

DESCRIPTION This chapter presents a description of the pressurization system (Figure 12-1). The function of various major components, their physical locations, and the operation of pressurization system

controls are discussed. Where necessary, references are made to the environmental system as it affects pressurization.

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Figure 12-1 Pressurization and Air Conditioning Distribution System

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PRESSURIZATION SYSTEM As airplane altitude increases, outside ambient air pressure decreases. At approximately 12,500 feet the environment outside the aircraft cannot supply enough oxygen for human needs. However, the pressurization system in the 1900 Airliner provides a comfortable cabin environment, containing sufficient oxygen, up to its design ceiling of 25,000 feet.

The pressurization system maintains an inside cabin altitude proportionally lower than the aircraft’s altitude. As shown by the Cabin Altitude for Various Airplane Altitudes graph (Figure 12-2), whenever cabin altitude and airplane altitude are the same, no pressure differential exists.

Whenever cabin pressure is the greater of the pressure, differential is positive. If cabin pressure

PILOT TRAINING MANUAL

is less than that of the outside ambient air, press u r e d i ff e r e n t i a l i s n eg a t ive . M a x i m u m differential is defined as a measure of the highest differential pressure the airplane structure can safely withstand for an extended period of time, expressed in pounds per square inch differential (psid).

The 1900 Airliner maintains a normal differential pressure of 4.8 ± .1 psi (5.1 psi—Series UE), providing cabin altitudes as indicated in Figure 12-1. Although the pressure vessel is designed to withstand a normal maximum differential of 4.8 ± .1 psi (5.1 psi—Series UE), the minimum differential allowable is zero. That is, the airplane structure is not designed to withstand negative differential pressure.

Pressurization and air conditioning systems (Figure 12-1) operate in conjunction with each other, or as separate systems. Together they maintain the desired cabin pressure and cabin air tempera-

Figure 12-2 Cabin Altitude for Various Airplane Altitudes Graph—Series UE

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ture. The cabin is pressurized, heated or cooled through common ducts.

The “Pressure Vessel” is that portion of the aircraft designed to withstand the pressure differential. In the 1900 Airliner, the pressure vessel extends from a forward pressure bulkhead, between the cockpit and nose section, to a rear pressure bulkhead just aft of the baggage compartment. Exterior skins make up the outer seal, and windows are oval for maximum strength. All cables, wire bundles, and plumbing that pass through pressure vessel boundaries are sealed to reduce leaks. The cabin and cargo doors are sealed by using pressurized cabin air to inflate the door seals; therefore, the greater the pressure differential, the tighter the seal.

AIR DELIVERY SYSTEM Bleed air from the engine compressor sections is used to pressurize the pressure vessel. The bleed air is discharged through one of two precooler valves on each engine, distributed through or around a precooler, then routed to the environmental bleed air regulator/ shutoff valve before it is ducted into the environmental bleed air system.

The bleed air regulator/shutoff valves are controlled by switches placarded BLEED AIR VALVES - LEFT - RIGHT; OPEN - ENVIR OFF - INSTR & ENVIR OFF. The bleed air switches are located on the Environmental Control Panel (Figure 12-3) on the copilot’s subpanel. When the switches are in either the ENVIR OFF or the INSTR & ENVIR OFF positions, no bleed air can enter the air cycle machine or the cabin.

When the switches are OPEN, bleed air flows through the environmental regulator/shutoff valves. If a complete electrical failure should occur, both environmental regulator/shutoff valves would fail to the closed position. No more bleed air would enter the pressure vessel and cabin pressure would leak down.

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Figure 12-3 Bleed-Air Valve Switches

Protective devices terminate bleed air flow in the event of failures which would allow the bleed air temperature or pressure to increase without control. Should specific temperature and pressure limits be exceeded, a bleed air valve deactivate relay is energized, closing both precooler valves and the regulator/shutoff valve on the affected side. A signal is simultaneously transmitted to the annunciator system illuminating the ENVIR FAIL and ENVIR OFF annunciators on the affected side. Placing the bleed air switch in ENVIR OFF resets the bleed air valve deactivate relay and restores the power circuits for the precooler valves.

After flowing through the regulator/shutoff valves, and before reaching the cabin, bleed air passes through the air cycle machine for further cooling, or bypasses the ACM if warmer air is required. The temperature-controlled, pressurized air then flows into outlets in the lower cabin sidewalls. Finally, it flows out of the pressure vessel through the outflow valves (Figure 12-4), located on the aft pressure bulkhead.

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Figure 12-4 Cabin-Air Outflow Valve

The temperature-controlled air is delivered to the pressure vessel at a rate which can vary from about 8 to 16 pounds per minute, depending upon ambient temperature and pressure altitude. Pressure within the cabin and the rate of cabinpressure changes are regulated by pneumatic modulation of the outflow valves (Figure 12-5).

The primary function of the vacuum-operated outflow valves is to control the rate at which air can escape from the pressure vessel. In addition, the outflow valves serve three other purposes: 1) to provide positive and negative differential pressure relief; (2) to depressurize the pressure vessel whenever the cabin pressure switch is moved into the DUMP position; and (3) to keep the pressure vessel unpressurized while the airplane is on the ground, with the left landing gear safety switch closed.

Cabin air outflow modulation is controlled by applying vacuum pressure to the outflow valves.

Regulated controller pressure (metered suction) is introduced into the center chamber of the out-

flow valves to oppose the internal spring which holds the valve closed. When this suction is increased (lower chamber pressure), the valve opens. Conversely, as suction is decreased, the outflow valve closes. In normal pressurized flight, the outflow valve is constantly being adjusted to maintain desired cabin altitude as engine power and cabin/ambient pressure differential changes occur. The center chamber pressure stabilizes inside the outflow valves at a slightly lower pressure than the actual cabin pressure once desired cabin altitude is reached. This differential between cabin pressure and the pressure inside the center chamber of the outflow valves holds the valves open at the precise setting appropriate for cabin air inflow.

The 1900 Airliner pressure vessel is not designed to withstand negative pressure differential. The outflow valves are designed to relieve negative differential pressure at 0.1 psi. This prevents outside atmospheric pressure from exceeding cabin pressure during rapid descents, even if bleed air inflow ceases.

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Figure 12-5 Pressurization Control Schematic

The outflow valves also protect cabin pressure Tom exceeding maximum positive differential (4.8 psid) (5.1 psid UE). By comparing ambient (static) pressure to controller chamber pressure, a poppet valve within the outflow valve can sense when maximum differential pressure is reached.

As pressure is applied to that relief valve, vacuum is applied against the normal spring tension of the outflow valve case. If this differential continues to increase, the valve opens and releases controller pressure air to the static line, thereby reducing pressure in the center chamber (blue) and modulating the outflow valve open. As the outflow valve opens, more cabin air is allowed to escape, thus raising the cabin altitude and reduc-

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ing the differential between ambient (static) altitude and cabin altitude.

When the outflow valves function as dump valves, vacuum is introduced into the main outflow valve chamber through the dump solenoid valve. The resulting pressure drop causes the valve to open, allowing cabin air to escape, thus raising cabin altitude. This is known as “dumping� cabin pressure. Dumping cabin pressure will cause the cabin to depressurize. Negative differential protection in both outflow valves ensures that cabin differential pressure remains at zero (unpressurized) and not less than outside atmospheric pressure.

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CABIN DIFFERENTIAL PRESSURE HI WARNING (SERIES UE) In the UE Series Aircraft, additional warning is provided to the pilots in the event the cabin differential should exceed 5.1 psid. To accomplish this, a differential pressure switch monitors inside pressure and compares this pressure to the pressure sensed from the alternate static sources.

Should the differential pressure reach 5.25 psid, the CABIN DIF HI annunciator would illuminate and warn the pilot to shut off the incoming bleed air to prevent damage to the pressure vessel.

CABIN PRESSURE CONTROL The pressurization control switch (Figures 12-6 and 12-10), located forward of the pressurization controller on the pedestal, is placarded CABIN PRESSURE - DUMP - PRESS - TEST. When the switch is in the DUMP position (forward, leverlocked), the outflow valves are held open by vacuum. With both valves completely open, the cabin is depressurized and will remain unpressurized as long as the pressurization control switch is locked in dump. When the switch is held in the TEST position (aft, spring-loaded), the landing gear safety switch is bypassed, allowing the outflow valves to be modulated by the pressurization controller during preflight testing. When the switch is in the PRESS position (center, on) the outflow valves can be controlled by the pressurization controller, so that the cabin will pressurize normally. Pressurization system circuit breakers (Figure 12-11) are located on the copilot’s circuit breaker panel under the heading “Environmental.”

The cabin pressurization indicators are mounted below the caution/ advisory annunciator panel above the power quadrant. Actual cabin pressure altitude (outer scale) and cabin differential pressure (inner scale) are continuously indicated by the cabin altimeter (Figure 12-7). Immediately to the left of the cabin altimeter is the cabin rate of climb indicator (Figure 12-8), labeled CABIN CLIMB. The cabin climb gage continuously indicates the rate at which cabin pressure altitude is changing.

Figure 12-6 Cabin Pressure Switch

Figure 12-7 Cabin Altimeter

Figure 12-8 Cabin Climb Indicator

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An adjustable cabin pressurization controller (Figure 12-9) is mounted in the pedestal. A volume tank is also located in the pedestal to increase the air volume needed at the rate control side of the controller. The pressurization controller commands modulation of the outflow valves.

A dual-scale indicator is mounted in the center of the controller. The outer scale (CABIN ALT) indicates the cabin pressure altitude which the pressurization controller is set to maintain. The inner scale (ACFT ALT) indicates the maximum ambient pressure altitude at which the airplane can fly while continuing to maintain selected cabin altitude. The indicated value on each scale is read opposite the index mark at the forward (top) position of the dial. Both scales rotate together when the altitude selector is rotated.

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The cabin pressurization controller compares current cabin pressure to selected cabin pressure and references vacuum input to provide the desired cabin altitude. A pneumatic relay amplifies vacuum input and directly applies vacuum to the outflow valves. As vacuum is applied to the controller, the controller’s balancing forces move toward equalization. Cabin altitude (pressure) stabilizes at the desired altitude when actual cabin pressure and selected cabin pressure are in balance.

Inside the controller an aneroid bellows provides the reference for all controller action. When the CABIN ALT knob is turned, it moves the bellows and a calibration spring which is connected to the rate diaphragm separating the two chambers in the controller. The position of the diaphragm regulates the amount of suction applied to the upper chamber through a metering valve. The upper chamber is also vented to the cabin through a filter.

Figure 12-9 Pressurization Controller

Cabin pressure altitude is selected by turning the cabin altitude selector knob until the desired setting on the CABIN ALT dial is aligned with the index mark. The maximum cabin altitude which can be selected is -1,000 feet to +12,500 feet MSL. The rate control selector is placarded RATE - MIN - MAX. Rotating the rate control schedules the rate at which cabin pressure altitude changes the current value to the selected value. The rate of change selected may be from approximately 175 to approximately 2,500 feet per minute.

Figure 12-10 Pressurization Control Electrical Schematic

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Figure 12-11 Pressurization System Circuit Breakers

When the controller’s upper and lower chamber pressures are equal and are approximately 0.05 psi less than cabin pressure, the metering valve holds steady suction to the controller to stabilize the controller and cabin pressure. This slightly lower controller pressure is required to hold the outflow valves open (against internal spring tension), and to maintain a dynamic balance between air inflow and cabin outflow.

When the two chamber pressures are not equal, a vacuum provides the dynamics to adjust them to higher or lower pressures, as required, to adjust cabin pressure altitude. The rate at which the chambers equalize their pressures is regulated by the RATE knob. Turning this knob adjusts a needle valve which changes the size of the opening in the line between the two chambers.

results in a temporary increase in controller upper chamber pressure, causing the outflow valves to close.

The preset solenoid valve allows the pilot to adjust the cabin controller prior to takeoff, but prevents the controller from pressurizing the cabin while the aircraft is on the ground. When on the ground with weight on the left landing gear safety switch, the preset solenoid removes the vacuum supply from the controller. Once airborne, the valve opens to admit vacuum, and the controller begins to readjust cabin altitude as requested by the CABIN ALT set knob. With the CABIN PRESS switch in PRESS, the preset solenoid valve is energized (closed) on the ground and de-energized (open) in normal flight.

Refer to the Pressurization Control segment in this chapter for more details.

A request for the cabin altitude to climb (move toward a lower pressure) results in a temporary reduction in pressure in the upper chamber of the controller, commanding the outflow valves to open. Conversely, a request for the cabin altitude to descend (move toward a higher pressure)

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PREFLIGHT CHECK

annunciator advises the pilot of operations which may require the use of oxygen.

During runup, the pressurization system can be checked by using the cabin pressurization switch.

With both bleed-air valves OPEN, adjust the cabin altitude selector knob so that the CABIN ALT dial indicates an altitude at least 500 feet BELOW field pressure altitude. Rotate the rate control selector knob to place the index at the 12 o’clock position. Hold the pressurization switch to TEST until the cabin altitude indicator shows a descent. Release the switch to the PRESS position after normal pressurization operation is confirmed.

Prior to takeoff, the cabin altitude selector should be adjusted so that the ACFT ALT scale on the indicator dial indicates an altitude approximately 1,000 feet above the planned cruise pressure altitude, and the CABIN ALT scale indicates an altitude at least 500 feet above the takeoff field pressure altitude. The rate control selector knob should be adjusted as desired. A setting of 12 o’clock on the index will provide the most comfortable cabin rate of climb. The cabin pressure switch should be checked to ensure that it is in the PRESS position.

CLIMB AND CRUISE As the airplane climbs, cabin pressure altitude climbs at the selected rate of change until the cabin reaches the selected pressure altitude. The system then maintains cabin pressure altitude at that value. However, if the airplane climbs to an altitude higher than the value indexed on the ACFT ALT scale of the controller, the cabin-toambient pressure differential will reach the maximum differential pressure relief setting of the outflow valves. Either or both valves will then override the pressurization controller in order to limit differential pressure. During cruise operation, if the flight plan calls for an altitude change of 1,000 feet or more, reselect the new altitude plus 1,000 feet on the CABIN ALT dial. If cabin pressure altitude reaches 12,500 feet (10,000 feet—Series UE), for any reason, a pressure sensing switch closes, causing the red CABIN ALTITUDE (CABIN ALT HI—Series UE) annunciator to be illuminated. The cabin altitude

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DESCENT During descent, and in preparation for landing, the cabin altitude selector should be set to indicate a cabin altitude of approximately 500 feet above the landing field pressure altitude. The rate control selector should be adjusted as required to provide a comfortable rate of cabin descent.

Again, the 12 o’clock position on the rate control knob will provide a comfortable rate of change in cabin pressure. The airplane rate of descent should be controlled so that the airplane altitude does not catch up with the cabin pressure altitude during the descent. If the cabin is allowed to depressurize before reaching the desired airplane level-off point, the cabin rate of descent will be the same as aircraft rate of descent until the desired airplane altitude is reached. However, it is desirable to set the controller so that the cabin is unpressurized upon reaching pattern altitude, to prevent the possibility of landing while the cabin is still pressurized.

The following pressurization situations are described in order to illustrate normal flight operation of the system. In each case, the given conditions will be outlined on the profile diagram.

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Situation 1 Climb from sea level to FL200, then descend to a field pressure altitude of 1,500 feet.

Conditions ●

Aircraft climbs at 1,000 ft/min to FL150, then 1,500 ft/min to FL200

●

Cabin climbs at 500 ft/min

●

Aircraft descends at 1,000 ft/min

●

Cabin descends at 500 ft/min

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Controller setup for descent—As the aircraft starts to descend, set the outer dial (CABIN ALT) to 2,000 feet altitude (500 feet above field pressure altitude). The rate knob should stay in the 12 o’clock position.

Operation—As the aircraft descends to 2,000 feet in approximately 18 minutes, the cabin descends to 2,000 feet, in approximately 11 minutes. At 2,000 feet, the cabin continues its descent unpressurized to 1,500 feet, descending at the same rate as the aircraft.

Remarks—All settings are normal and the system reacts properly.

Controller setup before takeoff—Prior to takeoff, set the inner dial (ACFT ALT) on the pressurization controller to FL210 (1,000 feet above cruise altitude). The outer dial (CABIN ALT) will show a 6,500 foot altitude. Set the rate knob at the 12 o’clock position.

Operation—As the aircraft climbs to FL200, which will take approximately 18 minutes, the cabin climbs to 6,500 feet in approximately 15 minutes, thus the cabin always stays “ahead” of the aircraft during the climb. (Staying ahead means that maximum differential pressure is not achieved while the aircraft is climbing since the cabin reaches its final altitude of 6,500 feet before the aircraft has completed its climb to FL200).

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Situation 2 Climb from sea level to FL200, then descend to a field pressure altitude of 1,500 feet.

Conditions ●

Aircraft climbs at 1,000 ft/min to FL150, then 1,500 ft/min to FL200

●

Cabin climbs at 500 ft/min

●

Aircraft descends at 1,000 ft/min

●

Cabin descends at 500 ft/min

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Operation—Pressurization is normal until the cabin reaches maximum differential pressure, causing pressure bumps in the cabin. The condition normalizes during descent.

Remarks—By not setting the pressurization controller properly, cabin pressure bumps are likely, and passenger discomfort results. The ACFT ALT dial should be set to at least 1,000 feet above aircraft cruise altitude.

Controller Setup—Same as Situation #1 except set ACFT ALT dial to FL200 (same as cruise altitude) which will put cabin at maximum differential pressure when the aircraft arrives at its final altitude. Set it as in Situation #1 for the descent.

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Situation 3 Climb from sea level to FL200 then descend to a field pressure altitude of 1,500 feet.

Conditions ●

Aircraft climbs at 1,000 ft/min to 15,000 ft then 1,500 ft/min to FL200

●

Cabin climbs at 500 ft/min

●

Aircraft descends at 1,000 ft/min

●

Cabin descends at 500 ft/min

PILOT TRAINING MANUAL

less than the aircraft’s rate. Because the aircraft is cruising at maximum differential pressure, the cabin will be subject to pressure bumps. Once the aircraft begins to descend below 14,250 feet and the cabin pressure is below maximum differential, the condition normalizes.

Remarks—By setting the controller for landing prior to takeoff, a problem similar to that described in Situation #2 occurs, resulting in discomfort to the passengers.

Controller setup—Same situation as #1 except set CABIN ALT dial to landing field pressure altitude plus 500 feet prior to takeoff. No readjustment required for descent.

Operation—Pressurization situation is normal until maximum differential pressure is reached as the aircraft climbs through approximately 14,250 feet. As the aircraft continues to climb, the cabin climbs at a rate proportional to, and only slightly

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Situation 4 The aircraft was held to 5,000 feet for 15 minutes during climb, then cleared to FL200. The aircraft was given a segmented descent to FL150 before being given final descent clearance for landing at sea level.

Conditions ●

Aircraft climbs at 1,000 ft/min to 5,000 ft, levels off for 15 minutes

●

Cabin climbs at 500 ft/min to 4,500 ft, levels off until reset, then climbs at 500 ft/min to 7,500 ft.

●

Aircraft descends at 1,500 ft/min with a 10-min level-off at 15,000 ft before continuing down for landing.

●

Cabin descends at 500 ft/min with a brief level-off at 3,500 ft until being reset for landing, then descends at 500 ft/min.

Controller setup before takeoff—Set CABIN ALT dial to 4,000 feet (1,000 feet below the aircraft’s intermediate level-off altitude of 5,000 feet). This prevents cabin altitude from catching up to aircraft altitude during climb.

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When cleared for final descent—Set the CABIN ALT dial to 500 feet above landing field pressure altitude.

Operation—As the aircraft climbs to 5,000 feet, the cabin climbs to 4,000 feet, thereby maintaining a slight pressurization differential. When the aircraft climbs to FL200, the controller is reset as in Situation #1, and the cabin climbs accordingly.

When the aircraft starts to descend to 15,000 feet and the controller is reset to 16,000 feet, the cabin starts down, leveling at approximately 3,000 feet. This allows the pressurization controller to maintain a cabin pressure below maximum differential, avoiding pressure bumps as in Situat i o n # 2 . U p o n b eg i n n i n g fi n a l d e s c e n t , pressurization occurs normally as in Situation #1.

Remarks—An alternate method for operating the pressurization system, when given the situation described above, is to set the rate knob to minimum until cleared to the final altitude. The rate control can be returned to a normal setting after final clearance altitude is given. However, if the aircraft remains at the intermediate altitude for a longer time than anticipated, it is still possible to create a situation that will result in passenger discomfort. Whatever method is used, forgetting to reset the controller will result in problems similar to those discussed in Situations #2 and #3.

When finally cleared to FL200—Set the ACFT ALT dial to 1,000 feet above assigned flight level as in Situation #1.

When cleared down to 15,000 feet—Set the ACFT ALT dial to 1,000 feet above the newly assigned altitude. The aircraft altitude is now set to 16,000 feet on the dial.

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SITUATION 5 Depart from a high altitude airport (in this case 6,000 feet), cruise at an altitude of 11,000 feet, and land at a sea level airport 15 minutes later.

Conditions ●

Aircraft climbs at 2,000 ft/min to 11,000 ft

●

Aircraft levels off for approx. 5 min

●

Cabin climbs and descends at 500 ft/min

●

Aircraft descends at 1,500 ft/min

Controller setting before takeoff—Set the CABIN ALT dial to 500 feet above the 6,000 foot takeoff field elevation (set cabin altitude at 6,500 feet).

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Operation—The aircraft will continue to climb to altitude while the cabin starts to descend to 500 feet pressure altitude. Aircraft levels at cruise, then descends for landing. By the time the aircraft is ready for landing, the cabin altitude is level at 500 feet (this assumes aircraft altitude vs. cabin altitude does not exceed 4.8 psid) (5.1 UE).

Remarks—It is important to set the cabin pressurization controller for a cabin altitude above takeoff field pressure altitude. If the cabin altitude is set to an altitude lower than the airport elevation at takeoff, a pressure bump will be experienced shortly after liftoff when the cabin begins to pressurize. Once cabin pressurization is stabilized after takeoff, the controller may be reset for landing, provided cruise altitude does not exceed the altitude in the ACFT ALT window.

Operation—After cabin altitude has stabilized at 6,500 feet (approximately one to two minutes after takeoff, depending upon cabin rate of climb) reset the controller to 500 feet above the landing field pressure altitude.

Controller setting during cruise—If the cabin altitude remains at 6,500 feet while flying at a cruise altitude of 11,000 feet, the differential pressure is 1.8 psid. Since only five minutes of the flight is planned at cruise, do not set the CABIN ALT dial higher than 500 feet above landing field pressure altitude.

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MALFUNCTIONS AND TROUBLESHOOTING Pilot controls are simple and straightforward and workload is minimal. The pilot has sufficient controls readily available to either regain control or minimize the effects of most problems.

If pressurization is lost in flight, follow the procedures outlined in the Emergency Procedures section of the POH/AFM. If the cabin climbs above 12,500 ft (10,000 ft—Series UE) pressure altitude, the CABIN ALT (CABIN ALT HI - UE) warning annunciator will illuminate, and pilots should consider the use of oxygen for crew and passengers. See the Oxygen section of this workbook for more details concerning use of the oxygen system.

Most pressurization malfunctions will show up shortly after takeoff. Three general symptoms are most common: ●

Rapid pressurization toward maximum differential

●

Lack of pressurization (i.e., the cabin climbs at the same rate as the aircraft)

●

Cabin leakdown (i.e., the cabin leaks pressurization slowly - 500 ft/min at lowpressure differentials and faster at highpressure differentials.)

The first two symptoms are generally caused by controller, control system or outflow valve malfunctions. The third is normally caused by air inflow problems. Refer to the Troubleshooting Information Chart at the end of this section for specific pressurization problems and recommended action (Figure 12-12).

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The outflow valves should modulate open as directed by controller pressure. If one or both outflow valves are stuck closed, cabin altitude will descend rapidly shortly after takeoff. Possible sources of this problem include: a stuck preset solenoid, a cracked pressurization controller, a diaphragm failure in the pressurization controller, disconnected or leaking plumbing, a cracked outflow valve, or a failed diaphragm in the outflow valve.

In this situation, the pilot should position both bleed-air valves to ENVIR OFF, stopping P3 bleed-air inflow, thus repressurizing the cabin at its leak rate. Once the cabin is stabilized, cycling the CABIN PRESS switch to TEST might free a stuck preset solenoid valve. Any additional troubleshooting should be accomplished on the ground.

Failure to pressurize after takeoff can be caused by: leaving the CABIN PRESS switch in DUMP, failure of the left main gear squat switch, or failure of the dump solenoid. The pilot should first check to ensure the CABIN PRESS switch is in PRESS. If it is already in PRESS, move the switch to TEST to override the squat switch. If the cabin begins to pressurize, hold the switch in TEST until cabin differential pressure exceeds 0.5 psid, then pull the PRESS CONTROL circuit breaker on the copilot’s circuit breaker panel.

Remember that the DUMP switch is inoperative as long as the circuit breaker is pulled. If the cabin does not pressurize with the CABIN PRESS switch in TEST, the problem is beyond the capability of inflight troubleshooting.

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Failure of the cabin to pressurize shortly after takeoff may also indicate inflow and/or outflow malfunctions. If the cabin altitude climbs with the airplane, one or both outflow valves could be fully open. One cause is premature opening of the preset solenoid. If the solenoid valve opened while still on the ground, the controller will hold the outflow valves wide open until the aircraft “catches up” to the controller during the climb. If this condition is suspected, the pilot can decrease cabin climb rate to minimum, or select a lower cabin altitude. However, be sure to reselect the proper setting once the cabin begins to pressurize normally.

PILOT TRAINING MANUAL

Unusual or excessive pressure bumps may be caused by sticking outflow valves, due to the buildup of contaminants on the valve seats. The filters associated with this system can also become contaminated causing large differences in cabin climb versus cabin descent without a change in rate knob position. These valves and filters should be checked at regular maintenance inspections, or more often if unusual conditions (heavy smoking, dusty atmosphere) exist.

Cabin leakdown because of air inflow problems is often caused by malfunctions of the bleed-air modulation system (precooler valves and bleedair regulator/shutoff valve). Check to see that the bleed-air valve switches are in the OPEN position. If both switches are open, the bleed-air modulation system may be inoperative or an electrical failure may have caused the environmental bleed-air valves to close. Check to see that the bleed-air control circuit breakers on the copilot’s circuit breaker panel have not tripped.

Reset the circuit breaker, if possible. If no reset is possible, continue the flight at a lower altitude or use oxygen if necessary. Maintenance troubleshooting procedures are outlined in the maintenance manuals.

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PILOT TRAINING MANUAL

Figure 12-12 Troubleshooting Information Chart—1 of 3

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Figure 12-12 Troubleshooting Information Chart—2 of 3

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Figure 12-12 Troubleshooting Information Chart—3 of 3

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See Chapter 14, “Landing Gear and Brakes,” for information on the hydraulic power systems.

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CHAPTER 14 LANDING GEAR AND BRAKES CONTENTS Page INTRODUCTION ................................................................................................................ 14-1 GENERAL ............................................................................................................................ 14-1 LANDING GEAR................................................................................................................. 14-2 Landing Gear Assemblies ............................................................................................. 14-2 Wheel Well Door Mechanisms ..................................................................................... 14-3 Hydraulic Landing Gear ................................................................................................ 14-3 Landing Gear Extension and Retraction ....................................................................... 14-5 Hydraulic Fluid Level Indication System ..................................................................... 14-7 Landing Gear Controls .................................................................................................. 14-8 Position Indicators......................................................................................................... 14-9 Landing Gear Warning System ................................................................................... 14-11 Manual Landing Gear Extension................................................................................. 14-11 Hydraulic Schematics.................................................................................................. 14-12 Tires............................................................................................................................. 14-17 Shock Struts................................................................................................................. 14-17 Landing Gear Operating Limits .................................................................................. 14-17 Landing Gear Switch Circuits ..................................................................................... 14-17 NOSEWHEEL STEERING ................................................................................................ 14-18 Manual Steering System.............................................................................................. 14-18 Power Steering System (Airight)—Series UA and UB............................................... 14-18 Power Steering System (Decoto)—Series UC and UE ............................................... 14-21

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WHEEL BRAKES ............................................................................................................. 14-23 Brake System .............................................................................................................. 14-23 Parking Brake ............................................................................................................. 14-23 Brake System Servicing.............................................................................................. 14-24 Brake Wear Limits...................................................................................................... 14-25 Brake Deice System.................................................................................................... 14-26 Cold Weather Operation ............................................................................................. 14-28 Antiskid System .......................................................................................................... 14-28

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ILLUSTRATIONS Figure

Title

Page

14-1

Main Gear Assembly ............................................................................................. 14-2

14-2

Nose Gear Assembly ............................................................................................. 14-2

14-3

Main Gear Door Mechanism ................................................................................. 14-3

14-4

Landing Gear Electrical Circuitry ......................................................................... 14-4

14-5

Hydraulic Landing Gear Plumbing Schematic ...................................................... 14-4

14-6

Checking Hydraulic Fluid Level ........................................................................... 14-5

14-7

Hydraulic Landing Gear Schematic....................................................................... 14-6

14-8

Hydraulic Landing Gear Powerpack ..................................................................... 14-7

14-9

Hydraulic Fluid Low Indicator .............................................................................. 14-7

14-10

Landing Gear Control Switch Handle ................................................................... 14-8

14-11

Safety Switch ......................................................................................................... 14-8

14-12

Landing Gear Position Indicator Assembly........................................................... 14-9

14-13

Landing Gear Position Indicator—No Illumination .............................................. 14-9

14-14

Landing Gear Control Switch Handle and Red In-Transit/Unsafe Indicator ...... 14-10

14-15

Landing Gear Handle Light Test ......................................................................... 14-10

14-16

Landing Gear Alternate Extension Placard ......................................................... 14-11

14-17

Landing Gear Relay Circuit Breaker ................................................................... 14-12

14-18

Landing Gear Retraction Schematic .................................................................... 14-13

14-19

Landing Gear Extension Schematic..................................................................... 14-15

14-20

Hand Pump Emergency Extension Schematic .................................................... 14-16

14-21

Power Steering Controls—Series UA and UB .................................................... 14-19

14-22

Power Steering System Schematic—Park and Taxi Mode.................................. 14-20

14-23

Power Steering Controls—Series UC and UE..................................................... 14-22

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14-24

Brake System—Left Brake Applied ................................................................... 14-24

14-25

Brake System—Parking Brake Set ..................................................................... 14-25

14-26

Brake Fluid Reservoir ......................................................................................... 14-26

14-27

Brake Wear Diagram........................................................................................... 14-26

14-28

Brake Deice System ............................................................................................ 14-27

14-29

Brake Deice Controls .......................................................................................... 14-28

14-30

Antiskid Controls ................................................................................................ 14-29

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CHAPTER 14 LANDING GEAR AND BRAKES

INTRODUCTION An understanding of the landing gear system will aid the pilot in the proper handling of landing gear operation and emergency procedures. This section, in addition to describing the overall system, identifies inspection points and abnormal procedures.

An understanding of the brake system will help the pilot operate brakes safely and with minimum brake wear. This section, in addition to describing the brake system, points out operating and servicing procedures.

GENERAL The landing gear system section presents a description and discussion of the landing gear system, landing gear controls, and system limitations. Landing gear indicating and warning systems and alternate landing gear extension operations are also described. In addition, this section is followed by a discussion of the standard manual and optional power steering systems.

The wheel brake system section presents a description and discussion of the wheel brake system. Correct use of the primary braking system and parking brakes is described. An overview of the brake system, brake deice system, and brake inspection procedures is also presented.

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LANDING GEAR

PILOT TRAINING MANUAL

Operation

LANDING GEAR ASSEMBLIES Components

The upper ends of the drag braces and two points on the shock struts are attached to the airplane structure. When the gear is extended, the drag braces are rigid components of the gear assemblies.

Each landing gear assembly (Figures 14-1 and 14-2) consists of a shock strut, torque links, drag brace actuator, wheel and tire, and brake assembly. A shimmy damper is mounted on the right side of the nose gear assembly on all aircraft with manual steering systems. Brake assemblies are located on the main gear.

The landing gear incorporates Beechcraft air/oil shock struts that are filled with both compressed air and hydraulic fluid. Airplane weight is borne by the air charge in the shock struts. At touchdown, the lower portion of each strut is forced into the upper cylinder; this moves fluid through an orifice, further compressing the air charge and thus absorbing landing shock. Orifice action also reduces bounce during landing.

Figure 14-1 Main Gear Assembly

Figure 14-2 Nose Gear Assembly

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A torque link connects the upper and lower portions of the shock strut. The torque link allows strut compression and extension but resists rotational forces, thereby keeping the wheels aligned with the longitudinal axis of the airplane. On the nose gear, the torque link also transmits steering motion to the nosewheel, and nosewheel shimmy to the shimmy damper.

The shimmy damper (manual steering system only) is a balanced hydraulic cylinder that bleeds fluid through an orifice to dampen nosewheel shimmy.

WHEEL WELL DOOR MECHANISMS The nose gear door is hinged at the front and is connected to the nose gear brace with two links.

When the landing gear is retracted, the door is pulled closed, and when it is extended, the door is pushed open.

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The main gear doors are hinged at the sides and are spring-loaded to the open position. As the landing gear is retracted, two rollers on each main gear engage the door actuating cams, pulling the doors closed (Figure 14-3). When the landing gear doors are closed, they cover the top braces of each landing gear. The lower portion of the tires remains exposed when the main gear doors are closed.

HYDRAULIC LANDING GEAR The retractable tricycle landing gear is electrically controlled (Figure 14-4) and hydraulically actuated (Figure 14-5). The individual landing gear actuators incorporate internal mechanical downlocks to hold the gear in the fully extended position. The landing gear is held in the retracted position by hydraulic pressure.

Figure 14-3 Main Gear Door Mechanism

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Figure 14-4 Landing Gear Electrical Circuitry

Figure 14-5 Hydraulic Landing Gear Plumbing Schematic

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Hydraulic pressure is supplied by a hydraulic powerpack located inboard of the left nacelle and forward of the main spar. A hydraulic fill can located in the left center wing section supplies hydraulic fluid to the reservoir in the power pack.

A dipstick in the fill can provides for a visual check of the fluid level (Figure 14-6).

Electrically actuated control valves route the flow of hydraulic fluid to the individual gear actuators.

The landing gear control switch provides electrical power to the control valves.

To prevent accidental landing gear retraction, a safety switch (squat switch) on the right main landing gear interrupts power to the landing gear motor when weight is on the wheels. The safety switch also provides power to a solenoid-operated latch which secures the landing gear control handle in the down position while the aircraft is on the ground.

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LANDING GEAR EXTENSION AND RETRACTION The landing gear is extended and retracted by a hydraulic powerpack in conjunction with hydraulic actuators located at each landing gear (Figure 14-7). The hydraulic powerpack (Figure 14-8) consists of a hydraulic pump, a 28-VDC motor, a two-section fluid reservoir, filter screens, a fourway gear selector valve, an up and down selector solenoid, a fluid level sensor, and a retract pressure switch.

The landing gear can also be extended manually with a hand-operated pump. The pump handle is located on the floor between the pilot’s seat and the pedestal.

Hydraulic lines for normal extension and retraction are routed from the powerpack to the nose and main gear actuators. Hydraulic fluid for manual gear extension is routed from the hand pump to the nose and main gear actuators through an independent system of tubing. The normal and manual extension lines are connected separately to the upper end of the nose gear actuator and to the lower end of the main gear actuators.

The hydraulic lines for gear retraction are fitted to the lower end of the nose gear actuator and to the upper end of the main gear actuators.

Hydraulic pressure is generated by the powerpack pump and supplied to actuator pistons which are attached to the drag braces; thus the gear is retracted and extended by hydraulic pressure.

When the actuator pistons have fully extended the landing gear, internal mechanical locks in all three actuators secure the gear in the down position. The internal locking mechanisms energize the actuator downlock switches, interrupting electrical current to the hydraulic pump motor.

The pump motor will continue to operate until all three landing gears are down and locked.

Figure 14-6 Checking Hydraulic Fluid Level

When the red gear-in-transit lights in the LDG GEAR CONTROL handle are extinguished, and all three green GEAR DOWN, NOSE–L–R indicators are illuminated, the landing gear is in the fully down and locked position.

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Figure 14-7 Hydraulic Landing Gear Schematic

Moving the landing gear control handle up energizes a solenoid mounted on the powerpack. As the solenoid is powered, the gear selector valve is actuated, allowing hydraulic fluid to flow to the retract side of the actuators. The actuators unlock when 200 to 400 psi of hydraulic pressure is applied to the retract port of the actuators. The gear selector valve is energized in either the up or the down position of the landing gear control handle.

Hydraulic system pressure holds the landing gear in the retracted position. When hydraulic pressure reaches approximately 2,775 psi, the retract pressure switch will cause the landing gear relay to open, interrupting electrical current to the pump motor. The same pressure switch will actuate the pump motor if hydraulic pressure drops below approximately 2,320 psi.

14-6

The landing gear control circuit is protected by a 2-ampere circuit breaker located on the pilot’s inboard subpanel. Power for the pump motor is supplied through the landing gear motor relay and a 200-ampere current limiter (UB-1 through UB-36), or a 60-ampere relay circuit breaker (UB-37 and after). The motor relay and circuit protection are located in the left nacelle power distribution panel. Electrical current through the 2-ampere circuit breaker and through the downlock switches energizes the motor relay.

The landing gear powerpack is powered off the center bus, and when activated, the battery bus tie HED is momentarily desensitized to prevent inadvertent opening of the battery bus tie during landing gear operation.

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Figure 14-8 Hydraulic Landing Gear Powerpack

During gear extension or retraction, if the powerpack motor runs longer than 16 ±0.5 seconds, a time delay will open the pump motor relay and stop the motor. This action also shorts out the 2ampere control circuit and trips the LANDING GEAR RELAY circuit breaker.

HYDRAULIC FLUID LEVEL INDICATION SYSTEM An annunciator placarded “HYD FLUID LOW” (Figure 14-9) in the caution annunciator panel is illuminated whenever the hydraulic fluid level in the landing gear powerpack reservoir is low. The fluid level sensor inside the powerpack senses the level of light within the reservoir, and provides the signal which illuminates the annunciator. The annunciator and the low level sensor can be tested by pressing the annunciator PRESS TO TEST button on the glareshield. The system is operating correctly when the annunciator is illuminated within four seconds after pressing the button.

If the low fluid annunciator illuminates in flight, the gear can be extended with the normal system upon reaching the destination. After the aircraft is on the ground, the hydraulic fluid level should be checked visually with the dipstick in the fill can. If the landing gear fails to extend normally, the gear can be extended with the alternate extension system.

Figure 14-9 Hydraulic Fluid Low Indicator

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LANDING GEAR CONTROLS The landing gear powerpack pump motor is controlled by the landing gear handle on the pilot’s right subpanel (Figure 14-10), placarded “LDG GEAR CONTROL–UP–DN.” The handle must be pulled out of a detent before it can be moved up or down.

The safety switch (Figure 14-11) on the right main gear torque link opens the landing gear control circuit when the oleo strut is compressed.

This switch also breaks power to a solenoid to activate a mechanical latch which secures the control handle in the down position when the airplane is on the ground. After takeoff, when

Figure 14-10 Landing Gear Control Switch Handle

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weight is off the gear, the downlock mechanism is automatically unlocked as power is applied to the solenoid. If the solenoid or the safety switch circuit malfunctions, press downward on the DOWN LOCK REL button to manually release the mechanical latch. The release button is located left of the landing gear control handle.

Although landing gear safety devices have been installed to prevent inadvertent gear retraction, the handle should not be moved out of the down position while the airplane is on the ground. If the handle is moved UP when weight is on the gear, the landing gear warning horn will sound

Figure 14-11 Safety Switch

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intermittently, and the red warning lights in the control handle will be illuminated (when the MASTER SWITCH is ON). When the handle is returned to the DN position, the red lights will be extinguished, and the warning horn will be silenced.

POSITION INDICATORS Landing gear position is indicated by a single unit containing three green annunciators located on the pilot’s right subpanel (Figure 14-12). The annunciator is marked “NOSE–L–R.” Four

PILOT TRAINING MANUAL

lights, two for the nose and one each for the left or right main gear, are illuminated to indicate the down and locked position of that particular gear.

Absence of illumination indicates that a gear is UP or unlocked (Figure 14-13). The green position indicator lights may be tested by pushing on the light housing. Any of the four bulbs can be replaced with any other annunciator bulb after removing the housing from the subpanel (Series UA, UB, UC); on the UE Series, the lights may be tested by pressing the annunciator test switch.

Removal of the assembly requires a special tool that is supplied with the aircraft loose equipment.

Figure 14-12 Landing Gear Position Indicator Assembly

Figure 14-13 Landing Gear Position Indicator—No Illumination

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Two red, parallel-wired indicator lights in the landing gear control handle are illuminated when the gear is in transit (Figure 14-14) or unlocked.

In addition, if the landing gear aural warning horn has been actuated, the red lights will be illuminated. If the horn has been silenced with the warning horn silence button, the red handle lights will continue to be illuminated until the landing gear is in a safe configuration. When the gear is UP or DN and locked, the red lights should be extinguished. The red warning lights may be checked by pressing the HDL LT TEST button (Figure 14-15) adjacent to the landing gear control handle.

PILOT TRAINING MANUAL

Each gear up-position switch is located in the upper portion of the wheel wells. When the gear is fully retracted, the up-position switch is activated, opening the circuit from the control handle warning lights to ground. As soon as the gear begins to extend, the up-position switch provides a path to ground, illuminating the red lights in the gear control handle. The lights go out when the drag brace on each landing gear actuates its respective down-position switch. All three gears must be down and in contact with the down-position switches to extinguish the red warning lights.

Figure 14-14 Landing Gear Control Switch Handle and Red In-Transit/Unsafe Indicator

Figure 14-15 Landing Gear Handle Light Test

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When illuminated, the red lights in the landing gear control handle indicate one or all of the following conditions:

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CONTROL handle. The lights in the LDG GEAR CONTROL handle will remain illuminated until the landing gear is down and locked, or until the power levers are advanced. Advancing the power levers above 84 to 86% N1 rearms the landing gear warning system.

●

Landing gear handle is in the UP position and the airplane is on the ground with weight on the landing gear.

●

One or both power levers are retarded below approximately 84 to 86 ±1% N1, and one or more landing gears are not down and locked. Aural warning horn will sound.

With the flaps beyond the APPROACH position, the warning horn and landing gear switch handle lights will be activated regardless of power lever position. In this configuration, neither the lights nor the horn can be canceled with the warning horn silence button.

●

Any one or all three landing gears are not in the fully retracted or down-and-locked position.

MANUAL LANDING GEAR EXTENSION

●

Warning horn has been silenced with the warning horn silence button with the gear unsafe for landing.

A hand pump, placarded “LANDING GEAR ALTERNATE EXTENSION” (Figure 14-16), is located on the floor between the pilot’s seat and the pedestal. The pump is located under the floor, below the handle, and is used when manual extension of the gear is required. The hand pump manually provides hydraulic pressure to the landing gear system. The landing gear cannot be retracted manually in flight.

Thus, the function of the landing gear handle lights is to indicate that the landing gear is in transit, the position of the landing gear does not agree with that of the handle, or the landing gear warning horn has been silenced and not rearmed.

The gear handle lights, gear-down indicators, and warning horn systems are completely independent of each other. A malfunction in any one system should leave the other two systems unaffected.

LANDING GEAR WARNING SYSTEM The landing gear warning system is provided to warn the pilot when the landing gear is not down and locked. The warning system is interconnected with flap position sensing.

When the flaps are UP or at APPROACH, and either or both power levers are retarded below 84 to 86% N1, the aural warning horn will be activated (the pilot will hear an intermittent warning tone), and the red landing gear control handle lights will be illuminated. The horn can be s i l e n c e d b y p r e s s i n g t h e WA R N H O R N SILENCE button adjacent to the LDG GEAR

Figure 14-16 Landing Gear Alternate Extension Placard

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To operate the system, first pull the 2-ampere LANDING GEAR RELAY circuit breaker (Figure 14-17) below and to the right of the landing gear control handle, and place the control handle in the DN position. This ensures that power is removed from the landing gear powerpack motor circuit. Then remove the pump handle from the securing clip, and pump the handle up and down until the green NOSE–L–R gear-down indicator lights are illuminated and further resistance is felt.

Approximately 80 up-and-down strokes will be required to move the gear to the down and locked position. Finally, place the handle in the fully down position, and secure it in the retaining clip.

PILOT TRAINING MANUAL

If all three green lights are not illuminated, or if the red light remains illuminated in the handle, the landing gear should be considered to be unsafe. If this situation occurs after manually extending the gear, the handle should not be stowed, and the pilot should ensure that sufficient resistance remains on the pump handle, indicating pressure in the hydraulic system.

After a practice manual extension of the landing gear, the gear may be retracted by pushing the LANDING GEAR RELAY circuit breaker in and moving the LDG GEAR CONTROL handle to the UP position.

HYDRAULIC SCHEMATICS

WARNING If for any reason the green NOSE–L–R gear-down indicator does not illuminate (e.g., in the case of an electrical system failure), continue pumping until sufficient resistance is felt to ensure that the gear is down and locked. After an alternate landing gear extension has been made, do not move any other landing gear controls or reset any switches or circuit breakers until the airplane is on jacks and the cause of the malfunction has been determined. The failure may be in the gearup circuit, and the landing gear could retract on the ground.

The hydraulic system schematics shown in this section are for gear extended, gear retracted, hand pump alternate extension, and maintenance retraction modes. Power is shown available to the contacts of the landing gear power relay.

Landing Gear Retraction When the aircraft is airborne, the pilot selects GEAR UP. This completes the circuit from the gear selector switch to the retract pressure switch (Figure 14-18). The pressure switch closes, completing the circuit to the gear-up switch and to the landing gear remote power relay.

Figure 14-17 Landing Gear Relay Circuit Breaker

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Figure 14-18 Landing Gear Retraction Schematic

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When the power relay closes, electrical power is provided to the powerpack pump motor. Electrical power is also routed from the pressure switch to the hydraulic selector valve up-and-down solenoid. Power to this solenoid moves the selector valve to the up position, and routes hydraulic fluid to the retract side of the gear actuator. During retraction hydraulic pressure in the retract line increases until reaching approximately 2,775 psi. The retract pressure switch then opens, interrupting the power circuit to the pump motor, stopping the hydraulic pump. If the pressure switch fails to interrupt power to the motor within 16 Âą0.5 seconds after the gear retraction cycle has begun, a timer will open the power circuit to the motor. A normal retraction cycle is completed in approximately six seconds.

In flight, due to normal hydraulic pressure leakdown, the retract pressure switch closes periodically. As pressure drops to approximately 2,320 psi, the powerpack pump motor is reenergized to restore the retract pressure needed to hold the gear up. Since there are no mechanical uplocks in the landing gear system, pressure is maintained between approximately 2,320 to 2,775 psi to keep the gear in the retracted position. An accumulator, precharged to 800 psi, located in the left wheel well, aids in maintaining pressure in the retract mode.

Landing Gear Extension When the landing gear power relay is open, power is supplied through the 2-ampere control circuit breaker, to the landing gear control switch, and to the three downlock switches. In Figure 14-19, each gear is depicted as down and locked; the three downlock switches are open with no current passing through them. This is the status of the system after a normal gear extension.

PILOT TRAINING MANUAL

Electrical power is also supplied to the selector valve solenoid to route hydraulic fluid to the extend side of the actuators.

Hydraulic pressure pushes the gear down until all three gear actuator downlock switches are depressed. When all three gears are down and locked, the control circuit to the pump motor is deenergized, and the pump motor stops. Notice that no pressure switches are involved. The pump does not cycle when the landing gear is in the down-and-locked position, and hydraulic pressure is not maintained after the downlocks have actuated. Internal mechanical downlocks in each actuator hold the gear in the extended position.

Hand Pump Alternate Extension Hydraulic fluid is pumped from the secondary reservoir by the hand pump and routed through the service valve to the extend side of each actuator through alternate extension hydraulic lines (Figure 14-20). Note that the alternate system is completely independent of the normal extension system.

For manual extension, electrical power to the powerpack pump motor is removed. The landing gear control handle is in the down position, and the control relay circuit is deenergized by pulling the 2-ampere control circuit breaker. When the landing gear has extended fully, the actuator downlocks secure the gear in the down position, and the green gear down-and-locked annunciators will be illuminated. For detailed operational procedures, refer to Landing Gear Manual Extension in the Abnormal Procedures section of the POH/AFM.

For normal gear extension, GEAR DN is selected with the landing gear control handle, completing the circuit from the landing gear control to all three actuator downlock switches, through the service valve, and finally to the landing gear power relay. The power relay closes and provides a power circuit to the powerpack pump motor.

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Figure 14-19 Landing Gear Extension Schematic

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Figure 14-20 Hand Pump Emergency Extension Schematic

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TIRES

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Left Main Gear Squat Switch

Series UA, UB, and UC The nose landing gear wheel is equipped with a single 19.5 x 6.75 x 8, 10-ply-rated, tube-type rim-inflation tire. The nose tire should be inflated to 60 psi +3/–0.

Each main landing gear wheel is equipped with a 22 x 6.75 x 10, 8-ply-rated, tubeless, rim-inflation tire. For increased service life, 10-ply-rated tires of the same size may be installed on the main gear wheels. The main wheel tires are inflated to 95 psi +5/–0.

Series UE The nose gear wheel is equipped with a single 19.5 x 6.75-8, 10-ply-rated tubeless tire. The nose tire should be inflated to 60 psi +5/–0. Main landing gear wheels are equipped with 22 x 6.7510, 10-ply-rated tubeless tires. The main wheel tires should be inflated to 97 psi +5/–0.

Maximum tire speed is increased from 160 to 190 mph.

SHOCK STRUTS Shock struts should always be properly inflated.

Do not over- or under-inflate, and never tow or taxi an aircraft when any strut is flat. Correct inflation is approximately 5.12 to 5.6” for the main strut, and 5.25 to 5.75” for the nose strut.

LANDING GEAR OPERATING LIMITS Consult current issue of POH/AFM for landing gear operating limits.

●

Stall vane heat control

●

Nosewheel steering disconnect—manual steering system only

●

Power steering system connect/disconnect

●

Pressurization system—preset and dump solenoid valves

●

Flight hourmeter ground path

Left Main Gear Drag Brace Switch ●

Brake deice time-delay sequencing

●

Power steering system connect/disconnect

●

Gear-extended signal for green light

Left Main Gear Actuator Switch ●

Landing gear powerpack motor relay and selector valve

●

Gear-locked signal for green light

Right Main Gear Squat/Safety Switch ●

Power steering system—“MAIN STEER FAIL” power

●

Ground idle low-pitch stop system

●

Landing gear powerpack motor relay and selector valve

●

Gear handle downlock solenoid

LANDING GEAR SWITCH CIRCUITS The following list will better acquaint the pilot with the various circuits controlled by the landing gear switches.

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Right Main Gear Drag Brace Switch ●

Gear-extended signal for green light

●

Air conditioner system—condenser blower ground

Right Main Gear Actuator Switch ●

Gear locked and signal for green light

Nose Gear Drag Brace Switch ●

Antiskid system—ground for pressure switch

●

Gear-extended signal for green light

Nose Gear Actuator Switch ●

Gear locked and signal for green light

Up-Position Switches ●

In-transit/unsafe light control

NOSEWHEEL STEERING Either of two systems are available to provide nosewheel steering for the 1900 Airliner.

Although manual steering is standard equipment, most of the aircraft have been provided with the optional power steering system.

MANUAL STEERING SYSTEM In airplanes with manual steering, direct linkage to the rudder pedals permits nosewheel steering when the nose gear is down. One spring-loaded link in the system absorbs some of the force applied to any of the interconnected rudder pedals until the nosewheel is rolling. When the nose wheel begins to roll, the resisting force decreases, allowing more pedal motion to result in more nosewheel deflection.

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Since pedal motion is transmitted to the rudder through cables and linkage, rudder deflection occurs when steering force is applied to the rudder pedals. Therefore, an electrical actuator and cam are incorporated into the steering linkage to remove nosewheel steering from the rudder pedals while the airplane is in flight. After liftoff, nosewheel steering is automatically disconnected. The nosewheel is self-centering upon retraction.

When force on the rudder pedal is augmented with differential power and braking, nosewheel deflection can be increased to its full range of 63°.

POWER STEERING SYSTEM (AIRIGHT)—SERIES UA AND UB The optional power steering system consists of an electric motor-driven hydraulic pump, hydraulic actuator, and servo valve assembly with electronic controls. The shimmy damper, standard with conventional nosewheel steering, is deleted when power steering is installed.

A rotary actuator is mounted on top of the nose gear assembly to rotate the nose gear shock strut and steer the nosewheel. The actuator consists of a two-position solenoid arming valve, a servo valve, a pressure-activated selector valve, and two pistons. A pump and motor assembly located in the left wheel well supplies hydraulic pressure to drive the actuator. Hydraulic fluid for the power steering pump is contained in the landing gear powerpack primary reservoir.

The steering command potentiometers, mounted on the rudder pedals, transmit input from either pilot’s rudder pedals to the power steering signal amplifier.

The nose gear follow-up potentiometers, above the nose gear assembly, monitor the nose gear turning action and relay this information to the signal amplifier. The amplifier then transmits these electrical signals to the actuator servo valve, which controls hydraulic pressure to either side of the nose gear rotary actuator as commanded by the pilot’s rudder pedal steering input.

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Power Steering Control Switch

TAXI Steering Mode

The power steering control switch is located on the lower pedestal and has three positions: OFF, TAXI, and PARK (Figure 14-21). The power steering activation switch is mounted on the left power lever. The power steering control switch must be in TAXI or PARK, and the power steering activation switch must be pressed for power steering to be operative. If not activated by the power lever switch, the steering actuator will remain in the caster mode. When castering, the actuator acts as a shimmy damper.

When the power steering control switch is in TAXI, 10° of power steering is available to the nosewheel after pressing the power steering activation switch.

Steering Modes The power steering system provides two steering modes: TAXI and PARK.

Caster Mode When power steering is not selected or if it becomes inoperative, the nosewheel operates in the caster mode. The pilot may use rudder, braking, and differential power as required to maintain directional control. Power steering can be turned off with the control switch, or it can be immediately disconnected by rapidly moving both power levers forward momentarily above 91% N1.

Annunciator System

PARK Steering Mode When the power steering control switch is in PARK, the power steering system provides 63° of nosewheel steering when activated (Figure 14-22). PARK is a solenoid-held switch position.

If both power levers are advanced beyond 91% Nl, the control switch will revert to the TAXI position, and the system will then be inoperative.

A PWR STEER FAIL annunciator in the caution/advisory panel will be illuminated if hydraulic pressure in the power steering system falls below 650 psi, or if an electrical fault is sensed when the gear is down and locked with the power steering control switch in TAXI or PARK.

Figure 14-21 Power Steering Controls—Series UA and UB

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Figure 14-22 Power Steering System Schematic—Park and Taxi Mode

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A MAN STEER FAIL annunciator in the caution/advisory panel illuminates if the rotary actuator selector valve is in the power steering mode and: ●

The airplane is on the ground with the power steering control switch OFF.

●

The airplane is on the ground with the control switch in TAXI or PARK and the activation switch off.

●

The airplane is in flight with the control switch in OFF or TAXI and the landing gear extended.

●

The airplane is in flight with the control switch in TAXI and the landing gear retracted.

Power Steering Fault Protection The power steering system is protected from electrical failures in the control system by automatic disconnect devices. If power steering becomes inoperative due to an electrical fault in the system, the power steering control amplifier will open. Power steering then reverts to the caster mode, and the PWR STEER FAIL annunciator will be illuminated.

If uncommanded steering occurs at any time, power steering will be automatically disconnected, and the PWR STEER FAIL annunciator will be illuminated.

If power steering does not disconnect when the control switch is placed in the OFF position, or when the power levers have been advanced with power steering selected, the MAN STEER FAIL annunciator will be illuminated. If this should occur, power steering should be selected, and takeoff is not authorized.

The fault protection system can be tested with the power steering test switch on the pedestal, located next to the power steering control switch.

The complete test procedure is in the After Starting engines checklist in the Normal Procedures section of the POH/AFM.

PILOT TRAINING MANUAL

Steering Control Circuits The power steering control circuit is protected by a 5-ampere circuit breaker located on the copilot’s circuit-breaker panel. Power for the power steering pump motor is supplied through the power steering motor relay and 35-ampere limiter. The motor relay, two-position solenoid valve, and actuator arming valve are energized simultaneously by current from the control relay.

The control relay is energized by the power steering control switch and is grounded through the locking relay and power steering activation switch on the left power lever.

The locking relay will retain the ground circuit through the steering disconnect switches as long as either power lever is retarded below 89% N1.

Once both power levers are advanced beyond 91% N1, the control relay is disconnected.

When the airplane leaves the ground, the left main gear drag brace switch and the left and right main gear safety switches open the control circuit, automatically disconnecting power steering.

POWER STEERING SYSTEM (DECOTO)—SERIES UC AND UE Beginning with the UC series, and including the UE series aircraft, an improved power steering system was introduced. This new electrically controlled, hydraulically operated system has many basic operating similarities of the older system but includes a number of hardware and limitation changes. The Decoto system provides two modes of operation: a TAXI mode, ±15° from center, and a PARK mode with a ±55° from center limitation.

This power steering system is electronically controlled and hydraulically actuated, with no mechanical connection between the rudder pedals and the nose gear. The system has three modes of operation: unpowered caster mode, powered TAXI mode providing ±15° of nose gear travel, and powered PARK mode providing ±55° of nose gear travel.

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The system consists of a hydraulic actuator mounted atop the nose gear strut, a hydraulic pump and associated plumbing in the nose wheel well, and an electronic amplifier with associated circuitry located under the copilot’s seat.

Hydraulic fluid is supplied from the brake reservoir in the nose avionics compartment.

The system may be operated by turning the power steering switch ON, selecting either the TAXI or PARK mode, using the two-position toggle switch placarded “POWER STEERING

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PARK–TAXI located on the pedestal, and then pressing the power lever steering switch on the left power lever (Figure 14-23). The system will remain on as long as either power lever is in the low-power position. If both power levers are advanced above approximately 89 to 91% N1, the power steering system is disengaged, and if the mode switch is in PARK, will automatically move to the TAXI position. The power steering system should not be used for any purpose other than parking or taxiing the airplane.

Figure 14-23 Power Steering Controls—Series UC and UE

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A green annunciator (PWR STEER ENGA) is provided on the caution/advisory annunciator panel to indicate when the system is engaged and operating.

Two amber annunciators on the caution/advisory panel are provided to caution the pilot that an abnormal power steering condition exists. An illuminated PWR STEER FAIL annunciator indicates an electrical failure or low system hydraulic pressure. An electrical failure will cause the system to deactivate, while low hydraulic pressure will cause sluggish response but will still permit system operation. An illuminated MAN STEER FAIL annunciator indicates that the nose gear has not returned to the caster mode after attempted disengagement of the power steering system. In the MAN STEER FAIL condition, the nosewheel will remain in the position existing when power was removed. In this circumstance, steering will only be available with power steering turned on.

Refer to the 1900 Airliner POH (Series UC-1 and after) or 1900D AFM (Series UE-1 and after) for operating instructions on this system.

WHEEL BRAKES The main landing gear wheels are equipped with dual, multidisc, hydraulic brakes, actuated by toe pressure on the rudder pedals (Figure 14-24).

Depression of either set of pedals compresses piston rods in master cylinders attached to the pedals. Hydraulic pressure, resulting from movement of pistons in the master cylinders, is transmitted through flexible hoses and fixed aluminum tubing to disc brake assemblies on the main landing gear. The pressure forces brake pistons to press against the linings and discs of the brake assembly.

Each rudder pedal is attached to its own master cylinder. The pilot’s and copilot’s right rudder pedals control the right brake. Similarly, the pilot’s and copilot’s left rudder pedals control the left brake. Differential braking can be used for taxiing and maneuvering on the ground. Either pilot can operate the brakes without taking exclusive control over braking action.

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As with any airplane, proper traction and braking control cannot be expected until the landing gear is carrying the full weight of the airplane. Excessive tire wear can be prevented by using extreme care to control skidding when applying brakes at high speeds. Braking should be applied smoothly and evenly to the end of ground roll.

BRAKE SYSTEM The dual brakes are plumbed in series (Figure 14-24). The pilot’s master cylinders are plumbed through the copilot’s master cylinders, thus allowing either set of pedals to perform braking action, eliminating the need for shuttle valves.

The effect of the brakes is cumulative; neither pilot has exclusive control.

PARKING BRAKE The parking brake holds hydraulic pressure in the brake lines through a set of check valves (Figure 14-25). Dual parking brake valves are installed adjacent to the rudder pedals between the master cylinders of the pilot’s rudder pedals and the wheel brakes. Control for parking brake valves is on the center pedestal. To set the parking brake, depress brake pedals to build up pressure in the brake system, depress the button in the center of the parking brake control, and pull the control handle up. This procedure closes both parking brake valves simultaneously.

Parking brake valves will retain the pressure previously applied to the system. The parking brake can be released by depressing pilot or copilot pedals (to equalize the pressure on both sides of the valves) and pushing down the parking brake handle (to allow the parking brake valve to open).

To avoid damage to the parking brake system, tires, and landing gear, the parking brake should be left off, and wheel chocks should be used to secure the airplane for extended periods. The parking brake system is designed to be used for temporary parking only, since it is not thermally protected. Ambient temperature changes can expand or contract brake fluid, causing excessive brake pressure or too little pressure. Excessive pressure may cause difficulty in releasing the

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parking brake, and low pressure can cause brake release, resulting in potential damage to the aircraft.

BRAKE SYSTEM SERVICING Brake fluid is supplied to master cylinders from a reservoir accessible through the nose avionics

PILOT TRAINING MANUAL

compartment door on the right side of the aircraft (Figure 14-26). The brake fluid reservoir is on the upper left side of the compartment.

Brake system servicing includes maintaining hydraulic fluid level in the reservoir. A sight gage and dipstick are used to observe fluid level. When the reservoir is low, add MIL-H-5606 hydraulic fluid to fill the reservoir to the full mark on the

Figure 14-24 Brake System—Left Brake Applied

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dipstick. Before flight, check all hydraulic landing gear connections for signs of leaks.

Brakes are equipped with automatic brake adjusters. Automatic adjusters reduce brake drag, thereby allowing unhampered roll. Airplanes with automatic adjusters tend to exhibit a softer pedal and a somewhat longer pedal stroke.

PILOT TRAINING MANUAL

BRAKE WEAR LIMITS To check brakes for wear, set the parking brake and measure the distance between the lining assembly and the piston housing (Figure 14-27).

When this distance measures 0.34 inch or more, the brake lining should be replaced.

Figure 14-25 Brake System—Parking Brake Set

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BRAKE DEICE SYSTEM A brake deice system is also installed (Figure 14-28). This system routes bleed-air heat into a distributor manifold attached to each brake. A lever-locked switch on the pilot’s subpanel, labeled “BRAKE DEICE,” activates the system (Figure 14-29). This switch opens two shutoff valves, permitting hot bleed air to enter the distributor manifolds for brake deicing. When the valves are open, the green L and R BK DEICE ON annunciators will be illuminated.

Figure 14-26 Brake Fluid Reservoir

The brake deice system may be operated as required on a continuous basis with the landing gear extended. However, if the BRAKE DEICE switch is not turned off manually after the landing gear has been retracted, a timer automatically turns off the brake deice system. A complete description of the brake deice system is in Chapter 10, “Ice and Rain Protection,” of this training manual and in the POH/AFM Supplements section.

Figure 14-27 Brake Wear Diagram

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Figure 14-28 Brake Deice System

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conditions cannot be avoided while taxiing, keep flaps retracted to avoid throwing snow or slush into flap mechanisms and to minimize potential damage to flap surfaces.

Refer to the FAA-approved Brake Deicing System in the POH/AFM for Emergency, Abnormal, and Normal Operation of the system.

ANTISKID SYSTEM The antiskid system provides a power brake mode for normal braking applications and an antiskid mode for maximum braking performance. In the power brake mode, brake pedal feel is much stiffer and master cylinder pressure is boosted, once it exceeds a preset level, to assist the pilot in braking effort. In the antiskid mode the above characteristics are combined with an antiskid control system to provide maximum stopping performance on a dry or reducedfriction runway while protecting the tires from undue scuffing or blowout. Both modes are available when the antiskid switch is turned ON and the landing gear is down.

The antiskid system is self-contained and completely independent of any other system except for electrical power. The pilot control equipment consists of one two-position antiskid switch located on the console and an annunciator labeled “ANTI-SKID FAIL� to indicate a failure in the antiskid system (Figure 14-30).

Figure 14-29 Brake Deice Controls

COLD WEATHER OPERATION When operating in cold weather, check brakes and tire-to-ground contact for freeze lock-up.

Before taxi, anti-ice solutions can be used on brakes or tires if freeze lock-up has occurred.

However, do not use anti-ice solutions containing oil-based lubricants, because these solutions may decrease brake effectiveness.

When possible, taxiing in deep snow or slush should be avoided. These conditions can force snow and slush into brake assemblies. If these

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The objective of the antiskid system is to closely approach, but not reach, the brake pressure which would produce a skid. This pressure is not constant and varies continuously during any braking process. In operation, the control computer continuously monitors wheel speed information that is transmitted by the transducers located on each wheel. When a skid is imminent, the computer signals the power brake relay valve which, in turn, adjusts brake pressure to obtain optimum braking effectiveness. When skid control is no longer required, the computer reverts to the monitoring mode, and braking forces are totally controlled by the pilot.

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Figure 14-30 Antiskid Controls

The system incorporates test functions that continuously monitor the system. If electrical faults are detected, the annunciator light will inform the pilot that he does not have antiskid protection. If the motor-driven pump should fail, the accumulator will provide sufficient fluid pressure for approximately 10 brake applications, after which the power brake relay valve will revert the system back to master cylinder control, and the lowpressure switch will cause the ANTI-SKID FAIL annunciator to illuminate.

Directional control is maintained with rudder input, nosewheel steering, and, when required, differential braking. A combination of these steering techniques may be used.

During periods of medium to maximum braking effort, steering corrections made with conventional differential (or asymmetric) braking techniques may not produce the desired effect.

The following caution is included in the POH/AFM:

CAUTION Do not use conventional differential braking techniques to maintain directional control during maximum braking performance. REDUCE the pedal force on the side opposite the desired direction of turn.

For airliners with the antiskid system installed, refer to the Supplements section of the POH/AFM, Abnormal and Normal Procedures.

Landing performance charts are different with the antiskid system also, so refer to the Supplementary information for correct data.

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CHAPTER 15 FLIGHT CONTROLS CONTENTS Page INTRODUCTION ................................................................................................................ 15-1 DESCRIPTION..................................................................................................................... 15-1 FLAP SYSTEM .................................................................................................................... 15-1 ASYMMETRICAL FLAP PROTECTION ......................................................................... 15-4 FLAP AIRSPEED LIMITS (SERIES UA, UB, UC)............................................................ 15-4 FLAP AIRSPEED LIMITS (SERIES UE) ........................................................................... 15-5

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ILLUSTRATIONS Figure

Title

Page

15-1

Flap Control System .............................................................................................. 15-2

15-2

Flap Control Lever................................................................................................. 15-3

15-3

Flap Position Indicator........................................................................................... 15-3

15-4

Flap System Circuit Breakers—Series UA, UB, UC ............................................ 15-4

15-5

Airspeed Indicator Markings—Series UA, UB, UC ............................................. 15-4

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TABLES Table

Title

Page

15-1

Airspeed Indicator Markings—Series UE ............................................................. 15-5

15-2

Airspeed Limitations—Series UE ......................................................................... 15-5

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CHAPTER 15 FLIGHT CONTROLS ;;;;;;;;;;;;; ;;;;;;;;;;;;; ;;;;;;;;;;;;; ;;;;;;;;;;;;; ;;;;;;;;;;;;; ;;;;;;;;;;;;; ;;;;;;;;;;;;; 20 ;;;;;;;;;;;;; ;;;;;;;;;;;;; 10 ;;;;;;;;;;;;; ;;;;;;;;;;;;; 5 ;;;;;;;;;;;;; ;;;;;;;;;;;;; 5 ;;;;;;;;;;;;; 10 ;;;;;;;;;;;;; ;;;;;;;;;;;;; 20 ;;;;;;;;;;;;; ;;;;;;;;;;;;; ;;;;;;;;;;;;; ;;;;;;;;;;;;; ;;;;;;;;;;;;; ;;;;;;;;;;;;; ;;;;; ;;;;; ;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;; ;;;;;;;;;;

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5

5

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INTRODUCTION Familiarization with the flap system operation and limits is necessary to provide optimum performance in takeoff, approach, and landing modes. This training unit identifies and describes flap action so the pilot will understand operation, controls, and limits.

DESCRIPTION

FLAP SYSTEM

This chapter presents a description and discussion of the flap system. The four-segment Fowler-type flap system, its controls, and its limits are considered with reference to operation as outlined in the Pilot’s Operating Handbook.

The flaps, two on each wing, are driven by an electric motor through a gearbox mounted on the forward side of the rear spar (Figure 15-1). The motor incorporates a dynamic braking system through the use of two sets of motor windings, which help to prevent overtravel of the flaps.

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Figure 15-1 Flap Control System

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The gearbox drives four flexible drive shafts connected to jackscrews at each flap. The flaps cannot be stopped at an intermediate point.

operated by a potentiometer driven by the right inboard flap. Flap position limit switches are also driven by the right inboard flap.

The UE flap system has the following three positions and degrees of travel: UP (0), APPROACH (17), and DOWN (35). The UA/B/C system has f o u r p o s i t i o n s , U P ( 0 ) , TA K E O F F ( 1 0 ) , APPROACH (20), and DOWN (35). Flap position is registered on an electric indicator on top of the pedestal.

The flap motor power circuit is protected by a 20-ampere circuit breaker, placarded FLAP MOTOR, located on the right circuit breaker panel (under floor boards - UE) (Figure 15-4). A 5-ampere circuit breaker, placarded FLAP IND & CONTROL, for the flap control circuit is also located on this panel.

The flaps are operated by a sliding lever located just below the condition levers on the pedestal (Figure 15-2). Flap travel, from 0ยบ (up) to 35 ยบ (landing) (down UE), is registered on an electric indicator on top of the pedestal near the flap control lever (Figure 15-3). The flap control has a position detent to select 100 flaps for takeoff and 20 ยบ flaps for approach (17ยบ UE). Full flap deflection, is equal to approximately 35ยบ of flap travel. The indicator is

When the flap handle is placed beyond the approach position with the landing gear up, the landing gear warning horn will sound continuously regardless of power lever position. In this configuration, the horn cannot be silenced with the landing gear warning horn silence button.

Returning the flaps to the approach position or extending the landing gear will then silence the gear horn.

Figure 15-2 Flap Control Lever

Figure 15-3 Flap Position Indicator

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ASYMMETRICAL FLAP PROTECTION Split-flap protection is provided by an asymmetrical flap switch system. This switch is rigged to shut off the flap motor for any out-of-phase condition of approximately three to six degrees between adjacent flap segments. This switch is spring-loaded to the normally open position, but is rigged so that the roller cam holds the switch in its momentary (closed) position. This provides electrical continuity to the flap motor when the outboard and inboard flap segments on both sides are parallel and in phase with one another.

FLAP AIRSPEED LIMITS (SERIES UA, UB, UC) Figure 15-4 Flap System Circuit Breakers—Series UA, UB, UC

Airspeed indicator (Figure 15-5) markings show the maximum speeds and operating ranges of the flaps (VFE). The open white triangle indicates a maximum flap speed of 198 KIAS for using the takeoff flap setting. The maximum speed with

Figure 15-5 Airspeed Indicator Markings—Series UA, UB, UC

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flaps at the approach position is 168 KIAS or at the white triangle. The upper limit of the narrow white arc is the maximum speed permissible with flaps beyond approach. Beyond the approach position, the maximum speed is 153 KIAS.

PILOT TRAINING MANUAL

(UE-79 and above, and earlier airplanes in compliance with SB 2512). Airspeed indicator marking are shown in Table 15-2. Airspeed limitations are show in Table 15-1.

Lowering the flaps will produce these results:

FLAP AIRSPEED LIMITS (SERIES UE)

●

ATTITUDE—Nose Up

●

AIRSPEED—Reduced

The maximum speed for flaps to the APPROACH p o s i t i o n ( 1 7 º ) i s 1 8 8 K I A S . B ey o n d t h e APPROACH position, the maximum speed is 143 KIAS (UE-1 through UE-78) or 154 KIAS

●

STALL SPEED—Lowered

●

TRIM—Nose down adjustment required to maintain attitude

Table 15-1 AIRSPEED INDICATOR MARKINGS—SERIES UE RANGE OR MARKING

KIAS VALUE OR RANGE

SIGNIFICANCE

84-188 84-143

Full-flap operating range. Lower Limit is the stalling speed (VSO) at maximum weight with flaps down (35°) and Idle power.

Maximum map extension/extended speed VFE: • Flaps 17° •Flaps 350 (UE-1 through UE-78 not in compliance with S.B. 2512) •Flaps 35° (UE-79 and after, and earlier airplanes in compliance with S.B. 2512)

84-154

Table 15-2 AIRSPEED LIMITATIONS—SERIES UE SPEED Maximum map extension/extended speed VFE: • Flaps 17° • Flaps 35° (UE-1 through UE-78 not in compliance with S.B. 2512) • Flaps 35° (UE-79 and after, and earlier airplanes in compliance with S.B. 2512)

KCAS

KIAS

190 145

188 143

155

154

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REMARKS Do not extend flaps or operate with flaps in prescribed position above these speeds.

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CHAPTER 16 AVIONICS CONTENTS Page INTRODUCTION ................................................................................................................ 16-1 GENERAL ........................................................................................................................... 16-1 AVIONICS POWER DISTRIBUTION ............................................................................... 16-3 KING SILVER CROWN II EQUIPMENT ......................................................................... 16-7 Audio Control System .................................................................................................. 16-7 Communications Transceiver System .......................................................................... 16-9 VOR/LOC/GS Receiver System................................................................................. 16-10 DME System............................................................................................................... 16-10 RNAV System ............................................................................................................ 16-11 ADF System................................................................................................................ 16-11 COLLINS PRO LINE II EQUIPMENT (SERIES UC) ..................................................... 16-13 NAV System ............................................................................................................... 16-13 DME System............................................................................................................... 16-14 COMM System ........................................................................................................... 16-15 ADF System................................................................................................................ 16-15 Transponder System ................................................................................................... 16-15 DB-407 Audio System................................................................................................ 16-17 Slaved Compass Systems............................................................................................ 16-19 AVIONICS INSTALLATION (SERIES UE).................................................................... 16-22 Cockpit Installation..................................................................................................... 16-23

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PITOT-STATIC SYSTEM.................................................................................................. 16-31 Introduction.................................................................................................................. 16-31 Description................................................................................................................... 16-31 Pitot and Static System ................................................................................................ 16-32

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ILLUSTRATIONS Figure

Title

Page

16-1

Nav/Comm Control Panel—Collins ...................................................................... 16-2

16-2

Avionics Master Switch......................................................................................... 16-2

16-3

Alternative Method of Powering the Avionics Bus............................................... 16-3

16-4

Avionics Master Switch ON/OFF Schematic........................................................ 16-4

16-5

Avionics Buses ...................................................................................................... 16-5

16-6

Inverter Power Supply ........................................................................................... 16-6

16-7

Nav/Comm Control Panel—King Silver Crown II ............................................... 16-7

16-8

King Audio Control System .................................................................................. 16-8

16-9

King Communications Transceiver ....................................................................... 16-9

16-10

King VOR/LOC/GS Receiver ............................................................................. 16-10

16-11

King DME System............................................................................................... 16-10

16-12

King ADF System................................................................................................ 16-11

16-13

Pro Line II NAV Control ..................................................................................... 16-12

16-14

Pro Line II DME Control..................................................................................... 16-14

16-15

Pro Line II Single DME Installation.................................................................... 16-14

16-16

Pro Line II Dual DME Installation ...................................................................... 16-14

16-17

Pro Line II COMM Control ................................................................................. 16-15

16-18

Pro Line II ADF Control...................................................................................... 16-16

16-19

Pro Line II TRANSPONDER Control ................................................................ 16-16

16-20

DB-407 Audio Switching Panel .......................................................................... 16-17

16-21

DB-407 Audio System Diagram.......................................................................... 16-18

16-22

Compass System Connections ............................................................................. 16-19

16-23

Slaved Compass System Block Diagram ............................................................ 16-20

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16-24

Collins PN-101 Compass System ....................................................................... 16-22

16-25

EFIS and Controls ............................................................................................... 16-23

16-26

Instrument Panel (Series UE).............................................................................. 16-24

16-27

Attitude and Compass Displays .......................................................................... 16-25

16-28

Compass and Weather Display, Flight Director Controls, and Course Heading Panel .................................................................................. 16-26

16-29

Radio Magnetic Indicator and Standby Horizon................................................. 16-27

16-30

Altitude Alerter, Cabin Briefing, and Transponder Controls.............................. 16-28

16-31

ADF, Navigation, and Communication Controls................................................ 16-29

16-32

Traffic Alert Collision Avoidance System.......................................................... 16-30

16-33

Pitot and Static System Schematic ...................................................................... 16-31

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Pilot’s Static Air Source Valve ........................................................................... 16-32

16-35

Pitot and Static System Schematic — UB/C....................................................... 16-33

16-36

Pitot and Static System Schematic — UE........................................................... 16-34

16-37

Airspeed Calibration - Alternate System Graph ................................................. 16-35

16-38

Altimeter Correction - Alternate System Graph ................................................. 16-36

16-39

Ice Protection Control Panel ............................................................................... 16-36

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CHAPTER 16 AVIONICS

INTRODUCTION Avionics systems, as a vital part of the airplane, are becoming more sophisticated and complex.

These systems lighten the pilot load, particularly during IFR operations. It is therefore important for the flight crew to understand how the various nav/comm systems function, and how to use them effectively. This section describes the standard avionics installation and how it operates.

GENERAL The Beechcraft 1900 Airliner avionics controls, along with the weather radar, are mounted on an isolation panel in the center of the instrument panel, easily accessible to the pilot or copilot.

Individual audio switches, across the top of the p a n e l , c o n t r o l a u d i o t o t h e s p e a ke r s o r headphones.

the main units typically installed in a 1900 Airliner will be addressed.

The Collins Pro Line II remote-mounted avionics package is installed in the Series UC and UE (Figure 16-1). The Series UE aircraft utilize the Collins EFIS-84 four-tube configuration which will be covered briefly in this section.

The King Silver Crown II line of panel-mounted avionic equipment is installed on many 1900 Airliners. Although not all equipment types in the Silver Crown II line will be discussed here, all of

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Figure 16-1 Nav/Comm Control Panel—Collins

Figure 16-2 Avionics Master Switch

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AVIONICS POWER DISTRIBUTION All avionics equipment may be turned on and off by the avionics master switch (Figure 16-2). In the event that this switch fails, power may be restored by pulling the avionics master circuit breaker, located in the upper right corner of the main circuit breaker panel (Figure 16-3).

The 1900 Airliner has three avionics buses (Figure 16-5) to feed DC power to the various types of avionic equipment. To determine specifically what equipment is being fed from a specific bus or power source, refer to the wiring diagram entitled “DC Power Distribution� which is supplied with each airplane. There are, however, some general rules of thumb which usually apply. For example: 1. Items numbered one (e.g., Comm 1, Nav 1, etc.) are fed by the No. 1 avionics bus, which

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in turn is fed from the electrical system triple-fed bus. It is important to note that in the event of a dual generator failure, the items fed by the No. 1 avionics bus would continue to operate for a limited period of time, being fed directly by the battery.

2. Items numbered two (e.g., Comm 2, Nav 2, etc.) are fed by the No. 2 avionics bus. The No. 2 avionics bus is fed by the left generator bus.

3. Additional avionic items which are not fed by the previous buses are fed by the No. 3 avionics bus. The No. 3 avionics bus is fed by the right generator bus.

During a normal engine starting sequence, as each generator is brought on line, the respective bus tie is closed. Therefore, assuming the avionics master switch is turned ON, all avionics systems will receive power from their respective buses under normal circumstances. Also, when

Figure 16-3 Alternative Method of Powering the Avionics Bus

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Figure 16-4 Avionics Master Switch ON/OFF Schematic

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running equipment checks on the ground with the external power switch ON and an APU connected, all three avionics buses will be powered.

In these instances, the bus ties are automatically closed (Figure 16-4).

However, assume the need to make a quick ground check of Comm 2, prior to starting engines, and without an APU connected. In this

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situation, manually close the bus ties with the appropriate switch located on the pilot’s outboard subpanel.

As a general rule of thumb, an APU should be considered essential for running avionic equipment on the ground. For electronic flight instrument system (EFIS) equipped airplanes, the avionic equipment and inverters require approxi-

Figure 16-5 Avionics Buses

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mately 50 amperes of current from the battery.

This amount of current drain would deplete the battery in a short period of time.

Also, especially with EFIS equipment installed, it is desirable to have the avionic nose compartment doors removed to allow sufficient cross ventilation and cooling of the equipment. Particularly during practice sessions with the avionic equipment which exceed 15 minutes in duration.

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AC power is available from either of two 400-Hz inverters. Under normal circumstances, the No. 1 inverter is fed from the left generator bus and the No. 2 inverter is fed from the right generator bus (Figure 16-6). However, in the event that the operating inverter loses power from its appropriate bus, the inverter automatically switches over to the center bus as its power source.

Figure 16-6 Inverter Power Supply

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KING SILVER CROWN II EQUIPMENT

installed in a King Air will be addressed. For additional information on each system, please consult the appropriate pilot’s guide.

The King Silver Crown II line of panel-mounted avionic equipment is installed on many 1900 Airliners (Figure 16-7). Although not all equipment types in the Silver Crown II line will be discussed here, all of the main units typically

AUDIO CONTROL SYSTEM The KMA 24 audio control system consists of a rotary microphone selector switch, speaker and

Figure 16-7 Nav/Comm Control Panel—King Silver Crown II

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phone switches for each receiver installed in the aircraft, and an integral marker beacon receiver with marker beacon lights (Figure 16-8).

appropriate comm receiver must be manually selected each time the microphone selector switch is changed to a different transmitter.

The microphone selector switch connects the microphone to each transmitter installed on the aircraft. On versions of the KMA 24 which do not have the capability of handling audio from a No. 2 ADF, the proper comm receiver audio switch may be automatically selected by simply pushing either the speaker or phone AUTO switch to the ON position and placing the microphone selector switch to the desired transmitter.

Each receiver may be connected to either the speaker and/or the phones by pushing the appropriate alternate action pushbutton switch to the in or ON position.

Versions of the KMA 24 which have the capability of handling audio from a second ADF do not have the AUTO buttons and therefore the

The marker beacon receiver has a HIGH (button pushed in) and a LOW (button out) sensitivity position. The marker beacon lights may also be tested by pushing and holding the test button in.

A built in photo cell automatically adjusts the lighting intensity depending on ambient lighting conditions.

Figure 16-8 King Audio Control System

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COMMUNICATIONS TRANSCEIVER SYSTEM The KY 196 is capable of transmitting and receiving a frequency range of 118.0 through 135.975 MHz in either 25- or 50-KHz steps (Figure 16-9).

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volume level. Pulling out this control “opens up” the receiver squelch circuit, enabling the pilot to hear weaker stations. This might be an appropriate action when attempting to receive a weak transmitter from a distance, such as listening to an ATIS at a distant point.

The large frequency knob changes the frequency to the left of the decimal point while the smaller knob changes the frequency to the right of the decimal point. The smaller knob makes 50 KHz changes when pushed in and 25 KHz changes when pulled out.

The left frequency display indicates the frequency to which the transceiver is actively tuned.

The right display indicates the standby frequency. In order to transfer or swap the two frequencies, the pilot pushes the transfer button momentarily. (NOTE: The frequency selector knob only changes the standby frequency.)

The ON/OFF/VOLUME control switch turns the unit on when rotated clockwise past the initial detent. Further clockwise rotation increases the

Transmitter operation is annunciated by the illumination of the letter “T” located between the active and standby frequencies.

Figure 16-9 King Communications Transceiver

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VOR/LOC/GS RECEIVER SYSTEM Operation of the KN 53 (Figure 16-10) is virtually identical to that of the KY 196 comm transceiver with the following exceptions: 1. Pulling out on the volume control knob activates the Morse code identification circuit, thus allowing the “Ident� to be heard through the audio system.

2. There is no transmit annunciator on this system.

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DME SYSTEM The remote-mounted KN 63 DME with the KD1 572 panel-mounted indicator operates in a straightforward manner. The indicator is capable of displaying DME distance, ground speed, and time-to-station simultaneously (Figure 16-11).

The mode selector allows the unit to be channeled by either Nav 1 or Nav 2. Selecting the HLD (hold) position allows the DME to remain channeled to the previously selected frequency and is annunciated by either H1 or H2 depending on whether Nav 1 or Nav 2 was previously used.

The mode selector also allows the DME to be turned off.

Figure 16-10 King VOR/LOC/GS Receiver

Figure 16-11 King DME System

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RNAV SYSTEM For operational information on the KNS 81 system, refer to the appropriate flight manual supplement.

ADF SYSTEM The KR 87 ADF (Figure 16-12) has two basic modes of operation, ANT (antenna) and ADF. In the ANT mode, the bearing pointer in the RMI/ADF indicator will not point to the station but provides improved audio reception. The ADF mode is used for navigation purposes, allowing the bearing pointer to point to the station. The ADF mode is selected by pushing the alternate action pushbutton in, and the ANT mode is selected by allowing the same pushbutton to remain in the “out” position. The selected mode is annunciated on the left side of the ADF display.

This unit incorporates a BFO (beat frequency oscillator) circuit which allows non-directional beacons to be identified which are not modulating the carrier with audio. These types of stations are sometimes used outside of the United States.

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The BFO circuit, when activated by pushing the BFO pushbutton to the “in” position, generates a 1020-Hz tone which will be heard each time the NDB transmitter is turned on. This allows the Morse code to be identified in a normal fashion.

As with the KY 196 comm and KN 53 nav, two frequencies may be displayed on the KR 87 ADF. The frequency on the left is always the frequency in use, however, the right display window is shared by several different functions. Like the comm and the nav, the right window may display the standby frequency. However, pushing the FLT/ET alternate action pushbutton changes the function of the right display window.

When FLT is annunciated to the right of the right display window, the display is being used to display flight time. Initially, the flight timer begins operation when the unit is turned on. Then, during takeoff, the flight timer is reset to zero and begins counting again when the weight of the aircraft is off the landing gear “squat switch.” The flight timer continues to operate until the weight of the aircraft is once again on the landing gear squat switch, at which time the display is “frozen,” and will remain so until power is removed or the aircraft takes off again.

Figure 16-12 King ADF System

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Pushing the FLT/ET button again will switch the right window to display an elapsed time function.

The elapsed timer may be reset to zero at any time by momentarily pushing the SET/RST (set/reset) button. Elapsed time will continue to accumulate until the SET/RST button is pushed again or power is removed.

The elapsed timer also has a “countdown� mode of operation, which may be initiated by holding the SET/RST button in for approximately three seconds, or until the ET annunciator begins to flash. Now, the countdown time (in minutes and seconds) may be set into the right display by rotating the two concentric knobs which are normally used to change the frequency. Set the minutes with the large knob and the seconds with the small knob.

In order to start the countdown cycle (as when passing the final approach fix) push the SET/RST button. Time remaining will now be continuously

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displayed until the timer reaches zero, at which time it will revert to a count-up mode of operation and will now automatically display the elapsed time above and beyond that which was originally set in. Additionally, when the countdown mode switches to the count-up mode, the right display window will flash for 15 seconds in order to alert the pilot to the fact that he has gone beyond the originally preset time.

With both the flight and elapsed timer, the display will initially be read in minutes and seconds (up to 59 minutes and 59 seconds [59:59]). After the first hour, these timers will display hours and minutes.

Pushing the FREQ (frequency transfer) button initially changes the right window back to the standby frequency display. Subsequent pushes of the FREQ button transfers the standby and in-use frequencies back and forth (flip-flops).

Figure 16-13 Pro Line II NAV Control

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COLLINS PRO LINE II EQUIPMENT (SERIES UC) A new series of Collins CTL control heads is used for the nav, comm, ADF, and transponder.

The ADF-60 and TDR-90 (ADF and transponder units respectively), however, are retained from the earlier Collins Pro Line System.

The Pro Line II family presently consists of a VHF comm (VHF-22), a VOR/LOC/GS/MB receiver (VIR-32), and a DME (DME-42). These units employ many state-of-the-art features, including extensive self-diagnostic capabilities and multiple frequency storage. Some of the features of this equipment will be described here.

For additional information, see the current Collins Pro Line II Pilot’s Guide.

The comm and nav units have many features in common; therefore, we will use the features of the nav (VIR-32/CTL-32) as a building block for the comm, which will be described next. Features common to both will be described under the nav explanation; differences will be pointed out under the comm explanation.

NAV SYSTEM The VIR-32/CTL-32 nav system is comprised of a VOR/localizer receiver, a glide-slope receiver, and a marker beacon receiver, all contained in one “black box” located in the nose avionics compartment.

The nav receiver (Figure 16-13) may be tuned to the correct frequency in any one of three ways:

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2. The PREset (standby) frequency may be initially selected and displayed in the lower frequency window. It may be necessary to cancel the direct tuning mode (described in 1 above) by again holding the ACT pushbutton for approximately three seconds. Once the PREset frequency is displayed in the lower window, it may be transferred up to the ACTive window by holding the XFR/MEM switch to the XFR position momentarily.

3. Up to four frequencies may be placed into the four channel slots of the memory. This is done by repeatedly pressing the XFR/MEM switch to the MEM position until the desired channel number appears in the upper (ACTive) window (e.g., CH-1). Now the frequency may be selected using the two concentric frequency select knobs and will be displayed in the lower (PREset) window.

Once selected, the frequency may be stored by simply pressing the STOre button twice.

Subsequent frequencies/channels may be stored in a similar fashion.

Regardless of the frequency selection method used, when a new frequency is selected, the compare annunciator (labeled ACT) will flash once if, in fact, the VIR-32 receiver has properly tuned to the frequency displayed in the active window. If the compare annunciator continues to flash, a tuning fault is indicated. The test button should be pressed momentarily in order to display the fault and diagnostic code (see Pilot’s Guide for further details).

DME hold may be selected by placing the mode selector switch in the HLD position. This topic will be further discussed under the topic of DME, to be covered later in this section.

1. The active frequency may be tuned directly by first holding down the ACT pushbutton for approximately three seconds. The lower (PRE set) frequency display will be dashed out. The two concentric frequency select knobs will now directly channel the active frequency. Features such as DME hold, preset channels, etc., are still operable in this situation.

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DME SYSTEM By using frequency scanning techniques, the DME-42 is capable of working with up to three DME stations simultaneously (Figure 16-14). It can display DME distance (NM), ground speed (GS), time to station (MIN), and station identification to any one of these stations; however, the DME stays locked onto all three stations.

In a single DME-42 installation, the three frequencies or channels are connected to the Nav 1 and Nav 2 control heads (CTL-32’s) (Figure 16-15).

In a dual DME-42 installation, the No. 1 DME42 is only connected to the No. 1 nav control head. Likewise, the No. 2 DME-42 is only connected to the No. 2 nav control head. In this configuration, each DME-42 is purposefully limited to displaying only two channels (Figure 16-16).

Figure 16-14 Pro Line II DME Control

Figure 16-15 Pro Line II Single DME Installation

Figure 16-16 Pro Line II Dual DME Installation

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COMM SYSTEM

ADF SYSTEM

In most respects, the VHF-22 COMM works just like the features previously explained on the VIR-32 navigation receiver (Figure 16-17). The primary differences are as follows:

The ADF control head also works like the nav control head in many respects, however, the modes on the mode selector switch are appropriate to an ADF (Figure 16-18).

1. There are six frequency memory positions instead of four.

Although the ADF utilizes one of the new Pro Line II control heads (CTL-62), the actual ADF unit is of an older generation and it does not display test codes.

2. In place of the HLD annunciator, there is a TX (unit transmitting) annunciator.

3. SQ OFF (squelch off) replaces HLD on the mode selector.

TRANSPONDER SYSTEM

4. Two short tones indicate a fault. Push the test button to display fault code.

Like the ADF, the transponder unit is of an earlier generation, has a new Pro Line II control head (CTL-92), and does not display test codes (Figure 16-19).

5. Continued turning of the small knob results in 50-kHz steps. When reversed one click, however, a 25-kHz step results.

The transponder control head can store one preselected code, such as 1200, ready for use at the push of the PRE button.

Figure 16-17 Pro Line II COMM Control

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Figure 16-18 Pro Line II ADF Control

Figure 16-19 Pro Line II TRANSPONDER Control

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DB-407 AUDIO SYSTEM The avionics installation has dual DB-407 audio systems which are totally independent of each other (Figures 16-20 and 16-21).

The following operating rules apply to the audio system. These rules will only be listed for the pilot’s audio system. However, they apply equally to the copilot’s audio system.

1. The speaker volume control regulates the speaker audio level.

2. The speaker switch turns the speaker ON and OFF. 3. The phone volume control regulates the headphone audio level.

4. The headphones are operational at all times (as long as they are plugged into their jack).

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one does not necessarily imply a failure of the other.

6. To select any audio source (e.g., Comm 1, ADF, etc.) turn ON the appropriate audio selector switch.

7. The switch labeled VOICE–BOTH–RANGE Figure 16-20 works in conjunction with both the ADF and nav receivers. When in the VOICE position, the voice portion of the audio will be heard and not the Morse code station identification. When in the RANGE (ident) position, only the Morse code station identification will be heard, not the voice portion. When in the BOTH position, both the voice and range portions of the audio will be heard.

If the pilot’s audio system has failed entirely, the pilot may still listen to audio through the copilot’s speaker.

5. The speaker and headphone audio channels are independent of each other and failure of

Figure 16-20 DB-407 Audio Switching Panel

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Figure 16-21 DB-407 Audio System Diagram

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SLAVED COMPASS SYSTEMS The most common compass system for the 1900 Airliner is the King KCS-55A; however, the Sperry C-14A-43 or the Collins MCS-65 or MCS-103 systems could be installed. As far as the pilot is concerned, all of these systems operate in a similar manner. They will be treated as one common system in the following discussion.

Occasionally, a Collins PN-101 system will be installed on the copilot’s side. This system operates in a slightly different manner and will be discussed separately in this section.

KCS-55A, MCS-65, MCS-103, and C-14A-43 Systems From an operational standpoint, all of these systems may be treated identically. All of these systems require 400-Hz electrical power from an inverter. In the unlikely event that both inverters fail, these systems would be inoperative. Each of them has the following components (Figure 16-23).

1. Flux sensor (also called a flux gate or flux valve)—The function of this device is to sense the earth’s magnetic field relative to the airplane and convert that information into an electrical signal which represents the airplane’s magnetic heading.

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2. Slaving amplifier—The magnetic heading signal from the flux sensor is too weak to be used directly; therefore, is amplified (made larger or stronger) by the slaving amplifier. The output signal is now strong enough to directly drive a torquer motor in the directional gyro and thus maintain the gyro rotor in alignment with magnetic north.

3. Directional gyro—Once the gyro rotor is aligned with magnetic north, it will have a natural tendency to stay there for a short period of time, due to a force called gyroscopic rigidity in space. This force will continue to keep the gyro mechanism in relatively good alignment as long as the gyro rotor continues to turn at its design speed.

When the gyro drifts out of alignment (precesses) the condition will be sensed, and the magnetic heading reference information from the slaving amplifier will again drive the gyro rotor back into alignment with magnetic north, using the torquer motor previously described.

4. Horizontal situation indicator (HSI)—The gyro heading information (which should be the same as magnetic heading) is sent to a compass card on the HSI to display the magnetic heading to the pilot. This heading information is then sent from the HSI to the compass card on the opposite radio magnetic indicator (RMI). In this way, gyrostabilized, magnetic heading information is displayed in front of each pilot from two independent sources, the pilot’s and the copilot’s compass systems (Figure 16-22).

Figure 16-22 Compass System Connections

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Figure 16-23 Slaved Compass System Block Diagram

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5. Slaving meter—The slaving meter compares the sensed magnetic heading at the flux sensor (system input) to the displaced magnetic heading at the HSI (system output). The difference, if any, is displayed on the slaving meter by displacement of the slaving needle from the center position (which indicates synchronization or zero error). It is normal for this needle to deviate occasionally due to precession, however, it should always come back to center. If it is displaced to one side for more than approximately one minute, the gyro may be precessing excessively and/or the slaving system may not be doing its job.

In any case, the accuracy of the compass system should be checked by cross referencing the heading information from the opposite system and/or the magnetic compass.

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7. Increase/decrease switch: This is a toggle switch which is spring-loaded to the center (OFF) position. The switch may be used when in the free mode of operation to manually change the directional gyro to the left or right, thus increasing or decreasing the displayed heading information. When in the slaved mode of operation, momentarily holding this switch in either position causes the system to “reset” itself to the fast-slave mode of operation, thereby correcting any displayed error at a rapid rate. This could be helpful if for any reason the gyro had tumbled or precessed excessively.

6. SLAVE/FREE switch—This lever-locking switch is used to select either the slaved or the free mode of operation for the compass system.

This switch should normally remain in the slaved mode of operation. In this mode, when power is initially applied to the system, it will automatically “slave” itself to the correct magnetic heading and remain there throughout the flight, correcting for precession as necessary.

The free mode of operation is generally reserved for occasions when the slaved (automatic) mode of operation has failed and the pilot wishes to revert to a directional gyro mode of operation. This mode may also be used for flight in polar regions where extreme levels of magnetic variation exist. In this mode of operation, the flux sensor and the slaving amplifier are disconnected from the rest of the system. The result is that the pilot now has a directional gyro (which will precess and must be corrected manually using the increase/decrease switch) which uses the HSI to display the heading information from the directional gyro.

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Collins PN-101 System This compass system (Figure 16-24) is frequently installed on the copilot’s side. It has the advantage of being directly powered by the 28volt DC electrical system. If both inverters fail, the system would continue to operate. However, the PN-101 system does not have a manual backup mode of operation (FREE) if the slaving system (flux sensor and/or slaving amplifier) fails.

The PN-101 system does have a fast-slave switch which may be momentarily held in the UP position to initiate the fast-slaving sequence (see fastslaving explanation under increase/decrease switch above). Except for the differences mentioned here, the basic operation of this system is virtually identical to that of the Collins MCS-65, MCS-103, and the Sperry C-14A-43 systems described previously.

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AVIONICS INSTALLATION (SERIES UE) The standard flight instrument system installation in the Beech 1900D Airliner encompasses the fully digitized Collins EFIS-84 four tube configuration and Collins Pro Line II advanced line of digital technology radios.

The primary display system consists of multicolor cathode ray tube (CRT) displays, remote display processor unit and system control units (Figure 16-25). The CRT displays provide conventional electronic attitude director indicator (EADI) and electronic horizontal situation indicator (EHSI) functions which replace existing electromechanical flight instruments.

Figure 16-24 Collins PN-101 Compass System

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Through advanced digital techniques, the Collins Pro Line II radios bring far more information to the cockpit than ever before made possible. The result is greater pilot confidence and increased safety of flight.

Fuel savings and increased payload are achieved through significant weight reduction in the entire avionics package. High reliability has been made possible through the extensive use of microprocessors resulting in fewer total parts for each instrument and radio.

COCKPIT INSTALLATION The standard instrument panel layout includes duplicate EFIS instruments for the pilot and copilot. The NAV/COM radios are installed on the center of the panel available to both the pilot and copilot. The audio control panel is located in the center of the panel above the radios.

Key Operator Benefits Collins 4-Tube EFIS—Identical pilot/copilot panels in an “airline mature” cockpit design with dual flight directors

State-Of-The-Art Cockpit 4 Tube EFIS Figure 16-26 shows identical instrument panel presentations for:

Figure 16-25 EFIS and Controls

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Both pilots w/dual flight directors

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Low pilot workload for more safety

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Less transition training lowers cost

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High reliability/fewer moving parts

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Cockpit voice recorder—standard

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Flight data recorder—standard

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Cabin briefer (heads-up solid state)— standard

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Figure 16-26 Instrument Panel (Series UE)

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Figure 16-27 through Figure 16-32 show flight and navigation instruments, and communication equipment.

Figure 16-27 Attitude and Compass Displays

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Figure 16-28 Compass and Weather Display, Flight Director Controls, and Course Heading Panel

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Figure 16-29 Radio Magnetic Indicator and Standby Horizon

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Figure 16-30 Altitude Alerter, Cabin Briefing, and Transponder Controls

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Figure 16-31 ADF, Navigation, and Communication Controls

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Figure 16-32 Traffic Alert Collision Avoidance System

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INTRODUCTION

DESCRIPTION

The pitot-static system is vital to safe operation of the airplane. This training unit describes the pitot-static system and identifies the instruments which are affected by it. Alternate static air use is discussed, and emergency procedures are outlined.

The Pitot-Static System section of the workbook presents a description and discussion of the pitot-static system. The dual pitot-static system and alternate static air source are described in detail. Abnormal indications are outlined, and performance graphs covering the alternate static air system will be discussed.

STATIC SELECTOR SWITCHES

COPILOT'S PITOT-STATIC TUBE

VSI ALTIMETER AIRSPEED IND VSI

ALTIMETER

ALTERNATE STATIC SOURCE

STATIC SELECTOR SWITCHES AIRSPEED IND

TEST PORTS ALTERNATE STATIC SOURCE PILOT'S PITOT-STATIC TUBE

Figure 16-33 Pitot and Static System Schematic

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Pitot and Static System The pitot and static system (Figure 16-33) provides a source of impact air and static air for operation of the flight instruments. A heated pitot-static mast is located on each side of the upper portion of the nose. Tubing from the left pitot-static mast is connected to the pilot’s airspeed indicator, and tubing from the right pitotstatic mast is connected to the copilot’s airspeed indicator. The pilot’s pitot pressure source is completely independent of the copilot’s pitot pressure sourse.

The normal static system provides sources of static air from both pitot-static probes to both pilot and copilot flight instruments. The static ports are incorporated into the sides of each pitotstatic mast and are open to the atmosphere, providing the source for normal static pressure.

Pitot-static lines are interconnected to provide a redundant static air source to both sides of the airplane. The pilot’s static air source is completely independent of the copilot’s static air source.

If the normal static source fails, alternate static air lines can be selected as the static air source for the pilot’s and copilot’s flight instruments. If, for example, ice accumulations obstruct the static air ports, the alternate source should be selected (Figure 16-34). The alternate line obtains static air from the alternate static air ports, located on the outside lower portion of the fuselage. The static air selector valve for the pilot is located on the left side of the instrument panel, below and to the left of the airspeed indicator.

The copilot’s static air selector valve is located on the right side of the instrument panel, just to the right of the vertical speed indicator. In Series UE aircraft, these Selector Valves have been relocated to the lower side walls just in front of Pilot and Copilot seats. When the alternate air source is required, the toggle switch is moved from the NORMAL to the ALTERNATE (ALTERNATE AND DRAIN - Series UE) position on the pilot’s or copilot’s static air selector valve. The need for drain valves is eliminated (Series UA, UB, UC) since the alternate static buttons are located in the lowest point in each line.

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The pilot’s altimeter, vertical speed indicator, and airspeed indicator are connected to the pilot’s static air source (Figure 16-35). When the system is switched to the pilot’s alternate air source, only the pilot’s flight instruments are affected. When using alternate air, the pilot’s airspeed indicator and altimeter will read higher than actual, and the vertical speed indicator will show a momentary climb.

Alternate static air for copilot’s airspeed indicator, altimeter, and vertical speed indicator is selected independently of the pilot’s normal system. When the alternate static system on the copilot’s side is selected, only the copilot flight instruments are affected. If both pilot and copilot alternate air sources are selected simultaneously, they then share a common alternate static air source.

Airspeed and altimeter indications change when the alternate static air source is selected. Refer to the Airspeed Calibration - Alternate System, and the Altimeter Correction - Alternate System graphs, in the Performance section of the Pilot’s Operating Handbook / Airplane Flight Manual (UE), for correct indications when using the alternate static air source.

Figure 16-34 Pilot’s Static Air Source Valve

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TEST PORTS P

TEST PORTS S

P

S

FUSELAGE STATION 84.00

PITOT-STATIC MAST

PITOT-STATIC MAST

PITOT

PITOT

STATIC

STATIC P

P

S2

S1

S1

S2

MANIFOLD

AS

MANIFOLD

ALT

VSI

AS

ALT

VSI

DP

ALTERNATE AIR SELECTOR VALVE ALTERNATE STATIC SOURCE BUTTON

PP

ALTERNATE AIR SELECTOR VALVE

PILOT'S

COPILOT'S P S2 S1 AS ALT VSI DP PP

ALTERNATE STATIC SOURCE BUTTON

= PITOT SOURCE = STATIC SOURCE (COPILOT'S) = STATIC SOURCE (PILOT'S) = AIRSPEED INDICATOR = ALTIMETER = VERTICAL SPEED INDICATOR = DIFFERENTIAL PRESSURE = PNEUMATIC PRESSURE

Figure 22-6.Pitot Pitotand andStatic Static System UB/C Figure 16-35 SystemSchematic Schematic — UB/C

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F.S. 42.75

S

NUMBER 1 ADS P P S

S

NUMBER 2 ADS P P S

F.S. 57.50

TP

TP

TP

TP

P ADC S F.S. 84.00 COPILOT'S PITOT-STATIC MAST PITOT P

PILOT'S PITOT-STATIC MAST PITOT P

S1

S2 S1

STATIC

STATIC

MANIFOLD

ASI

MANIFOLD

ALT

IVSI

PILOT ALTERNATE STATIC AIR SELECTOR VALVE

ASI

ALT

DPS

DPG

IVSI

ASI

COPILOT ALTERNATE STATIC AIR SELECTOR VALVE

S PILOT ALTERNATE STATIC PORT

LEGEND PILOT'S PITOT PILOT'S STATIC COPILOT'S PITOT COPILOT'S STATIC

S P S2 S1 ASI ALT IVSI DPG PPI DPS TP DV ADS FDR ADC

S2

= STATIC PRESSURE = PITOT PRESSURE = STATIC PRESSURE (COPILOT'S) = STATIC PRESSURE (PILOT'S) = AIR SPEED INDICATOR = ALTIMETER = INSTANT VERTICAL SPEED INDICATOR = DIFFERENTIAL PRESSURE GAGE = PNEUMATIC PRESSURE INDICATOR = DIFFERENTIAL PRESSURE INDICATOR = TEST PORT = DRAIN VALVE = AIR DATA SENSOR = FLIGHT DATA RECORDER = AIR DATA COMPUTER

S F.S. 120.00

COPILOT ALTERNATE STATIC PORT DV F.S. 468.25 DV FDR

F.S. 531.00

Figure 16-36 Pitot and Static System Sytematic — UE

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Figure 16-37 Airspeed Calibration - Alternate System Graph

A sample Airspeed Calibration - Alternate System graph from the Performance section of the POH/AFM is shown in Figure 16-37. When either system is switched to ALTERNATE, use this graph to determine the Indicated Airspeed required to maintain a desired Calibrated Airspeed. For example, to maintain a CAS of 125 knots with Approach Flaps and Gear Down, an IAS of about 130 knots is required.

maintain an Indicated Altitude of 15,290 feet MSL. The graph shows that in the alternate static mode, the actual airplane altitude is 285 feet lower than the altimeter indicates. In general, when using the alternate static air system for flight in the 1900 Airliner, indicated airspeeds will be higher than calibrated airspeed and indicated altitudes will be higher than actual for any given aircraft configuration.

A sample Altimeter Correction - Alternate System graph is shown in Figure 16-38. In this sample, to maintain an actual altitude of 15,000 feet MSL at 200 KIAS, it would be necessary to

When the alternate static air source is not needed, ensure that the STATIC AIRSOURCE valve switches are in the NORMAL position.

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Figure 16-38 Altimeter Correction - Alternate System Graph

The pitot-static masts can be heated electrically for flight in icing conditions. As a precautionary measure, it is customary to have the pitot heat on during high altitude flight when temperatures are below 5°C. It is not advisable to operate the pitot heat system on the ground except for testing or for short intervals to remove ice or snow from the mast. Operating the pitot heat on the ground for extended periods can damage the internal heating elements in the pitot-static masts. Two circuit breaker switches on the pilot’s right subpanel (ice protection control panel) control power to the heating elements in the pitot-static masts (Figure 16-39). Ice protection is also provided for the alternate static air system. The alternatic static heat switch, also a circuit breaker type switch, is located on the pilot’s subpanel.

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Figure 16-39 Ice Protection Control Panel

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CHAPTER 17 MISCELLANEOUS SYSTEMS CONTENTS Page INTRODUCTION ................................................................................................................ 17-1 GENERAL ............................................................................................................................ 17-1 OXYGEN SYSTEM ............................................................................................................. 17-3 Passenger Oxygen Deployment System........................................................................ 17-4 Flight Crew Oxygen Masks........................................................................................... 17-5 Oxygen Supply Cylinders ............................................................................................. 17-5 Oxygen Duration ........................................................................................................... 17-6 Servicing the Oxygen System ....................................................................................... 17-8 Physiological Training .................................................................................................. 17-9

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ILLUSTRATIONS Figure

Title

Page

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Oxygen System Schematic .................................................................................... 17-2

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Oxygen Cylinder Installation................................................................................. 17-3

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Oxygen System Controls ....................................................................................... 17-4

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Passenger Oxygen Mask........................................................................................ 17-5

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Crew Oxygen Mask ............................................................................................... 17-5

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Oxygen Fill Valve and Gage ................................................................................. 17-6

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Oxygen Pressure Gage........................................................................................... 17-6

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Percent of Usable Oxygen Capacity Graph ........................................................... 17-7

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Oxygen Duration Chart.......................................................................................... 17-7

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FAA Altitude Chamber........................................................................................ 17-10

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CHAPTER 17 MISCELLANEOUS SYSTEMS ;;;;;;;;;;;;;;;;;;;;;;; ;;;;;;;;;;;;;;;;;;;;;;; ;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;; ;;;;;;;;;;;; ;;;;;;;;;;;;;;;;;;;;;;; ;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;; ;;;;;;;;;;;; ;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;; ;;;;;;;;;;;;;;;;;;;;;;; ;;;;;;;;;;;; ;;;;;;;;;;;;;;;;;;;;;;; ;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;;; ;;;;;;;;;;;;

ANTI-ICE ON

12 16

8 4 0

RESET TEST

OIL

NO 1 FUEL TRANS

BLOWER OFF ENG 1 CHIP

XMSN OIL

NO 1 FUEL LOW

NO 1 FUEL FILTER

90° BOX OIL

BATT HOT

NO 1 BATT SYS

GEN 1 HOT

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INTRODUCTION Pilot and passenger comfort and safety are of prime importance in operating the 1900 Airliner.

Flight crewmembers must be prepared to use the oxygen system safely and effectively, when necessary, within the requirements of applicable Federal Aviation Regulations.

GENERAL This chapter deals with the oxygen system. It includes a general system description, operational considerations, and a discussion of emergency procedures. Use of the oxygen duration chart is demonstrated for a variety of flight situations.

FAR requirements for crew and passenger needs, and types and availability of oxygen masks, are discussed. Servicing procedures, referenced in the POH/AFM, are also included.

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Figure 17-1 Oxygen System Schematic

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OXYGEN SYSTEM The oxygen system (Figure 17-1) provides adequate oxygen flow for pressure altitudes up to 25,000 feet. Oxygen duration charts (POH/AFM Normal Procedures) for this system are based on a constant flow rate of 3.8 liters per minute (lpm) delivered to crew masks at all altitudes, and altitude-compensated flow delivered to passenger masks.

The oxygen system utilizes two interconnected 76.6-cubic-foot (77.9 on UE) cylinders mounted on either side of the nose under the floor of the nose compartment (Figure 17-2). Two cylinder pressure gages are mounted on the copilot’s right subpanel. Pressure from the cylinders is

PILOT TRAINING MANUAL

controlled by a regulator mounted on each cylinder, and by an altitude-compensated, constantflow regulator mounted on the cabin side of the forward pressure bulkhead.

Both cylinders are activated simultaneously by a push-pull handle (OXYGEN PULL ON) on the pilot’s subpanel, in the upper left corner (lower left corner on UE) (Figure 17-3). The handle controls pilot oxygen flow through the shutoff valve.

The OXYGEN PULL ON handle must be activated before the CABIN OXYGEN PULL ON control can supply oxygen to all 19 passenger outlets. The cabin oxygen control is mounted on the left lower portion of the pilot’s subpanel. This control has been relocated below the pilot’s control column on Series UE.

Figure 17-2 Oxygen Cylinder Installation

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Figure 17-3 Oxygen System Controls

PASSENGER OXYGEN DEPLOYMENT SYSTEM Oxygen is delivered to passenger masks through an altitude compensator, which varies the flow rate from a minimum of 0.1 lpm at 1,000 feet to 2.9 lpm at 25,000 feet.

When the cabin oxygen control knob is pulled out, a surge valve momentarily allows high pressure to reach passenger mask container assemblies located at elbow-level for seated passengers (Figure 17-4), and overhead for the

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center aft cabin passenger. On Series UE, the access panel is located at the outboard side of the cabin, above each window in the light assembly.

High pressure causes a plunger at the mask assembly to open the oxygen access panel, allowing flow to the outlet. In order to initiate oxygen flow to the masks, a lanyard valve pin must be pulled out. This is accomplished by pulling the lanyard tight. The pin must be reinserted or the control must be pushed back in to stop the flow of oxygen.

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Figure 17-4 Passenger Oxygen Mask

FLIGHT CREW OXYGEN MASKS Crew mask oxygen pressure is regulated through an orifice in the mask container assembly (Figure 17-5). Oxygen is available to the crew at a constant flow rate of 3.8 liters per minute, regardless of altitude (UA, UB, UC). A lanyard pin must be removed from the mask assembly to activate oxygen flow to the mask. Again, as with passenger masks, this is accomplished by pulling the lanyard tight. Reinsert the pin to stop the flow.

If either cylinder’s pressure exceeds 2,775 psi, its relief cap will blow out, relieving the pressure in both cylinders and providing a visual cue that overpressure has occurred. The cap must be replaced before the oxygen system can be reserviced. Two cylinder pressure gages, mounted on

On Series UE, optional diluter-demand crew masks are available to provide oxygen at approximately 2.5 lpm at 10,000 feet and 5.0 lpm at 25,000 feet.

OXYGEN SUPPLY CYLINDERS Both oxygen cylinders are serviced through a single filler valve covered by an access door, located below the left side of the nose baggage compartment (Figure 17-6). A gage adjacent to the filler valve is used for checking system pressure during filling. High-pressure overboard relief indicators are located on either side of the lower nose fuselage.

Figure 17-5 Crew Oxygen Mask

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the right side of the copilot’s instrument subpanel, indicate pressure in each of the cylinders.

A third gage, indicating oxygen pressure supplied to the cabin passenger masks, is on the upper right side of the copilot’s instrument panel (Figure 17-7).

OXYGEN DURATION

Figure 17-6 Oxygen Fill Valve and Gage

During preflight, the pilot should determine that available oxygen is sufficient to supply both crew and passengers during an unpressurized descent from 25,000 to 12,500 feet (10,000 on UE). Full system pressure is 1,850 ±50 psi, which is sufficient to supply pilot, copilot, and 19 passengers

Figure 17-7 Oxygen Pressure Gage

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for 60 minutes at 25,000 feet. Oxygen pressure can be determined by reading the cockpit gages.

To determine the percent of usable oxygen, note system pressure and refer to the Percent of Usable Oxygen Capacity Graph (Figure 17-8).

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The Oxygen Duration Chart (Figure 17-9) is based on 3.8 lpm delivered to the crew and altitude-compensated flow delivered to passengers. If the two gages in the cockpit indicate different pressures, use the average of the two for all calculations. To calculate total minutes of oxygen available, obtain the duration for a full bottle from Figure 17-9, considering the number of persons aboard and aircraft altitude, then multiply full bottle duration by percent of full bottle available.

Oxygen Duration Computation

Figure 17-8 Percent of Usable Oxygen Capacity Graph

To compute oxygen available at 25,000 feet cruise altitude, 32° F, for 18 passengers, the following example assumes that one pressure gage indicates 1,300 psi and the other indicates 1,600 psi. Average these two readings to 1,450 psi.

Enter the graph in Figure 17-8 at 1,450 psig, and read across to the 32° F intersection; then follow the graph down to read 82% of usable capacity.

Finally, multiply 82% by the duration in minutes obtained from the chart in Figure 17-9. With 18 passengers at 25,000 feet, 64 minutes oxygen would be available if both oxygen bottles were full; however, since the system is only partially full, it is necessary to compute 82% of the 64 minutes. The result is 52.48 minutes of oxygen available for pilot, copilot, and 18 passengers.

Figure 17-9 Oxygen Duration Chart FOR TRAINING PURPOSES ONLY

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Time of Useful Consciousness If decompression is experienced at high altitude, hypoxia can result from lack of oxygen. The body’s primary need is oxygen to keep the brain and other body tissues functioning properly.

Early symptoms of hypoxia, such as an increased sense of well-being, quickly give way to slow reactions, impaired thinking ability, unusual fatigue, and a dull headache. If pressurization is lost, the crew must act quickly to don oxygen masks and to provide oxygen to the passengers before hypoxia can cause impairment.

The following sets forth the average time of useful consciousness (time from onset of hypoxia until loss of effective performance) at various altitudes: ●

The following precautions should be observed when purging or servicing the oxygen system: ●

Avoid any operation that could create sparks. Keep all burning cigarettes or fire away from the vicinity of the airplane when oxygen is in use.

●

Inspect the filler connection for cleanliness before attaching it to the filler valve.

●

Make sure that hands, tools, and clothing are clean, particularly of grease or oil.

These contaminants are extremely dangerous in the vicinity of oxygen.

●

As a further precaution against fire, open and close all oxygen valves slowly during filling.

●

Use only aviators breathing oxygen (MIL-0-27210) for servicing the oxygen system. Do not use oxygen intended for medical purposes or for such industrial uses as welding. Such oxygen may contain excessive moisture that could freeze oxygen system lines and valves.

35,000 feet ...................... 1/2 to 1 minute

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30,000 feet ....................... 1 to 2 minutes

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28,000 feet ................. 2 1/2 to 3 minutes

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25,000 feet ....................... 3 to 5 minutes

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22,000 feet ..................... 5 to 10 minutes

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12,000 to 18,000 feet ............ 30 minutes or more

Individual reactions may differ from those shown above. The POH Emergency Descent procedure should be used to quickly descend to an altitude more compatible with human needs, and to reduce the potential for experiencing hypoxia.

SERVICING THE OXYGEN SYSTEM The oxygen system is serviced through a filler valve on the fuselage, below the left side of the nose baggage compartment (see Figure 17-6). A cylinder pressure gage at the filler port is used to monitor pressure indications during servicing.

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To recharge the oxygen system, remove the protective cap from the filler valve, and attach the hose from an oxygen recharging unit to the filler valve. Make sure that both the airplane oxygen system and the servicing equipment are properly grounded before servicing.

To prevent overheating, fill the oxygen system slowly by adjusting the recharging rate with the pressure regulating valve on the recharging unit.

All oxygen cylinders should be filled to 1,850 psi at 70° F. Pressure may be increased an additional 3.5 psi for each degree of increase in temperature; similarly, for each degree of drop in temperature, reduce the pressure for the cylinder by 3.5 psi. When the system is properly charged, disconnect the filler hose from the filler valve, and replace the protective cap.

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Oxygen Cylinder Retesting

Where Can You Get It?

Oxygen cylinders used in the airplane are of two types. Lightweight cylinders, stamped “3HT” on the side-plate, must be hydrostatically tested every three years and stamped with the retest date. This bottle has a service life of 4,380 pressurizations or 15 years, whichever occurs first, and then must be discarded. Regular-weight cylinders, stamped “3A”or “3AA,” must be hydrostatically tested every five years and stamped with the retest date. Service life on these cylinders is not limited.

A resident physiological training course at the FAA’s Aeronautical Center in Oklahoma City is devoted entirely to problems in civil aviation.

The National Aeronautics and Space Administration (NASA) in Houston, Texas, and many military installations conduct resident programs for non-government personnel.

Oxygen System Purging Purging may remove offensive odors from the oxygen system. The system should also be purged any time system pressure drops below 50 psi, or when the lines have been left open. To purge the system, connect a recharging unit, and allow oxygen to flow through lines and outlets until offensive odors have been carried away.

If any offensive odor still lingers, continue purging the system for an additional hour. If odors remain, the cylinder must be replaced. After purging, return the cylinder valve to its normal position, and service the system.

PHYSIOLOGICAL TRAINING What Is It? Physiological training is a program directed toward understanding and surviving in the flight environment. It covers problems occurring at both high and low altitudes and recommends procedures to prevent or minimize human-factor errors which occur in flight.

Who Needs It? The course primarily benefits pilots. It is also recommended for other aircrew personnel, air traffic controllers, aviation medical examiners, and other national aviation system personnel.

How Long Is the Course?

The course takes one full day.

What Is Contained in the Course?

Many topics are covered. They include the environment to which a person is exposed each time he flies, physiological functions of the body at ground level, and alteration of some of these functions by changes in the environment. The higher one flies the more critical are changes in the environment. The higher one flies the more critical is the need for extra oxygen. After the course, the trainee will understand why pilots cannot fly safely at altitudes in excess of 12,000 feet for prolonged time periods without some aid—either from supplemental oxygen or pressurized aircraft. Both oxygen equipment and pressurization are discussed.

When man is confronted with certain stressful situations, he tends to breathe too rapidly. This topic (hyperventilation) and methods of control are discussed. Ear pain on descent and other problems with body gases are described, and procedures to prevent or minimize gas problems are explained. Alcohol, tobacco, and drugs are also discussed as they apply to flying. Vertigo is discussed and demonstrated so that the trainee will understand why a non-current instrument pilot should never attempt to fly in clouds or other weather situations where visibility is reduced.

All resident courses include an altitude chamber flight (Figure 17-10), where the trainee may experience his individual symptoms of oxygen

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deficiency and rapid decompression. This simulated flight will demonstrate that: ●

Oxygen equipment will protect an individual from oxygen deficiency.

●

The pilot can experience and recognize symptoms identical to those in actual flight situations, and can learn to take necessary action to prevent loss of judgment and consciousness.

●

Decompression is not dangerous, provided necessary protective actions are taken afterwards.

What Are the Prerequisites for Training?

Trainees must have a valid FAA Medical Certificate. When training is scheduled at military facilities, a training fee of five dollars is required.

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Written parental consent is required for those under age 21.

How Do You Apply for Training?

All requests for the training course must be coordinated with: Physiological Operations and Training Section, AC-143 FAA Aeronautical Center P.O. Box 25082 Oklahoma City, Oklahoma 73125

How Can You Get Further Information?

Write to the Physiological Operations and Training Section at the above address, or phone (405) 954-4837.

Figure 17-10 FAA Altitude Chamber

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APPENDIX SYMBOLS, ABBREVIATIONS, AND TERMINOLOGY CONTENTS Page AIRSPEED........................................................................................................................ APP-1 METEOROLOGICAL...................................................................................................... APP-2 POWER............................................................................................................................. APP-2 CONTROL AND INSTRUMENT ................................................................................... APP-3 GRAPH AND TABULAR................................................................................................ APP-3 WEIGHT AND BALANCE ............................................................................................. APP-4 AVIONICS........................................................................................................................ APP-5

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APPENDIX SYMBOLS, ABBREVIATIONS, AND TERMINOLOGY AIRSPEED CAS—Calibrated airspeed is the indicated airspeed of an airplane corrected for position and instrument error. Calibrated airspeed is equal to true airspeed in standard atmosphere at sea level.

GS—Groundspeed is the speed of an airplane relative to the ground.

IAS—Indicated airspeed is the speed of an airplane as shown on the airspeed indicator when corrected for instrument error. IAS values published in this manual assume zero instrument error.

VFE—Maximum flap extended speed is the highest speed permissible with wing flaps in a prescribed extended position.

VLE—Maximum landing gear extended speed is the maximum speed at which an airplane can be safely flown with the landing gear extended.

VLO—Maximum landing gear operating speed is the maximum speed at which the landing gear can be safely extended or retracted.

TAS—True airspeed is the airspeed of an airplane relative to undisturbed air, which is the CAS corrected for altitude, temperature, and compressibility.

VMCA—Air minimum control speed is the minimum flight speed at which the airplane is directionally controllable, as determined in accordance with Federal Aviation Regulations.

The airplane certification conditions include: one engine becoming inoperative and windmilling, a 5° bank toward the operative engine, takeoff power on operative engine, landing gear up, flaps in takeoff position, and most rearward CG. For some conditions of weight and altitude, stall can be encountered at speeds above VMCA, as established by the certification procedure described above, in which event stall speed must be regarded as the limit of effective directional control.

V1—Takeoff decision speed.

VMCG—Ground minimum control speed.

V2—Takeoff safety speed.

VMO/MMO—Maximum operating limit speed is the speed limit that may not be deliberately exceeded in normal flight operation. V is expressed in knots and M in Mach number.

KCAS—Calibrated airspeed expressed in knots.

KIAS—Indicated airspeed expressed in knots.

M—Mach number is the ratio of true airspeed to the speed of sound.

VA—Maneuvering speed is the maximum speed at which application of full available aerodynamic control will not overstress the airplane.

VF—Design flap speed is the highest speed permissible at which wing flaps may be actuated.

VR—Rotation speed.

VS—Stalling speed or the minimum steady flight speed at which the airplane is controllable.

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V SO —Stalling speed or the minimum steady flight speed at which the airplane is controllable in the landing configuration.

VSSE—Intentional one-engine-inoperative speed is a speed above both V MCA and stall speed, selected to provide a margin of lateral and directional control when one engine is suddenly rendered inoperative. Intentional failing of one engine below this speed is not recommended.

VX—Best angle-of-climb speed is the airspeed which delivers the greatest gain of altitude in the shortest possible horizontal distance.

V Y —Best rate-of-climb speed is the airspeed which delivers the greatest gain in altitude in the shortest possible time.

METEOROLOGICAL Altimeter setting—Barometric pressure corrected to sea level.

Indicated pressure altitude—The number actually read from an altimeter when the barometric subscale has been set to 29.92 inches of mercury (1013.2 millibars).

IOAT—Indicated outside air temperature is the temperature value read from an indicator.

ISA—International standard atmosphere in which: ●

Air is a dry, perfect gas.

●

Temperature at sea level is 59º Fahrenheit (15º Celsius).

●

Pressure at sea level is 29.92 inches of mercury (1013.2 millibars).

●

Temperature gradient from sea level to the altitude at which the temperature is –69.7º F (–56.5º C), is –0.003566º F (–0.00198º C) per foot, and is zero above the altitude.

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PILOT TRAINING MANUAL

OAT—Outside air temperature is the free air static temperature, obtained either from the temp e r a t u r e i n d i c a t o r ( I OAT ) a d j u s t e d f o r compressibility effects or from ground meteorological sources.

Pressure altitude—Altitude measured from standard sea level pressure (29.92 inches Hg) by a pressure (barometric) altimeter. It is the indicated pressure altitude corrected for position and instrument error. In this manual, altimeter instrument errors are assumed to be zero. Position errors may be obtained from the altimeter correction graphs.

Station pressure—Actual atmospheric pressure at field elevation.

Temperature compressibility effects—An error in the indication of temperature caused by airflow over the temperature probe. The error varies, depending on altitude and airspeed.

Wind—The wind velocities recorded as variables on the charts of this manual are to be understood as the headwind or tailwind components of the reported winds.

POWER Beta range—The region of the power lever control which is aft of the idle stop and forward of reversing range, where blade pitch angle can be changed without a change of gas generator rpm.

Cruise climb—Cruise climb is the maximum power approved for normal climb. This power is torque or temperature (ITT) limited.

High idle—High idle is obtained by placing the condition lever in HIGH IDLE position. This limits the power operation to a minimum N1 of 70% (72% for UE).

Low idle—Low idle is obtained by placing the condition lever in LOW IDLE position. This limits the power operation to a minimum N1 of 58% (65% for UE).

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Maximum continuous power—Maximum continuous power is the highest power rating not limited by time. Use of this power setting is intended for emergency situations at the discretion of the pilot.

Power lever (gas generator N1 rpm)—The power lever serves to modulate engine power from full reverse thrust to takeoff. The flight idle position represents the lowest recommended level of power for flight operation.

Maximum cruise power—Maximum cruise power is the highest power rating for cruise and is not time limited.

Propeller control lever (NP rpm)—The propeller control is used to control the rpm setting of the propeller governor. Movement of the lever results in an increase or decrease in propeller rpm. Propeller feathering is the result of lever movement beyond the detents at the low rpm (high pitch) end of the lever travel.

Reverse—Reverse thrust is obtained by lifting the power levers and moving them aft of the ground fine range and moving them aft of the Beta range to the second detent.

SHP—Shaft horsepower.

Minimum takeoff power—Minimum takeoff power is the minimum power which must be available for takeoff without exceeding the engine limitations.

Takeoff power—Takeoff power is the maximum power rating and is limited to a maximum of 5 minutes operation. Use of this rating should be limited to normal takeoff operations and emergency situations.

CONTROL AND INSTRUMENT Condition lever (fuel shutoff lever)—The fuel shutoff lever actuates a valve in the fuel control unit which controls the flow of fuel at the fuel control outlet and regulates the idle range from low to high idle.

Propeller governor—The propeller governor senses changes in rpm and hydraulically changes propeller blade angle to compensate for the changes in rpm. Constant propeller rpm is thereby maintained at the selected rpm setting.

Propeller ground fine—The aft stop of the beta range. From here the power levers must be lifted to enter the reverse range. Propeller ground fine is used to provide deceleration on the ground during landing and accelerate-stop conditions by taking advantage of the maximum available propeller drag without creating negative thrust.

Torquemeter—The torquemeter system indicates the shaft output torque. Differential pressure from the mechanism within the reduction gearcase causes a bellows and servo system to indicate torque on a meter. Instrument readout is in foot-pounds.

GRAPH AND TABULAR

ITT (interstage turbine temperature)—Eight probes, wired in parallel, sense the temperature between the compressor and power turbines, and send the reading to the ITT indicator in degrees centigrade x 100.

Accelerate-go—Accelerate-go is the distance to accelerate to takeoff decision speed (V1), experience an engine failure, continue accelerating to liftoff, then climb and accelerate in order to achieve takeoff safety speed (V2) at 35 feet above the runway.

N 1 tachometer (gas generator rpm)—The N1 tachometer registers the rpm of the gas generator in percent, with 100% representing a gas generator speed of 37,468 rpm.

Accelerate-stop—Accelerate-stop is the distance to accelerate to takeoff decision speed (V1) and then bring the airplane to a stop.

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Route segment—Route segment is a part of a route. Each end of that part is identified by a:

AGL—Above ground level.

Best angle-of-climb—The best angle-of-climb delivers the greatest gain of altitude in the shortest possible horizontal distance with gear and flaps up.

Best rate-of-climb—The best rate-of-climb delivers the greatest gain of altitude in the shortest possible time with gear and flaps up.

Clearway—A clearway is an area beyond the airport runway not less than 500 feet wide, centrally located about the extended centerline of the runway, and under the control of the airport authorities. The clearway is expressed in terms of a clear plane, extending from the end of the runway with an upward slope not exceeding 1.25%, above which no object nor any terrain protrudes.

However, threshold lights may protrude above the plane if their height above the end of the runway is 26 inches or less, and if they are located to each side of the runway.

Climb gradient—Climb gradient is the ratio of the change in height during a portion of a climb to the horizontal distance traversed in the same time interval.

Demonstrated crosswind—Demonstrated crosswind is the maximum 90º crosswind component for which adequate control of the airplane during takeoff and landing was actually demonstrated during certification. This is not intended to limit operations in crosswind conditions.

MEA—Minimum enroute altitude.

Net gradient of climb—Net gradient of climb is the gradient of climb with the flaps in the takeoff position and the landing gear retracted. “Net” indicates that the actual gradients of climb have been reduced by 8% to allow for turbulence and pilot technique. The net gradient of climb graphs are constructed so that the value(s) obtained using the airport pressure altitude and outside air temperature will be the average gradient from 35 feet above the runway up to 1,500 feet above the runway.

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PILOT TRAINING MANUAL

●

Geographic location, or

●

Point at which a definite radio fix can be established

Takeoff flight path—Takeoff flight path is the minimum gradient of climb required to clear obstacles in excess of 35 feet, measured horizontally from reference zero, and vertically at the altitude above the runway. Reference zero is the point where the airplane has reached 35 feet above the runway, as determined from the accelerate-go graphs.

WEIGHT AND BALANCE Approved loading envelope—Those combinations of airplane weight and center of gravity which define the limits beyond which loading is not approved.

Arm—Arm is the distance from the center of gravity of an object to a line about which moments are to be computed.

Basic empty weight—Basic empty weight is the weight of an empty airplane, including full engine oil and unusable fuel. This equals empty weight plus the weight of unusable fuel, and the weight of all the engine oil required to fill the lines and tanks. Basic empty weight is the basic configuration from which loading data is determined.

Center of gravity—Center of gravity is the point at which the weight of an object may be considered concentrated for weight and balance purposes.

CG limits—CG limits are the extreme center-ofgravity locations within which the airplane must be operated at a given weight.

Datum—Datum is a vertical plane perpendicular to the airplane’s longitudinal axis from which fore and aft (usually aft) measurements are made for weight and balance purposes.

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Empty weight—Empty weight is the weight of an empty airplane before any oil or fuel has been added. This includes all permanently installed equipment, fixed ballast, full hydraulic fluid, full chemical toilet fluid, and all other operating fluids full, except that the engines, tanks, and lines do not contain any engine oil or fuel.

Engine oil—Engine oil indicates that portion of the engine oil which can be drained from the engine.

Jack point—Jack points are points on the airplane identified by the manufacturer as suitable for supporting the airplane for weighing or other purposes.

Landing weight—Landing weight is the weight of the airplane at landing touchdown.

Leveling points—Leveling points are those points which are used during the weighing process to level the airplane.

Maximum weight—Maximum weight is the greatest weight allowed by design, structural, performance, or other limitations.

Maximum zero fuel weight—The maximum weight allowable of the loaded aircraft (including payload) before adding fuel. Any weight above the value given must be loaded as fuel.

Moment—Moment is a measure of the rotational tendency of a weight, about a specified line, mathematically equal to the product of the weight and the arm.

Payload—Payload is the weight of occupants, cargo, and baggage.

PPH—Pounds per hour.

Ramp weight—Ramp weight is the airplane weight at engine start, assuming all loading is completed.

Station—Station is the longitudinal distance from some point to the zero datum or zero fuselage station.

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Takeoff weight—Takeoff weight is the weight of the airplane at liftoff from the runway.

Tare—Tare is the apparent weight which may be indicated by a scale before any load is applied.

Unusable fuel—Unusable fuel is the fuel remaining after consumption of usable fuel.

Usable fuel—Usable fuel is that portion of the total fuel which is available for consumption as determined in accordance with applicable regulatory standards.

Useful load—Useful load is the difference between the airplane ramp weight and the basic empty weight.

Zero fuel weight—Zero fuel weight is the airplane ramp weight minus the weight of fuel on board.

AVIONICS ADF mode—A mode of automatic direction finder operation allowing the ADF needle to point to the station.

NOTE In this mode of operation, on many receivers the audio fidelity is severely limited.

Air data computer—An electronic system primarily designed to gather information for an autopilot flight director system with outputs relating to pitot and static data. Possible information from this system includes: pressure altitude, indicated airspeed, total air temperature, static air temperature, and other information related to autopilot operation.

Altitude alert light—An amber light associated with an altitude alerter system. This light will be illuminated prior to intercepting a preselected altitude, or if for any reason the aircraft strays beyond a preset limit from the selected altitude once the aircraft has intercepted the altitude.

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Altitude preselector—An autopilot flight director subsystem that allows a pilot to preselect the altitude to which he desires to climb or descend. The controlling mechanism for an altitude preselect system is normally combined with the same device which controls the altitude alerter system.

Amplifier—A basic type of electronic device that seeks to make an electrical signal greater in strength. A public address system, for instance, is a type of amplifier. Amplifying devices are typically tubes or transistors.

Analog—A type of electronic circuitry that is characterized by smooth, continuous operation rather than discrete steps, as would be observed with digitally operated equipment.

Angle-of-attack (AOA) indicator—A supplemental flight instrumentation system that attempts to read out to the pilot the angle-of-attack or deck angle information. Several variations of this system are available.

Angular deviation—A means of showing displacement from a selected course either to or from a VOR station, TACAN, or NDB, showing displacement from the desired course in terms of angle. This is commonly used with the VOR system having a normal course width of 10° on each side of the course.

Annunciator—An indicator light with a message.

An annunciator makes an announcement as to the specific status of a system or subsystem.

ANT (antenna) mode—This mode of ADF operation allows improved audio fidelity in order to listen to the music or voice programs of an AM broadcast station. However, the ADF needle operation is defeated in this mode of operation.

Area navigation system (RNAV)—A system of direct point-to-point navigation having four further subclassifications: ●

APP-6

Course line computer—A computer util i z i n g i n f o r m a t i o n f r o m VO R a n d colocated DME stations, that allows the operator to change the location of the

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VOR station from its physical position to wherever the operator wants. ●

OMEGA/VLF definition.

●

Inertial navigation system—See related definition.

●

Loran system—Operationally similar to an OMEGA/VLF system.

system—See

related

Asymptotic—A design characteristic of an autopilot or flight director system. The function of this characteristic is to allow the autopilot to attempt to intercept a given course or altitude without overshooting. This is done by continuously reducing the intercept angle as the aircraft approaches the selected ground track or altitude.

Attenuation—The process of electrically reducing the size of a radio or audio signal (i.e., to turn down or make smaller.)

Attitude director indicator (ADI) (flight director indicator)—This instrument combines the basic functions of an attitude indicator with the steering commands received from the flight director system.

Attitude indicator (artificial horizon)—A gyroscopically controlled instrument used to display the aircraft’s pitch and roll attitude relative to the earth’s surface. The gyro used to display this information may be contained within the case of the displayed instrument, or it may receive its information from a remotely located attitude gyro.

Audio filters—An electronic means of removing a portion of the audio which the pilot does not desire to listen to. The pilot may choose to remove either the voice portion or the Morse code identifier of a VOR or an ADF system.

Audio selector switches—The system of switches which allows one or several audio syst e m s t o b e “ p i p e d - i n ” t o t h e s p e a ke r o r headphones of an aircraft.

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Audio system—The electronic system that serves as a switchboard and amplification system for the varied receivers that require the audio to be funneled to the speaker or headphones.

Autopilot/flight director modes: ●

●

Vertical modes—Vertical modes control changes in the pitch attitude of the aircraft using the elevator servo. Examples of vertical modes are: altitude hold, altitude preselect, indicated airspeed hold, and vertical speed hold.

Lateral modes—Lateral modes control autopilot operation by controlling the aileron and rudder servos. Examples of lateral modes are: heading hold, navigation modes and submodes (e.g., en route nav tracking, approach, backcourse, etc.).

Autopilot/flight director submode—Generally speaking, this concept represents two subclassifications of operation within a given mode.

Namely: ●

●

Arm—The process of activating a system or preparing it to operate at a future time.

For instance, if you push the Nav button to track a specific radial from a VOR station, but the CDI needle is displaced fullscale to the left or right at the moment you push the Nav button, then the autopilot flight director system will initially be activated in the nav-arm mode while the aircraft continues to intercept the selected radial.

Capture—A submode allowing the autopilot flight director system to track a specified lateral or vertical reference (e.g., altitude or glide slope as a vertical mode; VOR or localizer course as a lateral mode).

Avionics master circuit breaker—The circuit breaker that supplies power to the avionics master switch in Beechcraft factory-installed avionics packages on Baron, Bonanza, Duke, and King Air installations. This circuit breaker serves as a backup means of activating the avionics sys-

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tem should the avionics master switch fail for any reason.

Avionics master switch—A central on/off power switch for the entire avionics package in an aircraft. This switch conveniently allows the pilot to turn on the entire avionics package by turning on only one switch.

Beat frequency oscillator (BFO)—A device used on an ADF receiver that generates a tone allowing the pilot to identify the Morse code being transmitted by some nondirectional beacons.

NOTE This type of transmitter is not employed in the United States.

Carrier—That portion of the transmitted radio energy which “carries” the useful information (i.e., modulation).

Compass system slaving—The process of automatically aligning the directional gyro in a compass system with the earth’s magnetic field to display the aircraft’s magnetic heading. When the compass system is initially powered, slaving occurs at a fast rate to quickly align the compass system with magnetic north. Once the fast-slaving rate is accomplished, the system automatically goes into a slow-slaving rate for continuous operation. It will correct for precession errors of the compass system up to a maximum error of about 3° per minute.

Concentric—Two or more knobs mounted on one common system of shafts having the same axis. For example, most frequency selector knobs used in all avionic systems employ concentric knobs in the interest of conserving panel space.

Course deviation indicator (CDI)—An indicator used with a VOR/localizer receiver that shows only left/right deviation and to/from information.

This instrument has a knob called an OBS knob, meaning “omnibearing selector,” which allows the pilot to choose the course to or from a VOR station.

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Course knob—The name applied to the omni bearing selector on an HSI type of instrument.

The course knob is attached to the resolver and a course pointer on the HSI indicator.

Fast erect—A mode of operation whereby an attitude indicator may be quickly realigned with the earth’s horizon if for any reason the gyro has precessed or tumbled.

Course width—Displacement left or right of the desired course:

Flux valve (flux gate)—A component of a slaved compass system that senses the earth’s magnetic field and converts this information into an electrical signal representing magnetic north.

●

●

Angular—Degrees left or right of the desired course. Course width using the VOR system is 10° on each side of the desired course.

Linear—In the “en route mode” most course line computer RNAVs have a course width of ±5 nautical miles. In the approach mode most course line computer RNAVs have a course width of ±1 1/4 nautical miles on each side of the centerline.

Cross sidetone—Sending sidetone audio across the cockpit from one side to the other; for example, this allows the pilot to hear what the copilot is saying on the transmitter. Cross sidetone may be heard through either the phones or the speaker.

Digital—A type of electronic circuitry technology that operates in specific steps, as opposed to the smooth, sweeping type of operation employed in analog.

Double-cue flight director system (cross pointer)—A command presentation system using one vertical bar to indicate commanded roll-attitude instructions and one horizontal bar to indicate commanded pitch instructions.

Electronic flight instrument system (EFIS)—A type of flight instrumentation system employing cathode ray tubes (television screens) to display information.

Emergency/normal switch—In the event of the failure of the audio system, this switch (when placed in the emergency position) allows audio from the aircraft receivers to be “piped” directly to the headphones.

APP-8

Free operation—A mode of operation for a slaved compass system whereby the directional gyro is disconnected from the slaving system.

Normally this would be used when the slaving system fails or for operation in the polar regions where the earth’s magnetic field will not permit normal slaved operation. The concept here is that the directional gyro is free of its master, magnetic north.

Gain—The relative amount of amplification of a radio receiver. A gain control is commonly used on a radar indicator to control the relative amount of amplification of the received radar echo. This allows the pilot to optimize the information displayed, especially when the radar is used for terrain mapping purposes.

Go-around mode—An autopilot flight director mode intended to be used during a missed approach. This mode will command a pitch-up attitude appropriate for a climbout with an associated wings-level command. The autopilot may or may not remain engaged during the go-around mode, depending upon the type of autopilot installed in a specific aircraft.

Gyro erection—The process of an attitude gyro becoming aligned with the earth’s horizon or, viewed in another way, aligned with true vertical.

This happens automatically when the system first receives power.

Half bank—An autopilot mode of operation whereby the bank angle is limited during turns in such a way that the aircraft will only bank approximately half as much as normal. This is designed to give the passengers the perception of a smoother ride with no steep banks.

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Heading bug—An adjustable marker used on a heading indicator to direct an autopilot and/or flight director system according to the magnetic heading the pilot desires to fly. Also, this device may be used simply as a reminder to the pilot of what heading he is to fly when not using the autopilot flight director system.

Heading indicator (directional gyro)—A gyroscopically controlled instrument used to display an aircraft’s heading relative to magnetic north.

The compass card of this indicator may be receiving the information which it displays from a remotely located gyro and an associated slaved compass system.

Hertz—The unit of measure used to describe the number of cycles of alternating current per second.

Horizontal situation indicator (HSI)—This instrument, alternately called a CDI by some manufacturers, displays heading information from a compass system, left/right and to/from information from a VOR/localizer receiver, and deviation above and below a glide slope from a glide-slope receiver. The pilot’s workload is reduced by integrating these displays onto one indicator.

Inertial navigation system—This system allows direct point-to-point navigation via a great circle route. This system is completely self-sufficient, utilizing a group of gyros and accelerometers to sense movement along the earth’s surface.

Integrated autopilot/flight director system—A system utilizing both autopilot and flight direction information to respond to selected modes.

Interrogation—In the secondary surveillance radar system the ground-based radar unit is said to “interrogate” the transponders of all aircraft flying within reception range of that radar. Once a transponder has been interrogated, it should reply to the ground radar unit by sending a brief transmission of radio energy. For general aviation aircraft a transponder may be interrogated in both modes. A mode supplies azimuth and distance information, and altitude information is provided through mode C.

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Keying—The process of turning on the transmitter by means of the push-to-talk button located on the microphone or the control wheel.

Latitude—The angular displacement of a geographic location north or south of the equator.

This is normally expressed in terms of degrees, minutes, and tenths of minutes.

Linear deviation—A means of showing lateral displacement from the desired navigational course calibrated in miles. Linear deviation allows for parallel course boundaries whether far away from or near a station.

Longitude—The angular displacement of a geographic location east or west of the prime meridian located in Greenwich, England. This is normally expressed in terms of degrees, minutes, and tenths of minutes.

Magnetic bearing—The direction of a nondirectional beacon (NDB) or VOR station relative to magnetic north.

Meter movement—An application of an ammeter used in any instrumentation system to show deviation such as left/right, to/from, slaving indicator, etc.

Mode—One of several operating conditions of a system. For instance, most airborne weather radars have both weather mapping and terrain mapping modes of operation.

Mode A—That portion of the transponder reply which transmits azimuth and distance information for the radar controller.

Mode C—The portion of a transponder reply containing the pressure altitude of an aircraft as provided by an encoding altimeter.

Modulation—The addition of useful information to the carrier wave that is emitted from a transmitter; for example, talking into the microphone or the transmission of the Morse code identification from a VOR station.

Muting—The silencing of incoming receiver audio while one is transmitting.

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APP-9

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Nonintegrated AP/FD system—Two separate flight control systems, each using its own computer. Information coming from these two systems may or may not agree at any given time.

OMEGA/VLF system—A world-wide navigation system that allows direct great circle flight from one point to another. This system utilizes U.S. Navy VLF communication transmitters and the OMEGA system of navigation.

Parallax error—A problem that can cause inaccurate interpretations of an instrument reading. It is caused by the user’s viewing angle not being directly in line with the instrument.

Parallax error adjustment—An adjustment of some single cue flight director systems which allows the command bars to be adjusted up or down in order to “nestle” just above the aircraft symbol.

Parked or stowed ADF needle—The process of placing the ADF needle at the 3 o’clock relative bearing location to indicate that the ADF unit is in the antenna mode and that the needle is not operating.

Radio magnetic indicator (RMI)—The combined display of magnetic heading from a compass system and relative bearing to a nondirectional beacon or VOR, which results in displaying the product, called “magnetic bearing,” to or from the station.

Range filter—An audio filter designed to remove the Morse code identification from a radio transmission. Actually, range is something of a misnomer as pilots know it today—you should think of this as being an “ident” filter.

Relative bearing—The direction of a nondirectional beacon relative to the longitudinal axis of the aircraft.

Relay—An application of an electromagnet to perform switching duties. A relay may be used to switch large quantities of current. A multiple pole relay will allow a single pole switch to switch many circuits from a remote location.

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PILOT TRAINING MANUAL

Remote mounted avionics—Avionic equipment which is not fully self-contained and mounted on the instrument panel. Typically, the “black boxes” for these systems are located in the forward avionics compartment, forward of the front pressure bulkhead, or aft of the rear pressure bulkhead.

Resolver—The electronic device to which the course knob or OBS knob is attached. This device communicates the desired course, which the pilot selects, to the VOR receiver.

Servo system—Using an electric motor in any one of several applications to reduce pilot workload or allow automatic operation of some systems; for example: autopilot servos, electric elevator trim servos, servoed altimeters, compass systems, etc.

Sidetone—The ability to hear oneself talk while transmitting. The sidetone may be heard through either the headphones or the speaker. Additionally, sidetone may be considered as a means of verifying normal transmitter and receiver operation. If the receiver and transmitter are working properly, the sidetone will sound “normal.” If either the transmitter or receiver is malfunctioning, the sidetone will sound weak or garbled.

Single-cue flight director system (V-bar system)—A command display system using a pair of bars which work in unison to display the commanded attitude to the pilot.

Slant/range correction—A means of correcting for the inherent error in raw slant/range data which will result in a true lateral distance from the aircraft to the DME station. Many of the more sophisticated RNAV computers provide slant/range correction.

Slant/range distance—Conventional, uncorrected DME distance to the station.

Slaved compass system—A directional gyro system that is automatically synchronized to the magnetic heading of the aircraft. The concept of this system is that magnetic north is the master; therefore, the compass system is its slave.

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Slaved operation—The normal mode of a slaved compass system whereby the directional gyro automatically remains synchronized to magnetic north. This type of operation continually compensates for gyro precession and other compass system errors. The concept is that the compass system is a slave to magnetic north.

Slaving amplifier—An amplifier which takes the weak signal representing magnetic north, coming from the flux valve, and boosts that signal to a usable level in order to drive the directional gyro to the proper magnetic heading.

Slaving indicator—A meter used in some slaved compass systems that displays the difference between sensed magnetic heading and displayed magnetic heading. If the needle on this indicator is centered, there is no error between sensed and indicated magnetic heading. If the needle is off to the left or right, a small amount of error is indicated. Normal operation of the compass system causes the needle to sway to the left and right because of gyro precession and other factors.

Soft ride—A mode for an autopilot whereby the responsiveness of the autopilot to rough air is altered in such a way that the ride is perceived to be much smoother than it is.

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Voice terrain advisories—Voice callouts of pertinent altitude-above-ground information. The information announced will be determined by the type of system installed.

Waypoint—The geographic location of navigational fix used in area navigation. This may be used in either a VLF/omega system or a VOR/DME system utilizing a course line computer.

Waypoint address—The radial and distance of a waypoint from a VORTAC. Waypoint coordinates—The latitude and longitude of the waypoint used with a VLF/omega system.

Yaw damper—A system connected to the rudder servo that seeks to dampen or reduce oscillations of the aircraft about the yaw axis. The yaw damper system significantly reduces the level of motion sickness experienced by passengers. This system should be engaged soon after takeoff and under normal operations should remain engaged until just prior to landing.

Squelch—A silencing circuit employed in communication receivers that allows undesirable background noise to be omitted. Only a strong incoming signal from a transmitter will be heard.

Transponder code—A specific four-digit code that may be selected by the pilot on his transponder to identify his specific aircraft.

NOTE A common misconception is that this control changes the transponder reply frequency. The transponder always operates on the same frequency.

Voice filter—An audio filter designed to remove t h e v o i c e p o r t i o n o f a r e c e iv e d r a d i o transmission.

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APP-11

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ANNUNCIATORS The Annunciators section presents a color representation of all the annunciator lights in the airplane.

Please unfold to the right and leave open for ready reference as the annunciators are cited in the text.

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ANN-1

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L FUEL PRESS

CABIN ALTITUDE

BAGGAGE DOOR

INVERTER

R FUEL PRESS

L OIL PRESS

L ENVIR FAIL

FWD CABIN DOOR

R ENVIR FAIL

R OIL PRESS

A/P DISC

R BL AIR FAIL

AFT CABIN DOOR L BL AIR FAIL

A/P TRIM FAIL

SERIES UA, UB, UC

L DC GEN

L FUEL QTY

L FW VALVE

L FUEL FEED

L GEN TIE OPEN

L ENG ICE FAIL

L BK DI OVHT

L AUTOFEATHER

L CHIP DETECT

L IGNITION ON L ENG ANTI-ICE

L BK DEICE ON

L ENVIR OFF

FUEL TRANSFER

BATTERY CHARGE BATT TIE OPEN

R FUEL QTY

R DC GEN

R GEN TIE OPEN

R FUEL FEED

R FW VALVE

HYD FLUID LOW

MAN STEER FAIL

R BK DI OVHT

R ENG ICE FAIL

ANTI SKID FAIL

PWR STEER FAIL

R CHIP DETECT

R AUTOFEATHER

TAXI LIGHT

EXTERNAL POWER

ELEC TRIM OFF

MAN TIES CLOSE

R IGNITION ON R BK DEICE ON

R ENG ANTI-ICE

AIR COND N1 LOW

R ENVIR OFF

R FUEL QTY

R DC GEN

SERIES UA, UB BATTERY CHARGE BATT TIE OPEN

L DC GEN

L FUEL QTY

L FW VALVE

L FUEL FEED

L GEN TIE OPEN

R GEN TIE OPEN

R FUEL FEED

R FW VALVE

L ENG ICE FAIL

L BK DI OVHT

HYD FLUID LOW

MAN STEER FAIL

R BK DI OVHT

R ENG ICE FAIL

ANTI SKID FAIL

PWR STEER FAIL ANN PWR SOURCE

L NO FUEL XFR

R NO FUEL XFR PWR STEER ENGA

L AUTOFEATHER

L IGNITION ON

TAXI LIGHT

EXTERNAL POWER

R IGNITION ON

R AUTOFEATHER

L ENG ANTI-ICE

L BK DEICE ON

ELEC TRIM OFF

MAN TIES CLOSE

R BK DEICE ON

R ENG ANTI-ICE

L ENVIR OFF

FUEL TRANSFER AIR COND N1 LOW

SERIES UC

ANN-2

FOR TRAINING PURPOSES ONLY

R ENVIR OFF

FlightSafety international

BEECH 1900 AIRLINER

EXTINGUISHER PUSH

MASTER WARNING

MASTER CAUTION

PRESS TO RESET

PRESS TO RESET

HDG ALT

NAV

PILOT TRAINING MANUAL

ARM NAV APPR ARM APPR DSC CLIMB ALT ARM GS ARM GS GA

OK

DISCH

EFIS DRIVE XFR ALT ALERT

EXTINGUISHER PUSH DISCH

CABIN DIFF HI

R FUEL PRES LO

CABIN DOOR

R ENVIR FAIL

R OIL PRES LO

L AC BUS

CARGO DOOR

R AC BUS

A/P TRIM FAIL

ARM EMER LITES

A/P FAIL

L FUEL PRES LO

CABIN ALT HI

L OIL PRES LO

L ENVIR FAIL

OK

COMPARE PUSH TO RESET

GPS

L BL AIR FAIL

B/C

CLIMB

ALT

ALT SEL

VS

IAS

DSC

2

8

PUSH TO RESET

˚C X 100

7

˚C X 100

7

OBS

AFX

10

20

4 BARO

30

1

2

W

12

24

E

4

2

-11

20 0 -20 ˚C

50 0 PSI

OIL 200 140 150 100 100 60 20 0 -20 ˚C

30 27 24

21

18

21

ANT OFF

ADF TONE

STBY OFF

STO

V

D

0

33

CLR

CRSR

ENT

NAV D/T ACTV REF CTR

APT VOR NDB INT SUPL

AP/L YD

XFR

2

MEM

DIM OFF

RNG

8

ON ALT

60

7

3 HP

IN.HG 45

4

6

5

5

11 12 1 2 10 9 3 4 8 7 6 5 35

DRIVE XFR ALT ALERT

15

25

30

3

2

1 .5 UP

VERTICAL SPEED

3

MASTER WARNING

PRESS TO RESET

PRESS TO RESET

GYRO – SLAVING +

COMPARE

HDG ALT

PUSH TO RESET

NAV

ARM NAV APPR ARM APPR DSC CLIMB ALT ARM GS ARM GS GA

0 DN .5

ET

CRS

24.4 NM

00:00 12

V O R 1

15

SLEW

INSTANTANEOUS FPM X 1000

1

MODE

3

GPS

DG

2

A 30 40 50

ON

NORM

EADI/EHSI DIM

12.5

20K

20

ELAPSED TIME

15K

60

25K

N ALTITU BI

GPS APR

70

CRS TO

130

VG FAST ERECT

OXYGEN OUTLET PRESSURE PSIG

ADF TONE STO

V TEST

MASTER CAUTION

EFIS

PUSH HARD

ADF ANT OFF

IDENT

2 55

10

MEM

TEST

ACT

1

VG FAST ERECT

Collins

1

ALT 0900 TEST

50 0

BELOW G/S P/CANCEL

9

GS

ACT

ARM

ACTV

HDG

NAV

APPR

B/C

CLIMB

ALT

ALT SEL

VS

IAS

DSC

TEST

GPS CRS OBS

GPWS

BELOW G/S

P/TEST

P/CANCEL

LEG

ON

TEST

PSI

WSHLD ANTI-ICE NORMAL O F F PILOT PILOT AIR

DEFROST AIR

MAIN

PULL ON

HI COPILOT SURFACE DEICE SINGLE

LDG GR CONTROL

NOSE

L ENG ICE FAIL

3

R DC GEN

L

R

H

H

ARM OFF TEST

LO PITCH

L BK DI OVHT

HYD FLUID LOW ANTI SKID FAIL ANNPWRSOURCE R BK DI OVHT

L FIRE LOOP

L PITOT HEAT

XFR VALVE FAIL PWR STEER FAIL MAN STEER FAIL

L NO AUX XFR

AUTOFEATHER OFF

PITCH TRIM OFF

INBD WG DEICE

YD/RB FAIL

R FW VALVE

R PITOT HEAT

R FIRE LOOP

AFX DISABLE

R NO AUX XFR

UP DN

FUEL VENT LEFT RIGHT DOWN LOCK REL

HDL LT TEST

WARN HORN

RUD BOOST OFF OUTBD WG DEICE

TAIL DEICE

L AUTO FEATHER L IGNITION ON

PWRSTEERENGA

R IGNITION ON R AUTO FEATHER

L ENG ANTI-ICE L BK DEICE ON

MAN TIES CLOSE

R BK DEICE ON R ENG ANTI-ICE

FUEL TRANSFER

0 STALL WARN

LEFT

PITOT RIGHT

ALT STATIC

BRAKE DEICE

LANDING GEAR

UP 0˚

2

CABIN OXYGEN PULL ON

FLAPS

RELAY

2

1 .5

B IN CA ALT

4

35 30

CABIN CLIMB FT PER MIN

17˚

0

6

25

0 0

0

ON

3

LDG GR CONTROL

NOSE

VENT AIR - PULL ON

2

10

A F T

COPILOT AIR

PULL DECR

PULL ON

1000

1000 1500

500

FLIGHT 0 0 0 5

1500

500

PSIG 1/10

4 5

0

PSIG

2000

0

2000

OXYGEN

OXYGEN

CYLINDER PRESSURE

CYLINDER PRESSURE

L DC GEN

L FUEL QTY

STALL HEAT

BATTERY CHARGE PROP GND SOL

R FUEL QTY

R DC GEN

AUX ARM

AUX ON

L FW VALVE

L COL TANK LOW L GEN TIE OPEN

BATT TIE OPEN

R GEN TIE OPEN R COL TANK LOW

ANTI SKID FAIL

ANN PWR SOURCE

R BK DI OVHT

R ENG ICE FAIL

PWR STEER FAIL MAN STEER FAIL

R PITOT HEAT

R FIRE LOOP

AFX DISABLE

R NO AUX XFR

R FW VALVE

AUX TEST

GND FINE

ALT ALERT

LIFT

FEATHER

U P

FUEL CUTOFF FRICTION LOCK

R E V E R S E

REVERSE ONLY WITH ENGINES RUNNING

R

CABIN AIR

T

PUSH TO

LIFT

CAUTION

0

OFF

STALL WARN TEST

ENG FIRE TEST EXT TEST A O F F TEST B LEFT RIGHT LOOP O F F AMP DETECT

IDLE

C O N D I T I O N

P R O P TAXI

T R I M

DN

L

P O W E R

MAN COOL

HIGH RPM

INCR

E L E V A T O R

F W D

AUTO P TEST

DECR INCR

120 FURN ON

HIGH

VG FAST ERECT

MAN

20 PNEUMATIC PRESSURE

15

TAKEOFF LANDING AND REVERSE

D N

˚F CABIN TEMP

5

1 3

20

4

2

0

INCR

OVERSPEED WARNING TEST

0 FT

6 4

1

80

100

5

.5

35˚ DOWN

PSI

6

INCHES OF MERCURY

HOURS

R ENVIR OFF

EXTERNALPOWER

TAXI LIGHT

HI O F F AUTO

OFF AUTO TEMP

CABIN TEMP 40

SILENCE

RDR PWR ON

BLOWERS MAN TEMP

4 5

VACUUM

ENVIRONMENTAL BLEED AIR VALVES OPEN E NO VF IF R INSTR & ENVR OFF LEFT RIGHT

R ENG ICE FAIL

PULL ON

AUTOFEATHER

STARTER ONLY

R FUEL QTY

BATTERYCHARGE

L COL TANK LOW L GEN TIE OPEN BATT TIE OPEN R GEN TIE OPEN R COL TANK LOW

L FW VALVE

PROP AUTO MANUAL

MANUAL

PROP TEST OVERSPEED O F F

STALL HEAT

L FUEL QTY

L DC GEN ICE PROTECTION

ACTUATORS STANDBY

OPEN

RIGHT

OFF OXYGEN PULL ON

ALT

GPWS P/TEST

0 LOC2

DH200 ACT

PUSH ON

BRT

ATC

UP 0˚ F L A 17˚ P DOWN 35˚

FRICTION LOCK

L ENG ICE FAIL

L BK DI OVHT

HYD FLUID LOW

L FIRE LOOP

L PITOT HEAT

XFR VALVE FAIL

L NO AUX XFR

AUTOFTHER OFF

PITCH TRIM OFF

INBD WG DEICE

YD/RB FAIL

TAIL DEICE

L AUTOFEATHER

L IGNITION ON

PWR STEER ENGA

R IGNITION ON

R AUTOFEATHER

L ENG ANTI-ICE

L BK DEICE ON

MAN TIES CLOSE

R BK DEICE ON

R ENG ANTI-ICE

L ENVIR OFF

RDR PWR ON

TAXI LIGHT

EXTERNAL POWER

R ENVIR OFF

CANCEL

AILERON TRIM LEFT

UP

H

H

RIGHT

LEFT

0

3

RUDDER TRIM RIGHT

3

5

0

5

3

3

5

DN

EFIS TEST

EFIS POWER COPLT PILOT EADI DSPL EHSI PRCSR

PILOT EADI EHSI

ADC TEST

AP/L

COPLT DSPL PRCSR

OFF

ET

PRE

ACT

TTG

ALL ON NO. 2 DH SET

WX ARC HSI ARC MAP MAP

S/S

ET

TST

PRE

G

RN

P US H

HDG

CRS

P US H

HDG

CRS

DIR C E

CRS SEL

SY N C

MODE TEST STBY OFF

R

XFR

XFR

DIR C E

CRS SEL

L

RANGE

NORM WX MAP

TGT

HLD

50

STB

25 10

100 200 300

RUD BOOST OFF OUTBD WG DEICE

+10

DN SR YAW ENG

ACT

P US H

P US H

SY N C

GSP

NORM

G

TTG

TST

COPLT CMPST

NORM COPLT OFF ALL ON NO. 1

HDG

ALL ON NO. 2 TIMER SET

DH SET

WX ARC HSI ARC MAP MAP

DR XFR PLT NORM

NORM

RN

GSP

S/S

DSP PLT

COPLT OFF ALL ON NO. 1 ATT

OFF NAV DATA

TIMER SET

NAV DATA

PLT CMPST

AP/R

PLT COPLT OFF

OFF

T

HDL LT TEST

MIN

AP ENG

+5

MAX 0

1/20

GAIN

UP

TILT

-5 -10 TAKEOFF LANDING TURBULENCE OPTION AUTOBRIEF

PARKING BRAKE

ELEV TRIM

PARK

POWER ON

YAW DAMP

OFF

TAXI

OFF

PUSH FWD ON/OFF

O F F YAW CONTROL TEST

CABIN PRESSURE DUMP P R E S S TEST

2

3

WARNING

4

6

1 14

RATE INC

DEPRESSURIZE CABIN BEFORE LANDING

2

1

CABIN R CONTROLLE

8 10

ANTI SKID

CANCEL

FT T AL 00 0 X1

POWER STEERING RUDDER BOOST

SEND

CAB IN

DOWN LOCK REL

ACF T

GPWS FLAP OVRD

MSG

+02

5

GPWS INOP

CALC STAT SETUP OTHER

TA1.1 -06 TA2.6 +05

ADF

TEST

GPS

5 nm

0

G/S CANCLD

ON

DSC

00 MEM

OIL 200 140 150 100 100 60

IAS

10

30 TEST

VS

0

TEST

Collins

BFG XFR

MEM

ALT SEL

V TEST

Collins

1

L ENVIR OFF

LEFT

V

DIS 3 4 . 5 N M D AT E / T I M E • • • • • • • • • • 11 SEP 97 OBS IN –––˚ 02:04:00UTC O U T 3 1 5 ˚ A LT 0 1 5 0 0 F T RMI 130˚ BARO:30.10" O N A P P ROV E ?

ANNUN ENR–LEG CRSR

CRSR NAV FPL MODE TRIP

0

PPH X 100

3

ALT

LEG

STO

BENDIX/KING KLN 90B TSO

RIGHT

RIGHT

R GEN GEN TIES MAN CLOSE N O R M

IGNITION AND ENGINE START

OBS

DE

TEST

4

1

CLIMB

10

ON OFF

L GEN

2

B/C

COM

SQ ON OFF OFF

IDENT

TEST

ACT

APPR

OFF LEFT ENG ANTI-ICE LEFT

OFF

BUS SENSE RESET

XFR

MEM MEM

NAV ON HLD

3, 0 0 0

NAV

300 4060 280 KNOTS 80 260 100 240 120 220 140 200 180160

Collins

MEM MEM

OFF

ALTITUDE SET ALT ALERT

RESET

ARM NAV APPR ARM APPR DSC CLIMB ALT ARM GS ARM GS GA

HDG GPS CRS

ACTV

5 0

PPH X 100

3

ENG AUTO IGNITION ARM GEN RESET

MASTER SWITCH

OFF

VE

PUSH TO

VG FAST ERECT

AC BUS

AUX ON

XFR MB

DI

8 7 6 FUEL FLOW

8 7 6 FUEL FLOW GPWS FLAP OVRD

GPWS INOP

CRS

OFF

BATT

20 40

60

AUX ARM

Collins

CLI

CANCEL

PROP SYN ON

AVIONICS

0

% RPM 80

5 G/S CANCLD

TURBINE

TEST

NAV IDENT

V TEST

0.0

110 100

3

6

S

15

NA V

FPM X 1000

ON

ON RIGHT O F F TRANSFER

20 40

SILENCE

ARM

PA

AUX TEST

XFR OFF

INSTANTANEOUS

000

LEFT

0

% RPM 60

GPS APR

COMM 2 COMM 1

OFF

MEM

ON HLD OFF

STO

NAV

PH

STBY HRZN PWR ON

HORN

MEM

COM

SQ ON OFF

COPILOT AUDIO OFF

OFF

Collins

XFR

MEM MEM

HDG ALT

PUSH TO RESET

ADF

V NA

3

V1

TURBINE

OVERRIDE

2

12

12

V NA

33

ELAPSED TIME

AUX ON

L VO

DME 1

VOLUME

AUDIO SPKR

L VO

MKR BCN 1 & 2 HI L VO LO

XPNDR NORMAL

MKR BCN 1 2

2

9

9 00:00 N

0.0

110 100

3

GND COMM PWR

DIM

OFF

ENCD ALTM 1 ALTM 2

0 20 PROP 5 19 18 RPM X 100 10 17 16 15 1413

80

DN .5

2

NAV 1

GPS

BOTH RANGE PA

PUSH ON/OFF

6

6

20.0 NM L O C 1

EADI/EHSI DIM

SILENCE

VERTICAL SPEED

0

EFIS AUX POWER HORN

OFF

25

OFF DME

V

2

1 .5 UP ET

CRS

AUX TEST

EXT PWR

20

OFF

15

50

0 20 PROP 5 19 18 RPM X 100 10 17 16 15 1413

3

IN HG

30.03

5

3

OFF

ON

30

AUDIO SPKR

PH

1

15

35

VOICE

BOTH RANGE ANN HOT PUSH BRT INTPH

40

40

DG

AUX ARM

25

10

DH200 0

MODE

TEST

X 100

20

PA

5 TORQUE

40

Collins

MB

1017

6

0

33

ON

7

10

GYRO – SLAVING +

SLEW

30

ALT

10

VOICE

ADF

PRESS TO

COMPARE

10

10

15

25

30

15

35

0 9 500 1 1,400 2 8

VOLUME

COMM 1

0

50 45

5

L

5

35

2

KR

0

TORQUE

COMM 1 2

DME 1

VO

60 55

AP/L YD

MKR BCN 1 2

2

COMM 2

LEG

50

ALT

NAV 1

PILOT AUDIO OFF AFX

GS

COMM 1 2

4

6 5

GPS CRS

ARM ACTV

LOC1

2

8

4

6 5

TEST

MASTER WARNING

RESET

EFIS DRIVE XFR ALT ALERT

50

APPR

RADIO CALL

N669CM

SP

NAV

FIRE PULL

PULL TO CLOSE FIREWALL FUEL VALVE FIRE EXT ARM

12

9

X 100

11 12 1 2 10 9 3 4 8 7 6 5

OK

DISCH

ITT

45

45

R BL AIR FAIL

NA V

ARM NAV APPR ARM APPR DSC CLIMB ALT ARM GS ARM GS GA

START

12

ITT

40

PUSH HARD

A/P FAIL

SERIES UE EXTINGUISHER PUSH

MASTER CAUTION

HDG

300 4060 280 KNOTS 80 260 100 240 120 220 140 200 180160

ARM EMER LITES

R BL AIR FAIL

PRESS TO TEST

PRESS TO

START

COMPARE

GPS APR

FLT RCDR OFF

A/P TRIM FAIL

L BL AIR FAIL

KR

BELOW G/S P/CANCEL

R OIL PRES LO

R AC BUS

SP

GPWS P/TEST

R ENVIR FAIL

CARGO DOOR

FIRE PULL DRIVE XFR ALT ALERT

GPS

AUX TEST

R FUEL PRES LO

CABIN DOOR

L AC BUS

27

NAV

CABIN DIFF HI

L ENVIR FAIL

RESET

9 HDG ALT

CABIN ALT HI

L OIL PRES LO

24

PRESS TO

EFIS

AUX ON

L FUEL PRES LO

OK

DISCH

PRESS

LEG

DSC

MASTER CAUTION

MASTER G WARNIN TO RESET

OBS

FIRE PULL

20

IAS

10K

VS

CA

ALT SEL

EXTINGUISHER PUSH

AUX ARM

GPS CRS

ARM ACTV

ALT

FIRE PULL

TEST

CLIMB

T

P/CANCEL

B/C

21

P/TEST

APPR

18

GPWS

NAV

GPS APR

FLT RCDR OFF

15

HDG BELOW G/S

TEST

ERASE

OFF

HEADSET 600 OHMS

COCKPIT VOICE RECORDER

FUEL TRANSFER

SERIES UE

Figure ANN-1 Annunciators

FOR TRAINING PURPOSES ONLY

ANN-3