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INSD-2650-29

RTG POWER FOR VIKING LANDER CAPSULE

NUCLEAR SYSTEMS DIVISION A TELEDYNE COMPANY

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DISCLAIMER This report was prepared as an account of work sponsored by an agency of the United States Government. Neither the United States Government nor any agency Thereof, nor any of their employees, makes any warranty, express or implied, or assumes any legal liability or responsibility for the accuracy, completeness, or usefulness of any information, apparatus, product, or process disclosed, or represents that its use would not infringe privately owned rights. Reference herein to any specific commercial product, process, or service by trade name, trademark, manufacturer, or otherwise does not necessarily constitute or imply its endorsement, recommendation, or favoring by the United States Government or any agency thereof. The views and opinions of authors expressed herein do not necessarily state or reflect those of the United States Government or any agency thereof.


DISCLAIMER Portions of this document may be illegible in electronic image products. Images are produced from the best available original document.


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TAGS-85/2N RTG POWER FOR V I K I N G LANDER CAPSULE

INSD- 2650-29

AUGUST

1969

This document c Atomic Energy Act of 19 of its contents in an prohibited.

NUCLEAR SYSTEMS DIVISION

A TELEDYNE COMPANY Baltimore, Maryland

ta a s defined in the the disclosure


LEGAL NQTICE This report was prepared as an account of Government sponsored work. Neither the United States, nor the Commission, nor any person acting on behalf of the Commission:

A.

Makes any warranty or representation, expressed or implied,

with respect to the accuracy, completeness, or usefulness of the information contained in this report, or that the use of any information, apparatus, method, or process disclosed in this report may not infringe privately owned rights; or

a I

B.

Assumes any liabilities with respect to 'iI-8~

3f, or for

damages resulting from the use of any information, apparatus, method, or process disclosed in this report.

As used in the above, "person acting on behalf of the Commission"

includes any employee or contractor of the Commission, or employee of such contractor, to the extent that such employee or contractor of the Commission, or employee of such contractor prepares, disseminates, or provides access to, any information pursuant to his employment or contract with the Commission, or employment with such contractor.

INSD-2650-29 ii


i

c E L mmwom This report presents the results of studies performed by Isotopes, Inc., Nuclear Systems Division, to optimize and baseline a TAGS-85/2N

RTG for the Viking Lander Capsule prime electrical power source.

These

studies generally encompassed identifying the Viking RTG mission profile and design requirements, and establishing a baseline RTG design consistent with these

-

in this report were per.__"

formed under AEC Contrac

-- - - ~ ~

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INSD-2650-29 iii


CONTENTS Page

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Legal Notice Foreword Contents Summary

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....... ....... . ....... 11. Viking RTG Options . . . . . . . . . , . . . . . . A. TAGS-85/2N RTG . . . . . . . . . . . . . . . . 1. Design Description . . . . . . . . . . . . 2. Predicted Performance . . . . . , . . . . . B. 3P/2NRTG . . . . . . . . . . . . . . . . . . 1. Design Description . . . . . . . . . . . . 2. Predicted Performance . , . . . . , . . . . 111. Performance Studies . . . . . . . . . . . . . . . . A. RTG Optimization . . . . . . . . . . . . . . . 1. Radiator Parameters. . . , . . , . . . . . 2. TAGS-85/2N RTG . . . . . . . . . . . . . . 3. 3P/2NRTG . . . . . . . . . . . . . . . . B. Thermoelectric Supporting Data , . . . . . . . 1. Property Data. . . . . . . . . . . . . . . 2. TAGS Performance Data . . . . . . , , . . 3 , 3P/2N Performance Data . . . . . . . . . . . 4. Mechanical Properties . . . . . , , . . , . 5. Sublimation . . . . . . . . , . . . . . . A. SNAP 19 Nimbus RTG Design . . . B. Guidelines for Viking RTG S t u d y . C. Scope of Viking RTG Study . . .

INSD-2650- 29 iv

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iii

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iv

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vii

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1-1

I-5

. . . 1-6 . . . 11-1 . . . 11-1 . . . 11-1 . . . 11-10 . . . 11-28 . , . 11-28 . . . 11-34 . . . 111-1 . . . 111-1 . . . 111-1 . . . 111-3 . . . 111-9 . . .III-12 . . . 111-12 . . . 111-13 . . , 111-54 . . . 111-64 . . . 111-75


COrJTENTS (continued) Page C.

............... Analytical Model . . . . . . . . . . . . . . . Results . . . . . . . . . . . . . . . . . . . Viton 0-Ring Summary - . - -

Fill Gas Management 1. 2.

3.

..

.

... ............ General . B. Intact Impact Heat Source (IIHS) Design 1. Capsule Design. . . . . . 2. Heat Shield Design. . 3. Capsule Suitability to Solid Fuel Forms * C. Steady State Temperature Profiles * D. Launch Pad Abort Study . . 1. Environment Characterization. 2. Heat Source Response . 3. Composite Launch Pad Abort Probabilities . E. Flight Abort Study. 1. Flight Profile Analysis . 2. Re-entry Trajectory Evaluation 3. Heat Source Thermal Response . 4. Heat Shield Thermal Stresses .. . 5. Flight Abort Event Probabilities F. Impact and Post-Impact Study . 1. Impact Containment . 2. Post-Impact Containment . G. Design Considerations . . . . ... . . A.

V.

*

.............

Mechanical 1. Installation 2. Dynamic Considerations B. Thermal 1. RTG Temperature Limitations 2. Viking Study Results C. Electrical. 1. Load Condition 2. Ehvironmental Effects 3. Related Equipment 4. Instrumentation ,

........... ...............

..

111-78 111-78

111-83 111-104

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IV-1 IV-2 . . . . . . . . . . . . IV- - IV-144 . - 14 - - - IVIV-16 IV-20 - - . . . . 1v-21 - IV-29 . 1v-44 - - . . - - - . . IV-46 - - - - IV-46 - . .. IV-51 - - - - - IV-58 - 66 - - . IVIV-73 - - - - 1v-78 - - - - - - IV-78 - - . . IV-79

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RTG/VLC Integration Considerations A.

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INSD- 26 50 - 29 V

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IV-86

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V-4

v-6 V-6

v-a

v-11 v-11 V-14 v-20 V-28


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CONTENTS (continued) Page

.

.

.

D. Nuclear Radiation . .... .. ..... 1. Unshielded Fluxes . . 2. Shield Attenuation and Weight Optimization 3. Total Flux Criterion 4. Gamma Flux Criterion . 5. Discussion of Results . 6. Biological Dose Rates ... .

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-

E.

P F.

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B.

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C.

V-38

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.. V-58 v-52

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- - -

. . V- 60 . . v-66 v-83 .. . v-83 . . v-84 . v-go *

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0

9

INSD-2650-29 vi

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'

v-93

. . V-95 * 0

.. -

- - - - - -

- - - . .................... ............ ........ 3. Approach . . .:. . . . . . . . . . . . . . . . . RTG Analysis and Prediction . 1. T h e m 1 Reliability . 2. Mechanical Reliability . . 3. Electrical Reliability - .... 4. Generator Prediction - . Reliability-Power Distribution - .1. Mathematical Model ... 2. State Probability ............. 3. Power Distribution . 4. Results . . . . . . . . . . . . . . . . . . . Performance Criteria 1. Objective. 2. Criteria

0

3

e

. . . . V-38

.. Magnetics. . . . . . . ............. 1. Units.. . . . . " . . . . . . . . . . . . . . 2. Equivalent Circuit of RTG . . .. 3. Evaluation of Experimental and Calculational . Results.. . . . . . . . . . . . . . . . . . 4. conclusions . . . . . . . . . . . . . . . . . . Ground Handling . . . . . . . . . . . . . . . .. 1. Ground Support Test Console . - .- - - -

2. Power Supply Rack 3. Fuel Capsule Shipping Cask 4. Generator Subsystem Handling Adaptor 5. Mobile Carriage and Handling Sling 6. RTG Subsystem Shipping Container 7. ETG Shipping Container 8. Portable Monitor Package

A.

- -

v-97 V-99

v-99 V-101 V-101

V-104

V-104 V-104

v1-1

VI-1 VI- 2 VI- 3

VI-4 5 v1-6 v1-14 v1-23 VI-

VI-

24

v1-25 v1-25 v1-26

v1-27


SUMMARY The principal goal of the NASA Viking project is to send two instrumented spacecraft, each consisting of a Lander and Orbiter, to Mars during the launch opportunity that occurs in the summer of

Arrival at Mars will be early in

1973.

1974, at which time the Viking Lander

Capsule (VLC) of each spacecraft will descend to the Martian surface for a 90 day minimum surface mission.

The landed weight will be

approximately 1100 pounds. The Orbiter will remain in Mars orbit and provide data storage and communications support for the Lander as well as gather scientific data from orbit. The Titan 111-C Centaur launch vehicle will be used for earth orbit insertion and subsequent insertion into the Mars interplanetary trajectory. The prime power source for the VLC portion of the spacecraft will consist of multiple radioisotope thermoelectric generator (RTG) units. The anticipated RTG electrical power requirement for the VLC is a minimum of

140 watts during the 90-day surface mission. This

value is based on data currently being considered by the Viking integrator contractor, Martin Marietta Corporation-Denver Division. The baseline RTG design selected for the Viking mission employs

TAGS-85/2N thermoelectrics and an intact impact heat source (IIHS) containing a weighs

675 thermal watt inventory of PU-238 fuel. Each RTG

28.2 pounds and produces a nominal 44.1 watts at 2.9 volts at

beginning-of-life (Br)L). At end-of-mission (EOM),conservatively assumed in this study as occurring one year after launch, the nominal

INSD- 2650- 29

vii _I-_-.-_

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power output is 40.7. With reliability considerations, the nominal value yields

40.7 watt

38.5 watts at a 0.99 reliability. Thus, four of

the baseline RTG's make up the RTG complement for the VLC.

The base-

line TAGS-85/2N RTG configuration is similar to that of units presently on test at Isotopes (see frontispiece).

At the writing of this report,

three TAGS-85/2N engineering units, two of which have been hardmounted and vibrated to simulated Viking dynamic levels, have been fabricated and performance tested. Other units are scheduled for fabrication during this calendar year.

TAGS-85/2N thermoelectric

materials were selected for this mission because of their demonstrated stability and high conversion efficiency. The existing SNAP 19 configuration was retained to the fullest practical extent due to the demonstrated dynamic and performance capability on the Nimbus program. Table S-1 summarizes the performance and configx-ation of the baseline design. Each Viking RTG occupies an envelope

10.75 inches high by 24.4

inches across the fin tips. Ninety TAGS-85/2N thermoelectric couples convert the thermal energy from the heat source to electrical energy.

In the current Nimbus design the couples are arranged in three seriesparallel strings (three couples in a branch, thirty branches in series), but it is anticipated that the arrangement will be modified for Viking to a two series-parallel arrangement ( t w o couples in a branch, fortyfive branches in series) to increase the load voltage from 2.9 to volts, thus enhancing converter reliability and joule efficiency.

I

INSD-2650 - 29 viii

4.4


TABLE S-1 SUMMARY OF VIKING TAGS-85/2N RTG CONFIGURATION AND PERFORMANCE CHARACTERISTICS

RTG System Number of RTG units

4

Power output at 2.9 volts (watts) BOL (nominal) (1) EOM (0.96 reliability) (1) EOM (0.99 reliability) ( -1)

176.4 156.8 154.0

Weight (pounds)

112.8

RTG Unit Thermoelectric materials %leg

~M-TEGS-~N( M) (2)

P-leg

Isotopes-TAGS-85with 0.100 inch SnTe segment

Heat Source Type Fuel Thermal inventory (watts)

Intact impact ( IIHS) Pu02 microspheres

675

Dimensions Height ( in. ) Diameter, across fin tips (in.) Weight (pounds)

10* 75

24.4 28.2

Nominal Performance Power output (watts) BOL EOM Load voltage (volts) (1) - Assumes parallel electrical hookup of RTG's

(2) Supplied by

JM'Company

(3) - Value in parantheses is for RTG configuration with two series-parallel strings for the thermoelectric circuit.

INSD-2650 - 29 ix


At fueling the R E is sealed with approximately one atmosphere of argon internal pressure. With the passage of time, helium generated by

D

fuel decay is released to the thermoelectric converter through a filter vent in the fuel capsule.

Simultaneously the argon and released helium

permeate to space through the RTG Viton O-ring seals in a controlled manner. This behavior is well understood and has a predictable influence on electrical power output and generator internal pressure and gas

P

composition. The baseline intact impact heat source (IIHS) design for Viking is consistent with requirements for maximum enhancement of nuclear safety. Analyses herein permitted nuclear safety evaluation and decisions with regard to design selection. Component design criteria was provided by evaluation of three major postulated abort categories:

a

launch pad, flight (orbital and superorbital) and impact/post-impact containment. In each case environments consequent to the malfunctions were considered as well as the response of each candidate configuration. The specific objective was to estimte the fuel release probabilities for each environment for the design. These release probabilities were combined with indices which are measures of the resultant nuclear safety and s m e d for the candidate heat source designs. The outcome was the selection of the baseline IIHS design. Integration studies performed included evaluation of mechanical,

Ji

thermal, electrical, nuclear radiation and magnetic interfaces as well as associated ground handling equipment. These studies show the base-

INSD- 2650- 29 X


line RTG design is compatible with expected mission environments and will satisfy the Viking mission electrical power requirements. Analyses were performed for both single RTG units and stacked u n i t s ;

Martin Denver is presently evaluating both arrangements.

1~~-2650-29 xi . ..

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I. INTRODUCTION

D A.

n

SNAP 19 NIMBUS RTG DESIGN

The SNAP 19 radioisotope thermoelectric generator (RTG) was developed by Isotopes, Inc., under contract with the U.S. Atomic Energy Commission for use on the Nimbus weather satellite. RTG subsystem S / N 9, launched aboard the Nimbus I11 satellite in April

D 0

1969, is representative of the

present operational configuration. The Nimbus I11 RTG subsystem, shown in Fig. 1-1, consists of two electrically and mechanically independent RTG units bolted together at two end cover flanges in a stack configura-

tion. The stacked RTG units are mounted on the Nimbus spacecraft sensory ring through a support base and standoff. All electronic equipment

a a 3

associated with the complete RTG power supply system (i.e., power and telemetry signal conditioners) is contained within modules in the spacecraft temperature-controlled sensory ring. The SNAP 19 Nimbus I11 RTG power supply system is designed to provide approximately 50 watts(e) of dc power at a spacecraft bus volcage of VDC during orbital operation for a minimum period of one year.

24.5

The RTG

subsystem supplies approximately 57 watts of unconditioned power at 2.7 VDC.

The electrical output is obtained by direct conversion of thermal

energy, produced during the normal decay of the radioisotope fuel, to electrical energy in a thermoelectric converter consisting of lead telluride (PbTe) thermoelectric elements. Waste heat from the conversion process is rejected to the environment by the finned RTG unit housings.

t

INSD-2650-29 1-1


generator

Lower generator

Space craft

-Support

base

-Support b a s e Standoff S / N 007 P o w e r conditioner

Telemetry signal conditioner S / N 006 F I G , 1-1.

SNAP 1 9 NIMBUS I I I R T G POWER S U P P L Y SYSTEM S / N 9


i

Figure 1-2 shows t h e c o n f i g u r a t i o n of t h e SNAP unit.

19 Nimbus

I11 RTG

Each u n i t c o n t a i n s a 625-watt thermal inventory derived from

plutonium dioxide ( PUO2) microspheres.

The Pu02 r a d i o i s o t o p e f u e l i s

contained w i t h i n an i n t a c t r e - e n t r y h e a t source (IRHS) c o n s i s t i n g of a s u p e r a l l o y f u e l capsule and g r a p h i t e r e - e n t r y body.

The superalloy

capsule a s s u r e s containment of t h e r a d i o i s o t o p e f i e 1 by providing impact p r o t e c t i o n f o r i n a d v e r t e n t ground handling and launch pad a c c i d e n t s and low a l t i t u d e launch a b o r t s where t h e r e - e n t r y h e a t i n g p u l s e i s l o w .

The

g r a p h i t e r e - e n t r y body provides f o r f u e l containment during r e - e n t r y from h i g h a l t i t u d e launch a b o r t s and o r b i t a l decay from s h o r t - l i v e d , o f f optimum o r b i t s .

The g r a p h i t e housing remains i n t a c t f o r a l l c r e d i b l e

r e - e n t r y h e a t i n g c o n d i t i o n a t o a s s u r e that t h e f u e l i s not dispersed i n t h e upper atmosphere.

For t h e h i g h a l t i t u d e a b o r t and o r b i t a l decay

r e - e n t r y modes, t h e s u p e r a l l o y f u e l capsule i n some cases may m e l t w i t h i n t h e g r a p h i t e housing during r e - e n t r y t o form a s o l i d i f i e d fuel/metal m a t r i x p r i o r t o impact.

A f t e r e a r t h impact t h e f u e l i s j m o b i l i z e d i n

t h e m a t r i x and confined t o t h e p o i n t of impact,

INSD-2650-29

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a 1

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B.

GUIDELINES FOR V I K I N G RTG STUDY

This study of the a p p l i c a t i o n of SNAP 19 RTG technology t o t h e Mars Viking mission has been conducted by Isotopes i n accordance with t h e applicable p o r t i o n s of AEC Contract AT( 29-2) -2650.

In directing the

a c t i v i t i e s of t h i s study, t h e following t e c h n i c a l guidelines were established: The RTG/spacecraft e l e c t r i c a l i n t e r f a c e w a s assumed t o be a t t h e RTG e l e c t r i c a l connector plug, and t h e mechanical i n t e r f a c e t o be a t t h e RTG mounting flange. The TAGS-85/2N RTG u n i t s p r e s e n t l y being f a b r i c a t e d and t e s t e d a t Isotopes are t h e b a s e l i n e Viking u n i t s . Performance parameters f o r t h e SNAP

19 3P/2N Nimbus

I11

design a r e a l s o presented.

1 1

The b a s e l i n e heat source provides intact-impact c a p a b i l i t y (IIHS), with t h e a l t e r n a t e being t h e present SNAP i n t a c t re-entry heat source (IRHS)

19

.

RTG system configurations considered include s i n g l e generator

u n i t s as well as t h e present SNAP

19 stacked,

two-unit con-

f i g u r a t i o n t o permit s p a c e c r a f t i n t e g r a t i o n f l e x i b i l i t y .

J

End-of-mission i s considered t o occur approximately one year

a f t e r launch, however generator performance up t o f i v e years after launch i s presented.

J

INSD-2650 -29

1-5


C.

SCOPE OF VIKING RTG STUDY

The SNAP 19 Nimbus RTG system was developed and optimized for the Nimbus spacecraft and an Earth orbit operating environment.. It has been the purpose of this study to identify and evaluate those modifications required and/or desirable for adaptation of the RTG to the Viking Lander Capsule (VLC).

This work has been performed in conformance with the

guidelines set forth in the preceding section and guidance obtained through discussions with AEC, Sandia, NASA Project and Viking Integrator Contractor personnel. The principal consideration which gave direction to the study was the requirement for from two to four RTG units to satisfy the VLC power profile for a minimum operational period of

90 days

on the Martian surface. The number of units to be selected and power requirements depend on the results of studies now being performed by the Viking Integrator Contractor, Martin Marietta Corporation-Denver Division, which was awarded the Viking contract in May

1969. Martin Denver

is

currently evaluating an all-RTG power system in addition to the hybrid RTG/battery system presented in their Viking proposal. Present indications are that an RTG unit with a minimum power output of approximately

36 watts at end-of-mission is desirable for the all-nuclear, four RTG configuration, while a minimum output of approximately 25 watts per RTG is sufficient for the two-RTG, hybrid configuration. Current emphasis at Martin Denver and NASA is being placed on the four-RTG configuration, and thus has been given priority in this study.

INSD-2650-29 I-6


3

The Viking RTG study has been primarily concerned with the following areas: ( 1) Thermoelectric converter design and performance (2) Heat source design and re-entry performance

1

( 3) Housing/radiator weight optimization

1

(4) RTG/spacecraft interface considerations. A brief description of each area of study is presented below. 1.

Thermoelectric Converter The present SNAP 19 Nimbus I11 thermoelectric converter contains

lead telluride N and P-type thermoelectric elements. Thermoelectric material designations are 3M-TEGS-3PB for the P-type material and 3MTEGS-2N(M) for the N-type.* These materials have demonstrated stable

J

performance in long-term endurance testing of both fueled and electrically heated SNAP units on the SNAP 11, 1-9and 29 programs. Although the experience factor is an incentive for retaining these materials, a signif-

7

icant increase in system efficiency and specific power can be realized with a change in only the P-leg thermoelectric material. An Isotopes proprietary silver antimony telluride/germanium telluride P-type thermoelectric compound (designated TAGS-85) has demonstrated stable performance in life testing of couples, modules and commercial RTG units. This material has replaced the lead telluride P-leg previously used in the SNAP

29 based on module and couple testing performed on that program.

“mnese thermoelectric elements are supplied by 3M Company.

INSD-2650-29

1-7


E SNAP 29 RTG modules and couples a r e p r e s e n t l y on t e s t incorporating t h i s Also, t h r e e engineering e l e c t r i c a l l y heated t e s t u n i t s of a

material.

Viking type R E incorporating t h e TAGS-85 material i n the P-leg of t h e thermoelectric couples have been f a b r i c a t e d a t Isotopes, and i n i t i a l e l e c t r i c a l performance and dynamic t e s t i n g performed.

Additional Viking

type u n i t s are scheduled f o r f a b r i c a t i o n and t e s t i n g during t h e current calendar year. Power output from f u e l i n g of t h e RTG through f i v e years a f t e r launch has been determined f o r both t h e TAGS-85/2N and 3P/2N thermoelectric configurations.

Variations i n RTG performance due t o changes i n f u e l load-

i n g , s e l e c t i o n of i n e r t f i l l gas, and housing s e a l leakage have been

conducted. 2.

Heat Source The b a s e l i n e Viking RTG

u n i t can be mated with e i t h e r t h e e x i s t i n g

i n t a c t r e - e n t r y h e a t source (IRHS) or t h e i n t a c t impact heat source (IIHS) which i s c u r r e n t l y being developed a t Isotopes.

The b a s i c d i f f e r e n c e

between t h e IRHS and IIHS i s i n t h e f u e l capsule m a t e r i a l s technology. The SNAP

19 IRHS u t i l i z e s a single-wall superalloy vented capsule,

whereas a higher melting, double l a y e r , vented capsule i s required t o achieve t h e I s o t o p e s ' IIHS concept of containment of t h e radioisotope f u e l during r e - e n t r y and on E a r t h impact.

The IIHS concept incorporates

a hex-shaped g r a p h i t e h e a t accumulator block a s a n i n t e g r a l p a r t of t h e g r a p h i t e heat s h i e l d .

This modification enhances t h e success p r o b a b i l i t y

f o r i n t a c t re-entry from s u p e r o r b i a l v e l o c i t i e s by providing a d d i t i o n a l

INSD-2650-29

I-8


J

heat shield ablation/heat sink capability and for fuel containment on earth impact by lowering the impact velocity through increased drag.

3.

Housing/Radiator Optimization The SNAP 19 Nimbus I11 subsystem consists of two RTG units bolted

end-to-end in a stack arrangement but otherwise mechanically and electrically independent. This study has considered optimization of the housing and fins for the Viking mission thermal and dynamic environments,

B

and has considered both single RTG's and stacked configurations.

4.

Spacecraft Integration Use of RTG power for the Viking Lander Capsule requires an examina-

tion of the various technical interfaces between the spacecraft and RTG system. These interfaces can generally be categorized as either mechanical, thermal, electrical, magnetic, nuclear radiation o r operational support in nature. During the study these interfaces were examined quantitatively where possible and qualitatively where specific definition was not available,

INSD-2650-29 1-9


11. VIKING RTG OPTIONS This chapter presents the results of performance studies performed for the Viking RTG options, The baseline RTG configuration is a TAGS-

85/2~unit employing a 675 thermal watt heat source. The 3P/2N RTG configuration is similar to the SWP

19 Nimbus I11 units incorporating

625 thermal watts. Performance parameters are presented for both the baseline intact impact heat source (IIHS) and the alternate intact

.

re-entry heat source (IRHS)

A.

TAGS-85/2N RTG

The TAGS-85/2N RTG, considered as the baseline design for the Viking mission, is a modified SNAP

2

3

a 1

19 Nimbus RTG. Figure 11-1 depicts

the modified design, and shows t h e i n t e r n a l d e t a i l s of a u n i t . Each RTG unit weighs 28.2 pounds and produces a nominal beginning-of-life, decreasing to

40.7 watts

44.1 watts at

at end-of-mission, conser-

vatively defined in this study as one year after launch. The baseline heat source design is a 675 thermal watt IIHS.

An RTG unit occupies an

envelope approximately 11 inches l o n g by 24.4 inches across the fin tips. The internals of the generator are sealed by Viton O-rings at each end of the finned magnesium thorium alloy cylindrical housing. The initial fill gas is argon at approximately 1 atmosphere absolute pressure.

c3 3

1.

Design Description A detailed description of each major component in an RTG unit

follows:

INSD-2650-29 11-1


F I G . 11-1. V I K I N G RTG CONFIGURATION

I N S D - 2 6 50 -2 9 11-2


13 a.

J

3

ri

a 3

Thermoelectric converter

The thermoelectric converter consists of six thermoelectric modules containing a total of 90 thermoelectric couples connected electrically in three series-parallel strings. To enhance power conditioning equipment efficiency and to decrease joule losses, the 90 couples may be electrically connected in two series-parallel strings in future SNAP 19 units to increase the output voltage

5%.

Reliability analyses presented

in Chapter VI show that virtually no reliability penalty is entailed with this circuit configuration modification. The converter thermoelectric materials are Isotopes-TAGS-85 with a 0.1 in. thick SnTe segment at the hot end (P-leg) and 3M-TEGS-2N(M) (N-

leg), selected because of their demonstrated high conversion efficiency and long-term stability in various SNAP-type devices and the Isotopes

J

1 1

commercial RTG (Sentinel) line. The RTG designation TAGS-85/2N thus signifies the N and P leg materials.

Figure 11-2 is an exploded view

of the baseline couple. This couple design is identical in materials and fabrication to that being used in the SNAP

29 RTG design and in

t h r e e e x i s t i n g Viking t y p e e n g i n e e r i n g RTG's ( S / N

26, S / N 27, a n d S / N

28) presently on test.

J i 3

Figure 11-3 shows a single thermoelectric module with 15 couples. Also shown in the figure is the module cold end hardware which consists

of alignment buttons, springs, pistons and heat sink bars. This hardware provides compressive forces on the thermoelectric elements through alignment buttons adjacent to the copper connecting strap thermoelectric leg. One side of the button is flat to permit radial movement and the

3

-

INSD 26 50 -2 9 11-3


Electrical strap (oxygen- free CUI

a a SnTe bonding w a f e r

sPb o l d eIn r disc (2)

Fe cold c a p s ( 2 )

d

N i plate ( N only)

(Ponly)

TAGS-85 ( 0 . 2 7 0 i n . d i a m e t e r x 0 . 4 0 0 in. long) Bonding wafer ( P only)

3M-TEGS-ZN(M) ( 0 . 3 7 7 in. d i a m e t e r x 0. 500 in. long)

P

/

S n T e s e g m e n t ( P only) ( 0 . 2 8 5 i n . d i a m e t e r x 0. 100 in. long)

Fe cup ( 2 )

Braze wafer (2)

N i hot s h o e

FIG. 11-2.

T A G S - 85/2N C O U P L E CONFIGURATION

INSD- 2 650- 2 9 11-4


3

,-

Electrical connector

._

/

i

N element (typical) P element (typical)

3

'-

Module

eat sink bar Spring Piston Alignment button Connector strap

Hot shoe Thermoelectric couple

Thermal insulation

FIG. 11-3.

THERMOELECTRIC MODULE

INSD -2 650 -2 9 11-5

AND COLD END HARDWARE

"N "


e

other side is spherical to allow for angular misalignment, A springloaded piston with a concave surface bears on each alignment button to allow for axial movement and keeps the elements in compression. The springs and pistons are retained by a heat sink bar machined to match the generator housing inner diameter and to accept the pistons with minimal clearance to reduce the temperature drop across this interface. The buttons, pistons and heat sink bars are made of 6 0 6 ~ aluminum ~ 6 alloy, selected for its high thermal conductivity and low density. Each cold end component (except the spring) is hardcoated (a process which results in a thin oxide coating on the surface of the treated parts) to provide multiple redundant electrical insulation barriers. The six thermoelectric module assemblies are positioned at 60degree increments around the hex-shaped heat source. Electrical insulation at the hot end of the thermoelectric modules is provided by a mica sheet. The thermoelectric couples within each module assembly are wired in 3 series-parallel strings, with the six modules connected in series. Figure 11-4 shows the alternate two series-parallel string electrical schematic for the thermoelectric circuit. Also included in the figure is the cold strap hardware configuration required to achieve this electrical arrangement. Johns Manville Type 1301 Min-K thermal insulation is used for both the thermoelectric modules and at both ends of the heat source to minimize heat losses. The Min-K is cast into the required configurations at JohnsManville and trimmed upon final installation in the generator.

-

INSD-2650 29 11-6


1

a

. To electrical connector

Mod. 1

Mod.

Mod.

Mod.

Mod.

Mod.

2

3

4

5

6

---

---

b

To electrical connector

7--

n Module Interconnection (two parallel branch shown)

I I I

n-

I

I

I

I I I I I II I I

I I I

I I

I

I I

I I

I

I

I I I I Couple equivalent circuit

--------1 ; L II I+ I

I I I I

I

I

I

I

I

I I

I I

I

I I I I I I I

I I

I I I

I I I I I I I

I

I

I

I I I I

I I

I I I

I I

I

I L----------l Module 3

I I

I

I

I I I

I I I I I I I I I I

I

I I

I

L

---------. Module 4

Existing SNAP 19 Triple P a r a l l e l Branch Electrical Schematic

I

vi

D BU

I I I I

I I

I I

I

I I I I

I

I

I

I I I I I I

r!d

I

I I I

E+y %I I

L-,-------J

1

I----------.

Couple Interconnection (two parallel branch)

FIG. 11-4. ELECTRICAL SCHEMATICS OF T/E CONVERTER ASSEMBLY

INSD - 2650-29 11-7

Module 4

Modified SNAP 19 Two P a r a l l e l Branch Electrical Schematic

To next module

%.

I

I I I I

Module 3

To next module

n

I 1 I I

r----------

I I I I I


The thermoelectric converter (along w i t h t h e r e s t of t h e i n t e r n a l s of t h e generator) i s i n i t i a l l y s e a l e d i n argon.

With t h e generation of

E

helium a s s o c i a t e d with t h e radioisotope f u e l decay and t h e subsequent r e l e a s e of t h i s gas through t h e heat source f i l t e r , a helium/argon atmosphere i s e s t a b l i s h e d i n t h e RTG.

(Additionally, some nitrogen and oxygen

e n t e r from t h e ambient a i r during storage with t h e oxygen being g e t t e r e d by a zirconium sponge contained i n an annular s t a i n l e s s s t e e l cup next t o the heat source.)

The Viton O-rings i n t h e housing permit a c o n t r o l l e d

r e l e a s e of these gases by permeation.

The r e s u l t a n t equilibrium pressure

i n t h e generator i s c o n t r o l l e d by the leakage r a t e through the s e a l s and t h e r a t e of helium r e l e a s e from t h e f’uel. b.

Housing/radiator

The RTG housing ( f i n n e d c y l i n d e r and end cover d e t a i l s ) i s f a b r i c a t e d from a magnesium-thorium a l l o y .

The end covers a r e b o l t e d t o t h e c y l i n d e r

t o p and bottom flanges with t i t a n i u m b o l t s and nuts, and provide preload for

(1) t h e Viton O-ring s e a l s i n s t a l l e d a t t h e s e j o i n t s and ( 2 ) t h e f u e l

block through predetermined compression of t h e Min-K thermal i n s u l a t i o n . The e l e c t r i c a l power o u t l e t i n each u n i t , a l s o s e a l e d with a Viton O-ring,

i s mounted i n t h e side of t h e c y l i n d r i c a l housing s e c t i o n .

S i x f i n s are

welded a t 60-degree increments around t h e c y l i n d r i c a l housing t o provide t h e required r a d i a t o r area. Mounting of t h e i n d i v i d u a l RTG u n i t s t o t h e spacecract i s provided by s i x holes equally spaced on a 7.5 inch diameter c i r c l e i n t h e flanges of t h e RTG u n i t .

This attachment arrangement may e a s i l y be modified i f

deemed necessary t o s a t i s f y a p a r t i c u l a r mounting requirement.

INSD-2650-29 11-8

I


1 ' 7

The housing/radiator design employs t h e same b a s i c design f e a t u r e s

3

1 3

used on previously b u i l t and t e s t e d SNAP 1-9 u n i t s with some modifications t o reduce weight.

The c y l i n d r i c a l housing i s t h e same diameter and w a l l

thickness as previous u n i t s .

The f i n s and end cover thickness have been

weight optimized f o r t h e Viking mission environments.

All housing, f i n and end cover ( r a d i a t o r ) surfaces are coated with a zirconium oxide/sodium s i l i c a t e mixture developed under t h e SNAP

13 1,

I 3

3

9A

program and subsequently used on both t h e SNAF 11 and SNAP 19 RTG systems. The coating c o n s i s t s of f i n e l y m i l l e d zirconium oxide, which i s bonded t o the r a d i a t o r s u b s t r a t e by a sodium s i l i c a t e inorganic binder.

This

coating composition i s a t t r a c t i v e f o r space RTG r a d i a t o r s s i n c e it e x h i b i t s a high surface e m i s s i v i t y (0.83) i n conjunction with a low solar absorptivity (0.2).

SNAP

9A data have demonstrated t h e e x c e l l e n t

s t a b i l i t y of t h i s coating i n a space environment f o r a period now approaching s i x years. e.

Heat source

The proposed heat source f o r t h e Viking RTG i s a hexagonal r i g h t p r i s m i n configuration and u t i l i z e s a helium-vented,

two-layer capsule.

The capsule inner l i n e r i s a molybdenum-50$, rhenium a l l o y t o provide f o r

I

a J

J

f u e l compatibility.

The o u t e r l i n e r , which functions as t h e s t r e n g t h

member, i s constructed from TZM (titanium-zirconium-molybdenum) f o r p r o t e c t i o n a g a i n s t severe mechanical environments such as impact, b l a s t The capsule contains 675 thermal watts of PUO f u e l . For 2 r e - e n t r y p r o t e c t i o n , a g r a p h i t e heat s h i e l d f a b r i c a t e d from Carb-I-Tex

and shrapnel.

INSD-2650-29

11-9


c g r a p h i t e i s provided.

The h e a t source i s designed t o provide f o r i n t a c t -

impact r e - e n t r y of t h e h e a t source through r e - e n t r y heating and subsequent e a r t h impact.

A summary of design c h a r a c t e r i s t i c s f o r a TAGS-85/2N RTG u n i t i s presented i n Table 11-1 and a weight summary i s presented i n Table 11-2. 2.

Predicted Performance The performance of t h e Viking RTG depends p r i m a r i l y on t h e follow-

i n g parameters : (1) thermoelectric s t a b i l i t y ,

(2)

i n i t i a l f i l l gas,

(3)

percentage of f u e l decay-generated helium released

from the plutonium f u e l ,

( 4 ) t i g h t n e s s of t h e RTG housing seals, and

(5)

f u e l decay.

Thermoelectric s t a b i l i t y i s discussed i n Section I l l - B and parameters r e l a t i n g t o t h e RTG f i l l gas a r e discussed i n Section 1 1 1 - C .

The

nominal TAGS-85/2N RTG performance that results from v a r i a t i o n s on t h e s e parameters i s presented i n t h i s s e c t i o n .

For d e t a i l e d r e l i a b i l i t y

a s p e c t s of t h e RTG performance refer t o Chapter V I .

a , & I n i t i a l argon f i l l Power and temperature p r o f i l e s f o r generators f i l l e d with argon a t t h e time of f u e l i n g are presented i n Fig.

11-5, 11-6and 11-7 f o r helium

r e l e a s e r a t e s (from t h e f u e l ) of 106, 50% and 25%, r e s p e c t i v e l y .

Net

power output and RTG temperatures are presented parametric i n load voltage

INSD-2650-29 I1-10

i

E


a s 3

TABLE 11-1 TAGS-85/2N RTG UNIT DESIGN CHARACTERISTICS FOR

675 WATT IIHS

T/E Converter Conversion materials

J

Hot shoe material

Nickel

Hot end cups for T/E

Ferro-Vac-E iron

Cold end caps for T/E

Ferro-Vac-E iron

Number of T/E modules

6

Number of T/E couples

90

N-element dimensions Diameter (in.) Length (in.)

3 i

0.377 0.500

P-element dimensions Diameter (in.) Length (in.)

0.270 0.100 SnTe +

Thermal insulation

Johns-Manville Type 1301

Cold end electrical strap material

Oxygen-free Cu

Number of parallel strings

3

Heat sink assembly

J 3

Number of assemblies Pistons, hardcoated A1 6 0 6 ~ ~ 6 Alignment buttons, hardcoated AL 6 0 6 ~ ~ 6 Springs, SS Module bar, hardcoated A 1 6 0 6 ~ ~

a 3

0.400 TAGS

INSD-2650-29 11-11

6 180

6

180 180

6


E TABLE 11-1 (Continued)

Heat Source Radioisotopic fuel

Pu02 microspheres

Fuel power density rwatts/cm3)

2.7

Materials Carb-I-Tex graphite TZM Mo-5O% Re Ta felt Ta

Heat shield Strength member Inner liner Compliant support pads Compliant support cup Dimensions Hex-shaped heat shield Across flats (in.)

3.500 6.750

Length ( i n . )

0.562

Wall thickness (in.) Capsule strength member

ID (in.) Wall thickness (in.) Capsule thermal inventory at BOL (watts)

2.030 0.060

675

Capsule inner liner

ID (in.) Wall thickness (in.) Housing/Radiator Housing material Fin and end cover material Housing OD at flanges (in.)

8.15

Housing OD (in.)

6.45

Length, including end covers (in.)

10

INSD-2650-29 11-12

75


TABLE 11-1 (Continued)

9

Overall diameter, including fins (in.)

24.4

End cover minimum wall thickness (in.)

0.156

Housing w a l l thickness (in.)

0

Assembly bolts and nuts

No. 10 and 1/4 in. diameter Ti alloy

O-ring seals

Viton-A

093

Fin dimensions Length, root-to-tip (in.) Height ( in.) Root thickness (in.) Tip thickness (in.)

a

9.0

9.5 0.3 0.03

Number of fins

6

Coating on fins, housing and end covers

silicate binder

Bolts

Ti

Zirconia with sodium

Miscellaneous Hot end electrical insulation

Mica

3

Getter

Zirconium sponge in stainless steel container

J

Initial fill gas Argonat approximately 1 atmosphere

3

INSD-2650-29

11-13


TABm 11-2

TAGS-85/2N RTG UNIT WEIGHT SUMMARY

Item T/E Modules

Weight (lb)

7.86

(6)

Couples, straps and w i r i n g

3.22

Cold end hardware

3.91

Min-K i n s u l a t i o n

0.73 10.24

Hous ing/Radiat or

7.66

Housing and f i n s End cover, f u e l i n g end

1.16

End cover, joined end

1.02

Retaining hardware and O-rings

0.40

Heat Source (IIHS;

8.14

675 watts)

Fuel and Z r 0 2 f i l l e r

4.10

Capsule components

1.78

Heat s h i e l d c mponents

2.26

1.94

Miscellaneous Min-K a t ends

0.84

Getter

0.75

Miscellaneous ( shims, mica, p l q )

0.35 28.2

Total

I NSD-26 50 -29 I1-14


a

INSD - 2650 - .2 9. 11-15

.'6-


.

INSD-2650-29 11-16


i -J

3

- -

-.

..

INSD - 2 6 5 0 - 2 9 11-17

I ,

\L


f o r s i x mission times up t o

5 years.

The l o a d voltages shown i n t h e

f i g u r e s assume t h e present t h r e e s e r i e s - p a r a l l e l s t r i n g e l e c t r i c a l arrangement.

For t h e two s e r i e s - p a r a l l e l s t r i n g arrangement, t h e l o a d

voltage w i l l be

50% higher with t h e power output v i r t u a l l y i d e n t i c a l t o

t h a t shown f o r t h r e e s e r i e s - p a r a l l e l s t r i n g s .

For a l l performance analyses

i n t h i s r e p o r t , t h e present t h r e e s e r i e s - p a r a l l e l arrangement i s assumed except where noted otherwise. Generator power i s r e l a t i v e l y f l a t between

2.6 and

3.2 v o l t s , and

peak power p o i n t s occur between 2 . 8 and 3.0 v o l t s during the 5-year period analyzed and t h e range of helium r e l e a s e r a t e s expected.

Hot

junction temperature v a r i e s approximately l i n e a r l y w i t h load voltage and g e n e r a l l y decreases with t i m e a t any l o a d voltage. ness f a c t o r of 3 x

A nominal s e a l t i g h t -

E L

10-5 scc/sec, t y p i c a l f o r SNAP 19 type generators,

was assumed t o construct t h e curves.

Complete r e l e a s e (100% rate) of

a l l helium generated by f u e l decay, which apparently has occurred on t h e Nimbus 111 generators S/N

24 and S/N 25 ( s e e Section

111-C),

6

i s conserva-

t i v e f o r power p r e d i c t i o n analyses and was assumed h e r e i n i n conjunction with t h e 3 x 10-5 scc/sec seal t i g h t n e s s f o r nominal performance predictions.

E

The e f f e c t of seal t i g h t n e s s and helium r e l e a s e r a t e on power i s

m i n i m a l as shown i n Fig. 11-8. Power output a t t h e end-of -mission ( one year a f t e r launch) v a r i e s less t h a n

4% about

t h e nominal f o r the range of

s e a l t i g h t n e s s and helium r e l e a s e rates expected.

The maximum e f f e c t on

reducing power from t h e nominal expected value a t end-of-mission,

INSD-2650-29 11-18

40.7

E


H

2

v)

H a

H I I

N

w m wcn

0 I

N W


watts, is seen to be less than 276, and occurs for 100% helium release and the tightest s e a l , 1 x

scc/sec.

From Fig, 11-9, the beginning-of-life design hot junction temperature is 1 0 5 8 ~ at~ a cold junction temperature of 39OoF and a fin root temperature of 3309. Assuming 100% helium release from the fuel and

3 x 10-5 scc/sec tightness to be nominal values, the hot junction temperature decreases to lo099 over the period from nine months pre-

E;

E

ceding launch to one year after launch. During this same period, the nominal power changes from

44.1 watts per RTG to 40.7 watts at the 2.9 c

load voltage design point which results in maximum power at end-ofm i s s i o n f o r the nominal d e s i g n c o n d i t i o n s .

For the Viking mission, electrical power may not be required at all times, such as during interplanetary cruise, and the RTG may be

placed on short circuit if desired.

Thus the power profile shown in

Fig. 11-9 may be considered as an "available" power curve. The primary effect on the temperature profiles of short circuiting the RTG will be to decrease the hot junction by approximately l 0 O q from the values shown assuming the fin root temperature does not vary appreciably from the 33OoF design value (lOO?F effective environmental sink temperature). Table 11-3 summarizes the performance parameters for the Viking

TAGS-85/2N RTG for beginning-of-life (BOL) and end-of-mission (EOM) conditions. As is shown in Chapter VI, a 40.7 watt EOM nominal power output per RTG results in a

0.99 probability of obtaining 154 watts

from four such units connected in parallel

INSD-2650-29 11-20

(38.5 watts per RTG),

Thus

i


a

3

11-21

INSD -2 650 -2 9


TABLE 11-3

PERFORMANCE SUMMARY FOR TAGS-85/2N RTG DESIGN

(INITIALARGON FILL) B OL' -

E O2 -

Nominal input power (watts)

675

666

Nominal output power (watts)

44.1

40.7 (38.513

RTG efficiency (%)

6.53

6.11

RTG specific power (watts/lb)

1.56

1.44

2.9 (4.4)

2.9 (4.4)

Load current (amps)4

15.2 (10.0)

14.0 (9.2)

Hot junction temperature (OF)

1058

1009

Cold junction temperature

390

365

330

327

14.7

7.3

0

59

Load voltage ( volts)

4

Fin root temperature

(OF)

(OF)

Total internal pressure (psia)5 5

Helium concentration (%)

'Beginning-of-life, defined as at fueling

E E i

E

--

9 months prior to launch.

2End-of-mission, defined as one year after launch. butput power at EOM for a

0.99 reliability shown in parentheses.

)+Values in parentheses are for two series-parallel strings. ?Based on 100% He release from capsule and 3 x 10-5 scc/sec seal tightness (see Section 111-C).

/--

INSD-2650-29 I1-22


a 2.5 w a t t margin i s provided over t h e

36 watt minimum RTG power r e q u i r e -

ment based on present Martin Denver s t u d i e s f o r an a l l - R T G power system. A s i s shown i n Section 1 1 1 - A ,

B

t h e nominal BOL performance of a

TAGS-85/2N RTG with t h e a l t e r n a t e Nimbus I11 625-thermal w a t t IRHS i s

37.8 watts.

End-of -mission nominal power i s approximately 34.8 w a t t s .

No f u r t h e r performance increase with t h e present SNAP

r e a l i z e d , as b.

625 watts

19 IRHS can be

i s i t s present capacity.

I n i t i a l helium f i l l RTG

The p r i n c i p a l advantage of a n i n i t i a l helium f i l l i n l i e u of a n argon f i l l i s a somewhat f l a t t e r e l e c t r i c a l power p r o f i l e and lower hot j u n c t i o n operating temperatures as i l l u s t r a t e d i n Figs. 11-10, 11-11, and 11-12, a t t h e expense of a lower output f o r mission periods of less

than t w o years.

The five-year power and temperature l e v e l s , however,

a r e e s s e n t i a l l y equivalent t o those f o r argon f i l l e d generators, s i n c e by t h a t t i m e t h e equilibrium gas compositions and pressures a r e nearly

a

t h e same.

Post f u e l i n g , BOL, power of t h e helium generator i s 38.3

watts as compared t o 44.1 watts f o r an argon f i l l e d u n i t .

The temporary

power and temperature increases o f generators, shown i n Fig. 11-13, r e s u l t from t h e pressure and helium concentration changes f o r a period a f t e r launch. For t h e Viking mission, an i n i t i a l argon f i l l o f f e r s a s i g n i f i c a n t

1

performance advantage a t one-year a f t e r launch, 40.7 watts compared t o

38.7 watts ( a t 100% r e l e a s e r a t e ) , and t h u s i s t h e recommended fill gas approach.

1

:�

A s an i n i t i a l helium f i l l o f f e r s no performance advantage


INSD-2650-29 11-24


P

a a

e$

INSD -2650 -2 9 11-25


B

4.

@!

INSD - 2650- 29 11-26


I

I

a

a a 11-27

-. . . INSD -2650-29

'.

'-

-


over an argon fill f o r t h e Viking mission, t h e former was not considered f u r t h e r i n t h i s study.

B.

3P/2N RTG

I n t h i s s e c t i o n t h e performance of a 3P/2N RTG i s presented f o r t h e Viking mission.

With t h e present Nimbus I11 625 w a t t f u e l inventory,

t h e RTG weight i s 26.7 pounds and t h e nominal power output i s 26.6 w a t t s a t end-of-mission. 1.

Design Description The 3P/2N RTG u n i t design option evaluated f o r t h e Viking mission,

is similar t o t h a t o f t h e p r e s e n t SNAP 1-9 Nimbus I11 RTG w i t h the f o l l o w -

i n g exceptions: a.

The f i n s a n d h o u s i n g h a v e b e e n

o p t i m i z e d f o r the Viking

mission t o t h e configuration shown i n Fig. 11-1. b.

A 625 w a t t intact-impact heat source (IIHS), i d e n t i c a l

t o t h a t f o r t h e TAGS-85/2N design except f o r f u e l loading, has been s u b s t i t u t e d f o r t h e Nimbus I11 625 w a t t i n t a c t r e - e n t r y h e a t source (IRHS) with t h e l a t t e r considered

as a n a l t e r n a t e . With t h e exception of t h e thermoelectric converter P-leg and shoe m a t e r i a l s ( s e e Fig. 11-14), t h e d e t a i l e d d e s c r i p t i o n and construction m a t e r i a l s are i d e n t i c a l t o t h o s e previously presented for the TAGS-85/ 2N RTG i n Section 1 1 - A , A summary of design c h a r a c t e r i s t i c s and weight parameters f o r a

3P/2N RTG u n i t i s presented i n Tables 11-4 and 11-5.

INSD-2650-29 11-28


P

trap

Q-

R

Bonding wafer ( P only)

Ff 7 l

P and N element dimensions: 0.500 i n . long x 0.377 i n .

diameter

w-,

Bondin@;wafer ( P only) Fe hot shoe

J

Fig. 11-14, 3P/2N Couple Design

1. INSD - 2650- 29 11-29


TABLE 11-4

3 ~ / RTG 2 ~ UNIT DESIGN CHARACTERISTICSFOR 625 WATT IIHS

TIE Converter Conversion materials

3P/2N

Hot shoe material

Ferro-Vac-E iron

Cold end caps for T/E

Ferro-Vac-E-iron

Number of T/E modules

6

Number of T/E couples

90

N-element dimensions

0.377

Diameter (in.) Length ( in.)

0.500

P - e l e m e n t dimens ions

0.377

Diameter (in.) Length (in.)

0.500

Thermal insulation

Johns-Manville Type 1301

Cold-end electrical strap material

Oxygen-free copper

Number of parallel strings

3

Heat sink assembly Number of assemblies Pistons, hardcoated A1 6 0 6 ~ Alignment buttons, hardcoated

6 ~6

A1 6 0 6 ~ ~ 6 Springs, SS Module bar, hardcoated A1

606~~6

180

180 180 6

Heat Source Radioisotopic fuel

Pu02 microspheres

Fuel power density (watts/cm3)

2.7

Materials Carb -I-Tex graphite

Heat shield Strength member

TZM

INSD-2650-29 11-30


'&}

\,+

tL

TABLE 11-4 (Continued)

Mo-50$ R e Ta f e l t

Inner l i n e r Compliant support pads Compliant support cup

Ta

Dimensions Hex-shaped heat s h i e l d Across f l a t s ( i n . ) Length ( i n . ) Wall thickness ( i n . )

3- 500 6.750 0.562

Capsule i n n e r l i n e r I D (in.) Wall thickness ( i n . )

0.020

1 970

Capsule s t r e n g t h member I D (in.) Wall thickness ( i n . ) Capsule thermal inventory a t BOL (watts)

2.030 0.060

625

Housinglfladiator

I

Housing material F i n and end cover material

MgTh ( W U - T 8 )

Housing OD a t flanges ( i n . )

8.15

Housing OD ( i n . )

6.45

Length, including end covers ( i n . )

10.75

Overall diameter, including f i n s ( i n . )

18.0

End cover minimum w a l l thickness ( i n . )

0.156

P

Housing minimum wall thickness ( i n . )

0.093

Assembly b o l t s and nuts

1

No. 10 and 1/4 i n . diameter T i a l l o y

O-ring s e a l s

V i ton-A

F i n dimensions Length, r o o t - t o - t i p ( i n . ) Height ( i n . )

11

INSD-2650-29 11-31

5.8 995


TABLE 11-4 (Continued) 0.30 0.03

Root thickness (in.) Tip thickness (in.) Number of fins

6

Coating on fins, housing and end covers

Zirconia with sodium s i l i c a t e binder

Bolts

Ti

Miscellaneous Hot end electrical insulation

mica

Getter

Zirconium sponge in stainless steel container

Initial fill gas

Argon at approximately 1 atmosphere

INSD-2650-2 9 11-32


. .

. . .

-

D --?

w

TABLE 11-5

3P/2N RTG UNIT WEIGHT SUMMARY

13

Weight ( l b )

Item

8.35

T/E Modules (6) Couples, s t r a p s and wiring

3.71

Cold end hardware

3.91-

Min-K i n s u l a t i o n

0.73 8.30

Hous i ng/Radia t or

Housing and f i n s

5-72

Ena cover, f u e l i n g end

1.16

End cover, joined end

1.02

Retaining hardware and O-rings

0.40

Heat Source ( I I H S

i

a

-

8.09

625 watts)

Fuel

4.05

Capsule components

1.78

Heat S h i e l d components

2.26

1.94

Miscellaneous

Min-K a t ends

0.84

Getter

0.75

Miscellaneous (shims, mica, plug)

0.35 26.7

TOTAL

1. INSD-2650-29 11-33


2.

Predicted Performance The nominal performance of Nimbus I11 type 3P/2N RTG ' s e m p l o y i n g a

625 watt IIHS (or IRHS) is shown in Fig. 11-15 for various mission operating periods.

The nominal power output at end-of-mission (one year

after launch) is 26.6 watts at 2.4 volts. Hot and cold junction operating temperatures are 938OF and 390°F, respectively, at end-of-mission, decreasing from initial 9 78'3'

and 415OF values at beginning-of-life.

Figure 11-16shows the power and temperature profiles for the nominal helium release rate

(loo$)

and argon leak rate (3 x

scc/sec) values.

E

The performance is essentially identical to that predicted for the SNAP 19 N i m b u s 111

application where the RTG operates at a fin root temperature

of 354OF with a

625 watt fuel inventory.

Table 11-6 summarizes the nominal performance parameters for a 3P/2N RTG.

4

As is evident, the 26.6 watts obtained at end-of-mission is

significantly less than the 40.7 watts available from the TAGS-85/2N RTG previously discussed. However, as is shown in ChapterVI , approximately $0

watts at a

3P/2N

96% reliability is available at end-of-mission from two

RTG's connected in parallel.

T h u s , t h e 3P/2N

units may be satis-

factory for the RTG/battery hybrid power supply concept which requires approximately 25 watts from each of two RTG's. Additional power may be obtained from a 3P/2Ngenerator by increasing the fuel inventory and decreasing the operating fin root temperature by increasing the fin area.

This evaluation, d i s c u s s e d i n d e t a i l i n

Section 111-A, shows that nominal beginning-of-life power outputs up to

INSD -2650-29 I1-34

c


P

D

36 watts are feasible. Thus, a

3P/2N RTG cannot be satisfactorily

utilized for the present Martin Denver all-RTG system which requires

36 watts minimum per RTG at end-of-mission.

e Is

a

b


Q

INSD -2650-29 11-36


Y

a 1

4 1 11-37

INSD -2650-29


P

TABLE

PERFORMANCE

11-6

SUMMARY FOR 3 ~ / RTG 2 ~ DESIGN (INITIAL ARGON FILL)

BOL'

EO&

617

N o m i n a l input power (watts) Nominal output power (watts)

30.3

26.6 (24.8)3

RTG efficiency ($)

4.85

4.32

RTG specific power (watts/lb)

1.13

1.00

2.4 (3.6)

2.4 (3.6)

12.6 (8.4)

11.1 (7.4)

H o t junction temperature (OF)

978

938

Cold junction temperature (OF)

415

390

Fin root temperature

350

348

14.7

5.0

0

6.1

4

Load voltage (volts) Load current (amps)

4

(OF)

Total internal pressure (psia)5 5

Helium concentration (%)

Beginning-of-life, defined as at fueling nine months prior to launch. 2End-of -mission, defined as one year after launch.

3 Output power at EOM for a 0.96 reliability shown in parentheses. Values in parentheses are for two series-parallel strings.

5 Based on lo@ He release from capsule and 3 x

.

scc/sec seal

tightness (see Section 111-C)

.

I

INSD-2650-29

11-38 .

. . . .-.


111. PERFOFNA.NcE STUDIES ~

A.

~~

RTG OFTIMIZATION

This section presents the results of RTG optimization studies performed for TAGS-85/2N and 3P/2N RTG concepts for the Viking Lander mission.

In the studies, the radiator fin root operating temperature

was varied and the resultant effect on RTG fuel inventory, specific power, electrical power output and weight determined for a fixed hot junction temperature. inch and a longer present

Two generator housing lengths, the present

9.5 inch housing, were evaluated. Although the

8.9 inch long housing

housing length to

8.9

w i l l accommodate the IIHS, increasing the

9.5 inches is proposed to increase the thermal in-

sulation thickness at the ends of the heat source.

1.

Radiator Parameters The RTG radiator consists of the outer cylindrical housing, end

covers and six fins spaced at 60-degree increments around the housing. -7 i

The radiator emissive coating is a zirconium oxide/sodium silicate (binder) mixture developed on the SNAP 9A program.

Measured coating

properties show the thermal (infrared) emissivity to be the solar absorptivity to be

- 0.2.

- 0.83 and

Telemetered SNAP 9A S/N

04 RTG

housing temperature data received from Transit Satellite 1963-49B provide evidence of the excellent stability of this coating after approximately six years of space operation. Figure 111-1 shows radiator temperature data for the dirst three years of operation. APL reports no apparent degradation through the present.

INSD-2650-29 111-1


I

300 280 Outer Skin 260 Temperature (OF) 240 220

D

J

1963 1964

F

M

A

M

b I

J

A

S

O

N

D

J

1965

F

M

A

M

J

J

A

S

O

N

D

J 1966

FIG. 111-1. SNAP gA S/N

INSD-2650-29 111-2

F

M

A

M

J

04 ORBITAL RADIATOR

J

A

S

O

N

D

J 1967

TEMPERATURE STABILITY DATA


3 For the V king RTG, heat rejection was assumed to occ' r from the

3 3

finned cylindrical housing and one end cover.

s

a conservative maximum anticipated value for Martian daylight operation

3 rl

9 J I

J 1

a a

The end cover adjacent

to any mounting structure after installation was assumed to be adiabatic. The effective environmentalsink temperature was assumed to be 100°F,

based on Martin Denver proposal studies. Figures 111-2 through 111-4 show typical heat rejection capability curves for the finned cylindrical housing where Q/L is the heat rejected per inch of housing (fin) height (excludes heat rejected from end cover). 2.

TAGS-85/2N RTG The results of the optimization study for the TAGS-85/2N RTG are

shown in Table 111-1 and Fig. 111-5. A BOL design hot junction temperature of 1050°F was selected based on existing thermoelectric data (see Section 111-B).

Thermoelectric element size was maintained constant at

the present TAGS-85/2N values (P-leg 0.270 in. dia. x 0.500 lg., N-leg

0.377 in. dia. x

0.500 lg.) for the engineering models presently on

test at Isotopes. Specific power is seen to be rather insensitive over the fin root temperature range considered (325OF to 4OO0F), being approximately 1.6 watts per pound.

These results show that the trade-off is

essentially one of output power for weight and depends on the particular spacecraft requirements. The intact impact heat source (IIHS) proposed can easily accommodate the fuel inventories required for any of the cases whom.

3

I

Considering the power requirements of the Viking mission for a four

RTG system (minimum of 36 watts per RTG at end-of-missionat a high

INSD-2650-29 111-3


c

6 - -

.

- -

-.

.

.

INSD- 2650 - 29

111- 4

..

.-


-.-

.

J 1

..

i t

I'

I

3

a J J J

a 7 J

i I

'

!

. .

. , ..

i ...

.

INSD- 2650- 29 111- 5

I


7-7-

c

INSD - 2 650 - 29

111- 6


TABLE 111-1

TAGS-85/2N RTG BOL PARAME!TRIC EVALUATION DATA ( E f f e c t i v e Sink Temp. = 100°F, 9.5 inch housing l e n g t h ) Fin Root Cold Junction Tempeorature Temperature

F)

(OF)

Hot Junction Temperature

OF)

Net RTG Fuel Output , RTG Specific Fin S i z e ( i n . ) InvenBory Power Weight Power Fin Root Fin Root (watts) (watts) (lbs) watts/lb to-Tip Thickness

375

430

1050

634

38.9

24.0

1.62

5.6

0.1

350

405

1050

660

42.1

25.8

1.63

7.3

0.2

330

390

1050

675

44.1

28.2

1.56

9.0

0.3

325

385

1050

680

44.7

28.7

1.5.5

9.9

0.3


&

INSD-2650-29 111- 8

E


a 3 I 3 3 J 3 J

reliability--see Chapter VI), a nominal fin root operating temperature of 3304lwas selected. As is shown in Chapter VI this design condition, with a

3 3 9 J

power output, results in a nominal power output of

40.7 watts (38.5 watts parallel)

8 b EOM

at a

0.99 reliability for four units connected in

(one year after launch).

inventory is

675 watts,

and

36 watt minimum

9.0 inch wide fins (root-to-tip)with an 0.3

root thickness are required for the proposed

9.5 inch housing length.

These fins are approximately 1.5 wider than the present Nimbus I11 fins

(

7.5 inches long with 0.3 inch fin root).

Stress analyses performed

have shown that the 1.5 inch fin length increase w i l l not result in a structural dynamics problem for the anticipated Viking environment.

3.

3 ~ / RTG 2~ Table 111-2 and Fig.

111-6present the study results for a 3P/2N

For these studies the BOL hot junction temperature was maintained

at 1000°F.

The present SNAP 19 3P/2N thermoelectric element dimensions,

0.37" inch diameter by

0.500 inch long, were retained, adid the fin root

and fuel inventory varied so as to maintain the design 1000% hot junction temperature selected after examining the thermoelectric data prei sented in Section 111-B. BOL power outputs approaching

36 watts appear

practical for a 3P/2N RTG with an increase in fuel inventory and fin width and root thickness over the present Nimbus I11 configuration. Specific power is rather insensitive at approximately 1.2 watts per pound. for a

Comparison of the 3P/2N RTG with the TAGS-85/2N unit shows that

675 watt thermal input, power output is 35.3 watts and 44.1 watts

at BOL respectively.

3

Thus, the

requirement is satisfied and a 2.5 watt margin provided. The BOL fuel

RTG.

d

44.1 watt BOL

INSD- 2650- 29

111-9


TABU 111-2

3P/2N RTG BOL PARAMETRIC EVALUATION DATA (Effective Sink Temp. = 100°F, 9.5 inch housing length) P

Net RTG Fin Root Cold Junction Hot Junction Fuel Output IiTG Specific rature Temperature Temperature Inventory Power Weight Power (watts) (watts) ( l b s ) watts/lb (OF) ( OF)

*FF) HIU

Fin Size (in.) Fin Root Fin Root to-Tip Thickness

375

430

1000

627

30.5

24.2

1.26

5.7

0.1

350

4O5=

1000

655

33.2

25.9

1.28

7.3

0.2

330

390

1000

671

34.9= 28.5

1.22

9-2

0.3

325

385

1000

677

35.5

1.18

9.0

0.4

30.0


J

3

3

a a 3

a 3

a

3

INSD-2650-29 111-11


c B. 1.

Tâ‚ŹERMâ‚ŹlEXECTRIC SUEOR'I'ING U T A

Property Data In the past several years, Isotopes has developed a new P-type

thermoelectric material,

TAGS-85,composed of silver, antimony, ger-

manium and tellurium in the form (AgSbTe material operates in the general temperature range of lead telluride with a significantly higher energy conversion efficiency. The superiority of this material, as evidenced by the test data discussed in the following sections, justifies its use for the Viking mission. The rationale that supports this recommendation of TAGS is presented by

first examining initial performance and then the longer term behavior. A s might be expected, the largest body of evidence illustrating the

superiority of TAGS has been obtained from initial performance data. It is clearly evident from both analytical and experimental investigations that the initid and long term characteristics of TAGS are unequalled on an efficiency and output power basis. Three principal physical properties are indicators of thermoelectric performance: Seebeck coefficient, a ; electrical resistivity, p ; and thermal conductivity, k.

From these properties, a conclusion can be

dram on the potential worth of a material by calculating two performance

&

indices: figure-of-merit, Z = pk ; and the ZT product (where T = temperature).

Over the same temperature range, the material with the

greater figure-of-meritw i l l have the superior performance.

The ZT

index is of particular interest in that, as m rough approximation,

IIX-12


,-. I

i3

thermoelectric efficiency is related to it by the formula (ZT) (TH - TC)

where

' ~ T / E= thermoelectric efficiency

TH

J 3 J 3 J

TC

= hot junction temperature = cold junction temperature

Figures 111-7 through 111-ll present a, p , k, Z and ZT as a function of temperature for the materials PbSnTe (3P), PbTe (2P), SnTe and TAGS-85. At all temperatures, and even after an operational time in excess of khat typical of storage and the Viking mission (-1 year after launch) the ZT factor is highest for TAGS. Figures 111-12 and 111-13 show the change in Seebeck coefficient and

electrical resistivity with time obtained from greater than 10,000 hours of testing.

Seebeck and resistivity vary linearly on a logarithmic time

scale. These data are the fundamental thermoelectric property data used in performance prediction for long term operation. 2.

li

TAGS Performance Data Current hardware efforts utilizing TAGS thermoelectric material con-

tinue to demonstrate the superior performance achievable with this material over current lead telluride P-type materials.

A l l of Isotopes'

current RTG program have changed to TAGS as the P-leg material. The

SNAP 19 Nimbus I11 design has been modified from 3P to TAGS on three test generators oriented towards future applications such as Viking. SNAP 29 has fabricated sufficient TAGS-85/2N couples to build

1~~2650-29 111-13

8 modules,


160

140

120

. h

@ 100

>

I

Y

+.’

c

.2 0

80

.d

cr W 0

U

5

60

a,

P

a, W LI1

40

20

1

0

0

2 00

I

I

4 00

6 00

I 1200

I

1

1000

800

Temperature (OF)

FIG.

111-7.

INITIAL SEEBECK C O E F F I C I E N T

?+a

WSD-2650-26 111- 14 ,, ., . .-

...

..

.

..

.

~

....

... .

.. ..

. . .-..


a a

/

/

/

3P

/ /

/ /

J J 1

/ / / h

/

d

/

.d

I

2

/

1200

/

v

/ /

/

/

3

3 1

a J

/

/ /

/ / / 1

/

600

/

/

4 00

2oo

0

/

/

/

/

/ TAGS-8 5

t

0 0

3 FIG.

200

111-8.

400 6 00 Temperature

800

1000

(OF)

INITIAL EiXCTRICAL FC3SISTIVITY

INSD- 2650- 29

111-15

1200


2.5

-

2.0

-

1.5

-

L E

n

!3 & L G

i

\

3

8

v

h

s

--

TAGS-85

3P

s +-.

.A

a5 1.0 c 0

4

u

E

/

<--Yo , I

\

--4-* -22-

@ H

0

+ /

I

rl

cd

0.5

9

-

I

I

200

I

400

FIG. 111-9

I

800

600

I

1000

I N I T I A L THERMAL CONDUCTIVITY

INSD- 2650- 29 111-16 iâ&#x20AC;&#x2122;

I

1200


h

cr) I

0 I+

4

i3

3 3

J 1

3

-

x 1.0

/'

I

h v

4 .s

0.80

-

0.60

-

Lu

0

a, k

G /

0.40

0.20

,/-/

/ /

.'

/

0

'\

\ \

/'

\

\ \

0

'3P

0 0/

I

1

I

1

I

FIG. 111-10. INITIAL FIGURE-OF-MERIT

d

3

\

/

d 4

0

/

/

-

0

TAGS-85

INSD-2650-29 111-17

I


1.60

E

1.40

'TAGS%S / 1.20 c,

0

3

a 0

cl: Q

1.00

k 3

/

c,

cd

k Q

I$ 0.80 Q,

7

h

I

/

L

3 k

$

0.60

W

0

a,

2 0.40 0.20

0 0

2 00

400

600

800

1000

1200

Temperature (OF)

E FIG. 111-11. IIIITIAL FIGUFZ-OF-~RIT-TEIG'ERATURE PRODUCT

INSD- 2650- 29

111-18


0

(D

0 I

01

0 1 I

0

c 0 1

c.r

w 0

S e e b e c k Coefficient (vv/"F) c N I

0

1

c 0 0


I

I 0

0 Q,

1

0 0 03

I

1 W

0 0

c-

0 0

INSD-2650-29 111-20

I

v)

0 0

I

*

0 0

0

I

El

/-

J


.

employing w e l l over 100 watts each with t o r and

4 to

be l i f e t e s t e d .

4

.

.

.... .... .

.-.._-

t o be d e l i v e r e d as a genera-

The S e n t i n e l t e r r e s t r i a l RTG program has

d e l i v e r e d 13 modules u t i l i z i n g TAGS, a l l of which are s u c c e s s f u l l y operating i n t h e f i e l d .

i3

a.

SNAP

The SNAP

19

19

generator design which u n t i l r e c e n t l y used 3P/2N

couples f o r energy conversion, now incorporates TAGS-85/2N couples.

G3 il 1 3 3

This change has been i n i t i a t e d because of t h e s u p e r i o r thermoelectric e f f i c i e n c y over 3P over t h e temperature range of i n t e r e s t , roughly 300% t o 1000%.

Generators S/N 24 and 25, which were f a b r i c a t e d l a t e

i n 1968 f o r Nimbus 111, u t i l i z e d 3P/2N couples.

Generators 26, 27 and

28 have very r e c e n t l y been b u i l t with TAGS-85/2N couples and are d i s cussed below. (1) Couple power check

During t h e f a b r i c a t i o n of couples f o r use i n generators, sampling techniques are employed f o r q u a l i t y c o n t r o l of t h e process. The most meaningful performance i n s p e c t i o n i s t h a t provided by statistically selected couple power checks.

checked a t 1050"F hot junction,

J 3

13

matched load.

400%

~ ~ c s - 8 5 /couples 2~ a r e power cold junction temperatures and

F r o m t h e s e couple power checks, t h e subsequent generator

power may be a c c u r a t e l y p r e d i c t e d after s u i t a b l e c o r r e c t i o n s a r e made. These c o r r e c t i o n s t a k e i n t o consideration t h e a d d i t i o n a l connections and wiring i n t h e c i r c u i t , t h e f a c t t h a t a generator i s operated a t o p t i -

mum e f f i c i e n c y l l o a d and n o t matched load, and operating temperature d i f f e r e n c e s , i f any.

L W I YI

IVLI Y I I /

INSD-2650-29 111-21

\L


A s of t h i s r e p o r t , TAGS-85/2N couples hav

b en f a b r i c a t e d on

SNAP 19 from Lot 01 ( q u a l i f i c a t i o n l o t ) and Lots 02 through 05 (production l o t s ) .

These production l o t couples, a s well as Lot 06 f o r

which no power check d a t a are y e t a v a i l a b l e , have been used i n gene r a t o r s 26, 27 and 28.

A summary of t h e i r performance, based on t h e

d e s t r u c t i v e power checks, i s given i n Table 111-3. Based on t h e r e s u l t s of a comprehensive computer program,

a power p r e d i c t i o n of t h e couple w a s made and i s shown i n Fig.111-14 Noted on t h a t f i g u r e are t h e voltages corresponding t o m a x i m u m power

E E L

(when e x t e r n a l r e s i s t a n c e matches i n t e r n a l r e s i s t a n c e , hence t h e t e r m matched l o a d ) and maximum e f f i c i e n c y (always a t a higher voltage than maximum power).

Values of predicted open c i r c u i t voltage and power

are a l s o l i s t e d on Table 111-3 f o r comparison with t e s t r e s u l t s .

The

c o r r e l a t i o n i s good and thus lends credence t o other p r e d i c t i o n s a t t h e generator l e v e l . Note from Fig.

111-14 t h a t a t maximum e f f i c i e n c y load (.095

v o l t s ) , t h e couple power i s 0.52 w a t t s a t lO5WF/bOoF.

With 90

couples i n a SNAP generator, t h e gross output power (before s t r a p and

w i r e loss cokrections) would be 46.8 watts while t h e R E gross load voltage would be 2.85 v o l t s .

These values a r e i n e x c e l l e n t agreement

with t h e r e s u l t s observed on generators 26, 27 and 28. ( 2 ) Generators S/N 26, 27 and 28 Generators S/N 26, 27 and 28 were t h e f i r s t SNAP 19's b u i l t with TAGS-85/2N thermoelectric couples.

They d i f f e r from t h e Nimbus

#

INSD- 2650- 29 111-22

E


TABLE 111-3 SNAP 19 Couple Power T e s t Summary

(Th = lO’jO%;

H

N-Leg Output Power

P-Leg Output Power

Couple Output Power

(watts)

(watts)

(watts)

N-Leg open Circuit Voltage (volts)

P-Leg Open Circuit Voltqe (volts)

Couple Open Circuit Voltage (volts)

&ant i$y

Load

Lot No.

Lot Size

Power Checked

Voltage (volts)

01

18

5

077

319

,256

572

093

,062

.154

02

110

5

on

.311

.245

552

092

,061

153

03

68

5 -

.076

,316

.243

557

092

.061

153

04

39

4

.076

.318

.246

557

092

.060

.151

05

92

4

.076

.316

.245

553

.091

.061

.152

- H

T, = 400%)

H

Iu

w

,

*

0

Averqe (Test) Standasd Deviation ( T e s t )

Aver age Predicted

.061

153

.002

.002

.002

.og1

.064

155

.316

248

558

092

0I.l

,013

.017

0297

,250

545

’.


t #

. ., .

..

_ / . . _

-

.

,

.

. .

.. j

I

INSD- 2 6 5 0 - 29 111- 24

C E


I11 3P/2N configuration generators in that the 0.500" long and

.377"

diameter 3P elements have been replaced with 0.400" long and .270" diameter

TAGS-85elements, segmented with a .loo" thick SnTe segment

at the hot end.

il

Generators S/N

26, 27 and 28 are to be operated at hot

and cold junction temperatures representative of the Viking mission. Power checks af S/N

26 at Power inputs of 625 watts and 690 watts to

the heaters resulted in power outputs of

1 J 1 J -l

spectively.

Power checks of S/N

resulted in power outputs of

36.6 and 43.0 watts, re-

27 and 28 at TOO watts to the heaters

44.1 and 43.3, respectively. (Extra

thermal losses in an electrically heated generator at the higher power inputs due to reduced insulation for vibration capability amount to

25 watts thermal.)

Required power input for equivalent power outputs

of a fueled unit w i l l therefore be lower by 25 watts thermal.)

Results

of these tests are illustrated in Table 111-4. A parametric test in load voltage was completed on each generator (Figs. 111-15, 111-16 and 111-17). S/N and S/N

26 was tested after 340 hours

27 and 28 were tested after the initial power checks and in-

stallation of pressure transducers.

The peak power parameters a r e

recorded in Table 111-5 for comparison between generators. The thermoelectric material properties change exponentially with

a 1

time, with most of the change occurring early in life, as evidenced on Fig. 111-18. A greater change occurs from 0 to 3 months than from to 9 months.

Comparisons on S/N

3

26 between the initial power check

results and data fromthe parametric test indicate a burn-in during the

INSD-2650-29 111-25


RESULTS OF POWER CHECKS ON TAGS-85/2NGENEFUTORS Generator

Generator S/N 26 measured normalized*

Generator Generator Prediction S/N 27 S / N 28 (peak power p o i n t )

Parameter

S/N 26

Date of power check

5/8/69

5/9/69

5/9/69

5/29/69 6/1/69

--

Heater power (equiv. f'ueled)

watts

600

665

675

675

675

675

Power output

watts

36.6

43.0

44.0

44.1

43.3

44.1

$

6.1

6.5

-6.5

6.5

6.4

6.5

Efflciency (overall)

Load c u r r e n t

amps

14.0

16.2

16.2

15.5

15.3

15.4

Load voltage

volts

2.61

2.66

2.72

2.84

2.83

2.85

Open c i r c u i t voltage

volts

4.26

4.59

4.62

4.62

4.59

4.74

temperature

OF

987

1038

1048

1021

1016

1054

Cold j u n c t i o n temperature

OF

395

404

408

361

360

390'

Fin r o o t temperature

OF

339

349

353

36

304

--

Generator argon pressure

psia

14

14

14

14

14

14

6

17

17

9

6

0

Hot j u n c t i o n

ROUS

of o p e r a t i o n &

* Excluding outgassing time e Normalized t o 675 watts h e a t e r input


1

a a a

3

fi INSD- 2650-29 111- 2 7


E

A-

c

-

-

.

- -

--

-

INSD- 2650-29 III- 28


" I

a I J

IN=- 2650-29 111- 29


TABLE 111-5

SUMMARY OF PEAK POWER POINTS OF PARAMETRIC TESTS Generator

Generator

S/M 26

S/N 27

Heater power (equiv. f u e l e d ) watts

675

675

675

675

Power output

watts

43.5

43.1

42.1

42.9

E f f i c i e n c y (o v e r a l l )

Q

6.4

6.4

6.2

6.4

Load c u r r e q t

=PS

15.4

15.4

15.0

15.3

Load voltage

volts

2.88

2.79

2.81

2.83

Open c i r c u i t voltage

volts

4.91

4.63

4.58

4.71

Hot j u n c t i o n temperature

?I?

1060

1026

1030

1039

Cold j u n c t i o n temperature

9

396

380

382

386

Fin r o o t temperature

OF

336

331

337

335

Hours of o p e r a t i o n

hrS

340

60

60

Y..

Generator argon p r e s s u r e

psia

14

-29

- 29

Parameter

H

H ‘H IH 81

(W

Generator S/N 28 ‘

Average

!O

...-


R

a a

111- 3 1

INSD-2650-29


i n i t i a l 340 hours on t h e order of 0.5 w a t t s . A @ e a t e r change was observed between i n i t i a l power check and parametric t e s t over a s h o r t e r period (-60 hours) on S/N 28.

27

and S/N

The d i f f e r e n c e was an average of 1.1watts and i s a t t r i b u t a b l e

t o a change i n generator argon pressure from 29 p s i a (parametric t e s t ) .

14 p s i a (power check) t o

This pressure i n c r e a s e causes g r e a t e r

thermal l o s s e s through t h e MinK i n s u l a t i o n .

The i n c r e a s e i n i n t e r n a l

gas pressure was accomplished during i n s t a l l a t i o n of t h e pressure transducer, which was performed after t h e power check but before t h e parametric t e s t . The average power output f r o m the f i r s t power checks of t h r e e

r e p r e s e n t a t i v e TAGS-85/2N generators b u i l t t o d a t e i s 43.8 watts f o r t h e 675 w a t t ( e q u i v a l e n t f u e l e d ) thermal inventory.

Generator

S/N 26

i s being maintained near i n i t i a l optimum load v o l t a g e (2.86) w i t h 675

w a t t s ( e q u i v a l e n t f u e l e d ) e l e c t r i c a l i n p u t and a f i n r o o t temperature of 325OF.

S t a b i l i t y of s i g n i f i c a n t performance parameters w i l l con-

t i n u a l l y be monitored as time accrues.

On June 1 2 and 13, generators S/N

27

and 28 were s u c c e s s f u l l y

hard-mount v i b r a t e d on l o a d i n t h e heated condition t o t h e l e v e l s shown i n Table 111-6. Vibration was performed i n all t h r e e axes with s i n u s o i d a l i n p u t f i r s t , followed by random input, i n each a x i s .

Out-

put power was unchanged as a r e s u l t of t h e t e s t s , b. R

SNAP 29

Four SNAP 29 90-couple T A G S - ~ ~ /modules ~N (S/N 001, 002, 007, arid 008) have been f a b r i c a t e d and placed on l i f e t e s t .

IEBSD-2650-29

Four

E


TABLE 111-6

is

VIBRATION LEVELS FOR GENERATORS S/N 27 AND 28 Sinus oi dsl Frequency Range Direction

P

- 500 500 - 2000

z (axial) axis

15

20

x, y ( l a t e r a l axes)

20

500

10

- 500 -

Amplitude 0 t o peak)

(g's,

(CPS)

10

2000

7

Vibration l i m i t e d t o 0.25 inch s i n g l e amplitude f .025 inch Sweep r a t e : 2 octaves/minute Single sweep per axis

J I

a a

Random

Frequency Range Direction

kPS>

Two minutes i n each axis; t o t a l ,

INSD-2650-29

111-33

Level k2/CPS)

6

minutes


additional modules are being fabricated and will be delivered as an electrically heated thermoelectric generator (ETG) system. module performance on SNAP (1)

Couple and

29 are discussed below.

Couple power check

Sampling techniques similar to those employed on SNAP 19 are Qsed on SNAP 29 for quality assurance. Of this sampling, the most significant is the couple power check.

The couple performance pre-

dicted by computer analysis for SNAP 29 couples operating at Th = 1050°F and T, = 35OoF is presented in Fig. 111-19. The SNAP 29 Program has fabricated 940 TAGS-85/2N couples, of which matched l o a d .

c

85 were power checked at

A compilation of data f r o m these power checks is shown

in Table 111-7. Comparison of the analytical and test results illus; trates the excellent correlation obtainable. Good correlation between couple power checks and module performance is also obtained by subtracting the predicted strap and wire losses from the product of the average couple power and the number of couples.. The advantage of a TAGS-85/2N couple over the 3P/2N couple is indicated in Table 111-8. Over the SNAP 29 temperature range(1O5O0F/ 35OoF), the TAGS couples have a

24% better

efficiency and

output power per unit couple area than the 3P couple.

45% higher

The difference

is almost as great over the SNAP 19 range of temperature (1050DF/4000F). It should be noted that the SNAP 29 couple is not of optimum geometry.

The P-leg was sized with the existing

0.560 inch diameter

N-leg to satisfy module power requirementslmand thus could not be

i

reduced in diameter to the optimum efficiency geometry.

INSD-2650- 29 111-34

I


5 E: -111 62-0592 -(ISNI

Q


TABLE

SNAP

111-7

29 COUPLE POWER TEST SUMMARY

( Th = 1 0 5 O O F ; Tc =

Lot No.

Lot

- Size

,

Quantity Power Checked

Load Voltage (volts)

N-Leg Output Power (watts)

P-Leg Output Power (watts)

Couple Output Power (watts)

N-Leg open Circuit Voltage (volts)

.080

790

895

1.58

.og6

,065

.161

777

.886

1.56

097

.065

.162

.799

.g11

1.60

.098

.65

.163

I

20

60

40

,

21

40

3

.080'-'+

22

64

3

.080

23

80

3

24

80

6

.080

25

80

4

.080

26

80

3

.080

27

64

3

.om i

28

80

3

.080

I

I

I

350%) P-Leg open Circuit Voltage (volts)

Couple open Circuit Voltage (volts)

I

*

~

I I

v

29

80

4

.om-

30

80

3

.081

31

20

4

.083 . .

si

i

b


TABLE 111-7 ( Continued)

Lot No.

Lot - Size

32

Q u a n t i t y Load Voltage, Power Checked (volts)

56

33 Total

940

N-Leg

P-Leg

Couple

N-Leg Output Power (watts)

P-Leg Output Power (watts)

Couple Output Power (watts)

Copen ircuit Voltage (volts)

Copen ircuit Voltage (volts)

Copen ircuit Voltage (volts)

3

.080

*&5

.966

1.65

099

.068

.167

3

.080

.813

931

1.64

09

0067

.I64

.om

.788

898

1.58

097

.065

.lo2

.O27

.044

.002

.002

.002

729

.889

0%

.068

.164

85

Average (Test)

Standard Deviation (Test)

Average (Predicted)

047

1.62

C W A R I S O N OF OF’I’IMIZED TAGs-85/2N AND 3P/2N COUPLES AT BEGINNING-OF-LIFE Output Power per U

T/E E f f i c i e n c y

Hot/Cold Junction

2t Couple Area 1

Temperatures ( 9)

3P/2N

(73 TAGS-85/2N

9 Improvement

3P/2N

TAGS-85/2N

1050/300

7-31

9-12

24.8

2.91

4.26

46.4

1050/350

6.87

8.53

24.2

2.53

3.68

45.5

1050/400

6.41

7.91

23.4

2.17

3.14

44.7

(din

Improvement


(2)

Couple thermal cycle and life'tests

Nineteen SNAP 29 TAGS-85/2N couples have been life tested up to 2700 hours at a cold junction temperature of

375OF and hot junction

temperatures of 1025, 1050, and 1075OF. These couples are being operated at matched load conditions.

P

Five of the couples underwent

three deep thermal cycles and all of the couples received a two-hour power check prior to the life test. Analytical predictions for beginning-of-life and 2160 hours (3 months) are illustrated in Fig. 111-20 for the three temperatures of interest

(1075, 1050, and lO25'F).

Ex-

perimental power-time maps of the individual legs and couples for each of the hot junction temperatures are presented as Figs. 111-21 to 111-26.

In each case the mean power is indicated. In general a burn-in was experienced during the first 1000 hours, after which time virtually stable power was observed.

Couple life test results are summarized in Table

111-9. On all nineteen couples, the internal resistance and open circuit voltage of the TAGS leg increased, thus verifying the expected increase with time of the electrical resistivity and the Seebeck voltage as previously discussed.

The five couples which underwent thermal cycling

prior to the life test experienced slightly higher power reductions. Correlation between predicted and measured endurance test data is good as illustrated in Table 111-10.

( 3)

Large TAGS module

A 90-couple 2N-TAGS module of an earlier design was endurance tested at Th = lOl5OF and Tc = 35OoF for about 4200 hours.

2650-29 111-38

INSD-

The test


a

3

INSD-2650-29 111- 39


III- 40

INSD-2650-29

/-

L s


III-41

INSD-2650-29

t

..

t 4

-I

-1

i

-7

I -I


m-44

INS-2650- 29

.

__ ..

I

.. .. .{

. . ._. . .

E


0 Q 3

a 1

a II

a

Jli

I I

1 .I- __

._. ..

INSD- 2650-29 111- 45

-

).

.I...

I

.

2

!

. . . . . . . . . . .

..


INSD-2650-29 111-4 6

M (u

* * d ri

ci

L-

*

0;

#

f0

E

r 0 c


TABLE 111-10

OF PREDICTED AND TEST RESULTS OF SNAP 29

COMPARISON

Predicted

* Note

peak voltages.

LIFE TESTS

*

Average Measured

N-Leg output Power

P-Leg output Power

Couple output Power

(watts)

(watts)

(watts)

that PN + P

corn

N-Leg output . Power

P-Leg output

(watts)

Power (watts)

Couple output Power (watts)

69

77

1.31

67

9

75

1.42

70

.81

1.51

9

78

-92

1.52

70

89

1.58

.80

98

1.58

63

72

1.35

.62

75

1.22

.66

78

1.43

71

83

1.42

69

83

1.52

71

87

1.42

P

#

PCm

9

s i n c e individual leg powers are a t t h e i r r e s p e c t i v e

'


was interrupted by a test loop failure. Operational performance of the module, known as the large TAGS module (LTM), is illustrated on Fig. III-27.

This module actually appreciated 2.4% in power over the 4200

hours tested.

(4)

Module life tests

Four SNAP 29 TAGS-85/2pJ modules have been placed on life tests. Two of these modules, S/N 001 and S/N 002 have, at this writing, com-

pleted over 1500 hours while the other two modules, S/N 007 and S/N

008, have just completed the initial power check. These 90-couple modules are being tested at

= 1050°F and

Tc = 375OF.

Modules S/N 001 and 002

developed leaks early in the testing (following thermal cycling); were purged with argon after -S/N 001 experienced an unexpected power reduction, prob?bly due to oxidation of the thermoelectrics. The power level of

S/N 001 appreciated after the argon purge was initiated, and the power l o s s after

1880 hours is 5.8%. The severity of the leak in S/N

002

was not s o great as S/N 001, and the power loss is minimal, about 1.7% after 1640 hours. Fig. 111-28.

The performance of both modules is illustrated in

The trend is toward stability.

The initial power check of modules S/N 007 and 008 indicate initial power outputs of 117.1 and

117.4 watts, respectively. Average

initial power output of the Iffour life test modules is 116.5 watts, which compares well with the predicted output'of 114 watts. c.

Sentinel

The Sentinel commercial RTG program has fabricated and operated

.- .-

-.

INSD-2650-29 111-lr8

. . ..

._


................

......................

w

INSD- 2650 - 29 111- 49

- . -.

.

. . . . . . . .


INSD- 2 6 5 0 - 29

111- 50

.


23 thermoelectric conversion assemblies (13 for the field and 10

for the laboratory) which utilize series-connected TAGS/Isotopes N couples.

The TAGS assemblies have approximated, 190,000 hours (as of

6/15/69) without failure, These systems have a 5-year predicted reliability of

0.977 with a value of 0.85 or greater at a

50% confidence

level being demonstrated to date. A l l of the Sentinel RW's utilize an earlier design for the couples.

An improved couple using exothermic bonding techniques is now being produced.

Figure 111-29 presents test results of the TAGS P-leg of a pre-

prototype exothermic bonded couple.

Over the 5500 hour test, the TAGS

demonstrated excellent stability after the initial burn-in period. d.

Long eerm %ouple 'tests

Couples of various geometries utilizing TAGS as the P-leg material have been life-tested as long as 21,000 hours.

The ratio of normalized

couple power output to initial couple power (P/Po) is shown as a function of time in Fig. 111-30.

The average ratio of P/Po is presented in

Table 111-11. As is evident, excellent long term power stability is demonstrated.

TABm 111-11

Time (hours)

No. of Couples

Average P/Po

0 5000 10000

6 6 5

0.99

20000

3

0.93

INSD-2650-29 111- 51

1.00 1.02


f

INSD- 2650 - 29

111- 5 2


3.

3 ~ / Performance 2 ~ Data

Data for evaluating the performance of 3P/2N thermoelectrics were primarily from the SNAP S/N

24 and S/N 25,

SNAP

19 S/N 20 generator and Nimbus I11 flight units

from SNAP 11 generator and module testing, and from

29 module testing. Currently, the SNAP 19 S/N 20 generator has

operated in excess of 11,000 hours, while two SNAP 11 generators (Q/N-

1M and Q/N-3) have each accumulated test times in excess of 18,000 hours. Additionally, two SNAP 11 modules operating at 900° and 10000F, respectively, have each accumulated operating time of 17,000 hours.

On the

SNAP 29 program, a total of five modules have been fabricated and test-

ed with a maximum operating time o f 7000 hours.

a.

SNAP

The SNAP

19 S/N 20 generator

19 S/N 20 generator is an experimental unit built to

evaluate 3P/2N material in a SNAP 19 configuration. Endurance testing was performed at a thermal input of

630 watts since shipment to

JPL.

570 watts while at Isotopes, and

The

operating hot junction temperature of the hot junction is

570 watt input resulted in an

..,900°F.

With a 630-watt input

~960'~.

E - I characteristics were determined initially for various thermal

power inputs (570,

630, and 700 watts) at a

34OoF fin root temperature.

The curves are shown in Fig. 111-31. The maximum power outputs for these power inputs were

26.0, 3 1 . 1 and 36.9 watts, respectively. Cor-

responding hot junction temperatures were 905O, respectively.

INSD-2650-29 111-54

955' and lOl5'F,

P


Initial p a r a m e t r i c s

1050 1000

v

;.E k a,

950 9 00

348

6

cI

0

wo

-

'

Y

I340

/ / &

- 5.0

--

----

h

m +

d

0

>

-e-

7 P o w e r In ( w a t t s ) 630 - 570 700

- - --_-

\

18

- 4.0

0

u

c

-

3.0

14

cd

+ c

5

12

u \ 10

/

)A&-

----\

\

\\

0 0

8

1.4

0

0

' '\

I

-\

1.6

2.0

2.4

\

\

I

I

1

2.8

3.2

3.6

Load Voltage (volts) FIG. 111-31.

SNAP 19 GENERATOR S J N 20

T1

INSD- 2650 - 29 111-55

;.

k

h

a, k

+

.-I

16

-

Y

20

&


Figure 111-32 i s a p l o t of t h e endurance d a t a during t e s t i n g a t Isotopes.

The i n i t i a l power output w a s 25.8 watts, decreasing t o

E

24.7 watts a t 5235 hours. I t i s notable from t h e data received from JPL, t h a t t h e 570-watt

peak power output, 25.4 watts ( F i g . 111-33), i s approximately t h a t measured a t Isotopes when generator t e s t i n g w a s discontinued (24.7

watts) at.-5300

accumulated operating hours.

As shown i n Fig.

111-33,

t h e measured power output a t 2.4 v o l t s and a 630-watt power input ( t h e same as t h a t for t h e Nimbus 111 generator) i s 29.9 watts. SNAP

b.

19

S / N 24 and S / N 25 generators

Figure 111-34 shows t h e results of acceptance t e s t i n g performed

on t h e Nimbus I11 fueled, f l i g h t RTG u n i t s S/N 24 and S/N 25.

E E

Testing

w a s performed i n a vacuum chamber i n which t h e chamber w a l l temperat u r e w a s adjusted t o y i e l d t h e design f i n r o o t temperature of -354OF. The peak power outputs a t acceptance f o r t h e two u n i t s were 31.4 and 29.6 watts for t h e S/N 24 and S/N 25, r e s p e c t i v e l y .

A t the writing

of t h i s r e p o r t , S/N 24 and S/N 25 have been i n e a r t h o r b i t over two months and produce t h e same e l e c t r i c a l power as j u s t a f t e r launch

( -25

watts per g e n e r a t o r ) . c.

SNAP 11 generators

Performance d a t a for two SNAP 11 3P/2N generators, Q/N-lM and

Q/N-3,

c u r r e n t l y on t e s t a t Sandia Corporation and JPL, r e s p e c t i v e l y ,

are presented i n Figs. 111-35 and 111-36.

Neither of t h e s e generators

was intended f o r endurance t e s t i n g , thus, d a t a comparison can be made only f o r instances when similar t e s t p o i n t s occurred during parametric testing

. <%.s

&'

INSD- 2650- 29 111- y6

I f


R 3

s 3

J 3 J 3

0.

a l 00

al 0)

0 0

0 0 0 0

0 0

0 0 0 0 0 0

0

0 0

0. 0. 0.

h

&&+I 0 0 0

I1

I

I1

I

II

0

met-

-

-.

..

.I

INSD- 2650- 29 111-5 7

.h7."W


c

Total a c c u m u l a t e d

test h o u r s * 8930

C

h Tot a1 a c c u m u l a t e d test h o u r s *

--------

6500 8546 9778

G 630-watt input

57 0 - w a t t input

*Total a c c u m u l a t e d o p e r a t i n g h o u r s i n c l u d e -5300 at I s o t o p e s

V o l t a g e (volts)

FIG. 111-33. R E S U L T S O F JPL S N A P 19 S/N 20 TESTING

INSD- 2 6 5 0 - 2 9 111- 58


1

Tcold

3

J

3 3 3 3

INSD-2650-29 111- 59

= 420" F


-

0 Raw d a t a 0 Normalized

Isotope s checkout

E

to 1000" F, 350%'

RTG defueled a n d air t e s t i n g continued at Sandia C o r p . 10,000 h o u r s

I:@*

8 0

O0

0 0 0000

oo 0 0

15,000 h o u r s

0

0

O O 000

RTG fueled and tested at ORNL at s i m u l a t e d l u n a r conditions ( H J = 900" t o 1000" F)

0

0

S

G

0 0

Hot junction t e m p e r a t u r e = 1000" F C o l d junction t e m p e r a t u r e = 350" F L o a d condition - m a t c h e d ( R loa = R inter rial1 C o n s t a n t p o w e r input = 815 f5 w a t t s E n v i r o n m e n t = air

I

3000

I

I

6000 9000 Operating T i m e ( h r )

I

12,000

I

15,000

I FIG. 111-35.

S N A P 1 1 GENERATOR Q/N-1M T E S T DATA 3 P / 2 N

I INSD-2650-29 111- 60


7/20/66

Isotopes test in g

e

1 / 2 5/ 6 7

Initiation of JPL testing

11/14/67

-- ----- -0----

27

Pin = 810 w a t t s , air

I

UI

$ 23

e

n

4

Pin = 715 w a t t s , s i m u l a t e d l u n a r night

v

a

t:

;.q

0

L

a"

(

Pin = 634 w a t t s , s i m u l a t e d l u n a r night

20

:I E

0

I

I

0

0

Pin = 590 w a t t s , s i m u l a t e d l u n a r night

20

1000

2000

I

I

3000

4000

1

5000

I

6000

/ I

'I

1

1

8000

9000

F I G . 111-36. S N A P 11 Q / N - 3 P E R F O R M A N C E SUMMARY-3P/2N GENERATOR

A~

A A A

Cold junction t e m p e r a t u r e = 400" F L o a d condition-matched (Rload = i n t e r n a l1 0 Box N o . 4, h o t junction t e m p e r a t u r e = 900" F A Box N o . 2 , h o t junction t e m p e r a t u r e = 1000" F

I

0

1

3000

I

I

9000 O p e r a t i n g Time ( h r )

6000

FIG. 111- 37. POWER VERSUS T I M E FOR S N A P 11

INSD- 2650 - 29 111- 6 1

I

12,000

I

15,000

MODULE BOXES-3P/2N


The thermoelectric elements i n t h e SNAP 11 generator a r e 0.346 inch i n diameter by 0.500 inch long; a generator contains a t o t a l o f 72 thermoelectric couples.

This compares with the 0.377 by 0.500 inch

long elements and 90 couples i n t h e 3P/2N SNAP 19 generator. The Q/N-lM generator w a s a demonstration u n i t which was fueled a t

ORNL with a curium-242 heat source and was subjected t o a 90-day simul a t e d lunar mission i n a vacuum environment.

After s u c c e s s f u l l y

completing t h e simulated mission, t h e generator w a s defueled and s e n t t o Sandia Corporation where it i s p r e s e n t l y on t e s t a t a hot junction operating temperature of

.d

1000°F

(815 w a t t s power i n p u t ) .

Total

operating time to date, including a l l previous Isotopes and ORNL t e s t i n g ,

i s g r e a t e r than 18,000 hours.

A s shown i n Fig.

power output has decreased from -25.8

111-35, t h e generator

t o 23.3 watts during t h e 15,000

hour t e s t time f o r which data have been published. The SNAP 11 Q/N-3 generator t e s t data are shown i n Fig. 111-36 f o r t h e f i r s t 9300 hours; a c t u a l t e s t time i s i n excess o f 18,000 hours. I n i t i a l a i r and vacuum t e s t i n g w a s performed by Isotopes; t h e generator vas $hen shipped t o JFL, where it has been subjected t o air, vacuum and dynamic t e s t i n g .

( A s t h i s generator i s undergoing a i r and vacuum t e s t i n g ,

only similar t e s t p o i n t s can be compared, thus explaining t h e i n f r e quency of data points i n t h e f i g u r e . )

Examination of t h e t e s t p o i n t s

shows t h a t a maximum decrease i n power of 0.5 w a t t has occurred i n t h e generator a f t e r 9300 hours of t e s t t i m e .

-50%

It should be noted t h a t during

of t h i s t e s t t i m e , t h e hot junction temperature operated i n t h e

WOOF range due t o t h e a i r parametric t e s t i n g performed, and a t tempera-

INSD- 2650- 29

111-62


.

tures in excess of 900'3' during the remaining 50% of the test time. Although data are presently not available beyond 9300 hours, communication with JPL has verified that generator performance has remained virtually unchanged during subsequent testing. Testing of this gen-

a

erator at JPL is continuing. d.

SNAP 11 modules

An investigation of the performance of 3P/2N thermoelectric couples

I

at h o t junction temperatures of 900'

and 1000°F was initiated on the

SNAP 11 program and later continued on the SNAP

19 program. Two modules

each containing six thermoelectric couples connected in series, were fabricated and tested. The N and P element size (0.377 in. diameter by

1

0.500 in. long) of these couples are the same as for the SNAP

19

generator. Figure 111-37 shows the results of module testing for the 3P/2N

3

materials at hot junction temperatures of 900'

and 1000°F.

In these

tests the hot junction is maintained at a constant temperature by controlling the power input. Initial power output for the 900°F module was 1.81 watts, decreasing to

1.71 watts after ~17,000hours. However,

power output remained virtually unchanged after the first 3000 hours

3

of operation.

A comparison of the results of these gOO°F data w a s made with SNAP

19 generator S/N 20. Based on the 5000-h0ur test data point from the 9G0°F module, the predicted power output of the SNAP 19 generator S/N 20 is 25.0 watts.

This result is in excellent agreement with the actual

S/N 2G generator normalized power output (at 900'

INSD- 2650- 29 111-63

and 400°F junction

,


D

Atemperatures), being 25.2 watts at

-u5000

I I

hours.

Initial power output for the 1000째F box was 2.60 watts, decreasing to 2.25 watts after

l 7 , O C O hours.

Again, power stability was demon-

strated after ,- 4000 hours of operation.

A comparison of the results of the 1000째F module data was made with SNAP 11 generator Q/N-lM.

c

Eased on the l5,OOO-hour test data point

from the 1000째F module data, the predicted power output of the SNAP 11 Q/N-lM generator is

E E

26.1. This is in excellent agreement with the Q/N-

1M data where the normalized power output (at 1000째 and 35OoF junction

E

temperatures) is 25.8 watts after w l 5 , O O O hours of testing. From these discussions, the ability to correlate generator and module data for 3P/2N thermoelectrics is evident. e.

SNAP

29 modules

The design hot and cold junction temperatures of the SNAP 29 3P/2N modular generator are 1050' and 35OoF, respectively.

Figure

111-38 shows the normalized (to lO5o0F hot junction; 3 5 0 ' ~ cold junction) power histories of typical SNAP 29 modules. Operating times up to 7000 hours have been accumulated, with power decreases on the order of 8% noted.

The data trend indicates the power output to be stabiliz-

ing in the 5000 to 6000 hour test period.

4.

Mechanical Properties In handling and working with TAGS elements, it is obvious that the

mechanical properties of the material are superior to the $-type lead tellurides. In order to compare the mechanical strength of the two

INSD-2650-29

111-6& -

..

.

. ., . ....

..

.. . ..-

. - ~


cc

c3

mv,

c3

v)

r m

c3

03

w 'j

(D

x

P

D D D D D D D

B

D

D D D D D D D D D D D D

D

D

D D D

D

D D D D D

o

D D

D

D

Power (watts) S I N 003

0

S / N 002*~

s9 -111 62 -059Z -CISNI

0

0 0 0 0

0 0

0

C 0 0

0 0 0 0 0 0 0 0 0 0 0 0 0 0

S / N 005


materials more quantitatively, several tests were performed.

Both cast

E

and hot pressed TAGS samples were used while cold pressed and sintered PbTe (purchased from 3M) was used.

G

Three types of tests, tensile tests

with couples and compression and drop tests with elements, were per-

h E E

formed. The tensile tests were performed by bonding metal hooks to the cold caps of the couple, restraining the shoe in a fixture and then applying a tensile force until the thermocouple leg was pulled apart. The drop tests were subdivided into two types; in one case elements were dropped onto

a steel bal1,and in the other case the ball was

dropped onto the element.

br

The compression test consisted of placing an

element between the platens of a hydraulic press and compressing it

E

until failure occurred. Table 111-12 shows the results of the tensile tests on couples.

r

A l l of the couples subjected to this test were operated 100 hours prior

to testing, The couples used in these tests were from production lots, and the two groups were run at different times.

The validity of the

test is indicated by the near duplication of the average value of the pressure required to pull off the N-leg of each lot. Of the twelve N-legs removed, eleven broke in the bond and only one broke in the element, indicating the relative strength of the two.

There is a

dramatic difference between the average strengths displayed between the TAGS and 3P, a factor of five.

The difference is all the more empha-

sized by the fact that all six TAGS couples broke at the bond while all but one of the 3P couples broke in the elements.

I

Further, the data

INSD-2650-29

111-664 ..

.. .-

.

.-

.......

.-

.


-1

TABLE 111-12

29 COUPLES

TENSILE TESTS ON SNAP

ij Couple

u

3P/2N

3

J

Identification

N-leg, psi

P-leg, psi

26-21

203

122

26- 22

396

142

26-23

357

101

26-25

272

170

26-27 26-28

388

105

418

91

Av 339 (1)

- - - _- - - - - - - - - - - - - - - - -

3

TAGS-85/2N

a 1 J

a

- - - - - - - - - - - -

137

462

857

138

453

524

139

162

a62

67-25

283

469

67- 69

337

415

67-125

371

626

Av 344 (1)

v

Av 122 ( 2 )

(1) Characterized by bond-failure

( 2 ) Characterized by element-failure

1

INSD-2650-29 111- 67

Av 625 (1)


J-

/-

i n d i c a t e s t h a t t h e TAGS bond i s s u p e r i o r t o t h e 2N bond by a f a c t o r of two.

No q u a n t i t a t i v e statement about t h e r e l a t i v e s t r e n g t h s of

TAGS and 2N can be made, b u t a q u a l i t a t i v e i n d i c a t i o n r e s u l t s from t h e

f a c t t h a t TAGS has withstood a f a c t o r of two t e n s i l e load without f a i l u r e .

L

Finally, both TAGS and 2N a r e c l e a r l y s u p e r i o r t o 3P i n t e n s i l e loads. Two types of drop t e s t s were conducted.

Type 1 consisted of drop-

ping t h e element a known d i s t a n c e onto a s t e e l b a l l embedded i n a wooden block.

The impact a r e a on t h e element was c o n t r o l l e d by dropping t h e

element i n s i d e a g l a s s tube l o c a t e d over t h e b a l l .

m e 2 was similar t o

type 1 except that t h e b a l l w a s dropped onto t h e element.

weighed 24.7 grams and was dropped a distance of

4-2inches.

The s t e e l ball

The ele-

ment was placed on a T r a n s i t e block one inch t h i c k . Figure 111-39 and Table 111-13 show t h e results of t h e type 1 drop test.

I n Fig. 111-39 t h e f i r s t hot pressed TAGS-85 sample shows t h e i m -

p a c t spots on t h e element surface.

This element was dropped with t h e

bottom of t h e tube approximately one i n c h above t h e s t e e l b a l l .

As a

r e s u l t , t h e element cocked as it emerged and w a s n o t s t r u c k i n t h e c e n t e r ;

it chipped a t t h e edges.

The o t h e r two elements were centered when t h e y

s t r u c k t h e b a l l ( t h e impact mark i s on t h e bottom) and no failure occurred a f t e r f i v e drops.

The element bounced when it s t r u c k t h e b a l l .

The

PbTe elements show impact p o i n t s and f r a c t u r e s a f t e r only one drop.

The

c a s t TAGS-85 elements were dropped f i v e times and only one s p l i t on t h e f i f t h drop. New PbTe elements were used i n t h e type 2 test, b u t the elements which survived t h e type 1 t e s t were reused.

TAGS-85

The r e s u l t s are

E


1

I

7

J -l

a 1

INSD- 2 6 5 0 - 2 9 111- 69

m m


TABLE 111-13

DROP TESTS OF TAGS AND

LEAD

TELLURIDE

t

Element dropped on s t e e l b a l l Drop height 1 2 i n . Material

TAGS-85

Diameter (in.)

0.380

Length (in.)

0.435

(hot p r e s sed )

TAGS-85

0.374

0.444

(cast)

mop

i

Remarks

E 5

Chipped on edges; element not h i t t i n g i n c e n t e r

5

No f a i l u r e

5

No f a i l u r e

5

No f a i l u r e

5

No f a i l u r e

5

No f a i l u r e

5

Cracked on 5th drop

i

Cracked on 5 t h drop

PbTe (2p)

0.380

0.445

Chipped on edge Large c h i p off t o p Large c h i p off t o p Cracked, then s p l i t

INSD-2650-29

111-70 -

.....

.


TABLE 111-13 (Continued)

Gi 3 J

24.7-g;m s t e e l b a l l dropped on element Drop height 4 . 5 i n . Material

Diameter (in.)

Length (in.)

*TAGS-85

a a PbTe

0.380

1 3

a

0.410

Drop

No. -

5

No failure (hot pressed)

5

No f a i l u r e ( h o t pressed)

5

No f a i l u r e ( h o t pressed)

1

S p l i t on 1st drop ( c a s t )

2

Cracked on 1 s t drop; s p l i t

2

Cracked on 1st drop; s p l i t

2

Cracked on 1st drop; split

2

Cracked on 1st drop; s p l i t

"Samples from type1 t e s t

3 J

J

3

Remarks

INSD- 2650-29

111-71


E t a b u l a t e d i n Table 111-13.

The PbTe and c a s t TAGS-85 samples s p l i t

and cracked a f t e r one o r two drops of t h e b a l l .

None of t h e hot

E

pressed TAGS samples f a i l e d . A r a t h e r s t r i k i n g d i f f e r e n c e was noted i n t h e compressive s t r e n g t h

of TAGS versus 2P and 3P PbTe as shown i n Table 111-14.

I n most . i n -

stances, as t h e pressure increased, t h e element chipped and f r a c t u r e d along t h e outer edges and f a i l u r e occurred by massive cracking or shear. Figure 111-40 shows some of t h e test.

TAGS-85and PbTe elements after

Note t h e completeness of f r a c t u r e i n t h e PbTe elements.

These

elements sheared and crushed a t almost t h e same pressure t h a t f i r s t i n d i c a t e d f a i l u r e was occurring.

On t h e o t h e r hand, t h e TAGS m a t e r i a l

showed l a r g e d i f f e r e n c e s i n p r e s s u r e from t h e time edge cracks ( u s u a l l y a t a pressure higher than t h a t a t which t h e PbTe elements f a i l e d ) f i r s t

appeared t o t h e time when f i n a l f a i l u r e occurred.

Note a l s o t h a t only two

samples f a i l e d i n a manner s i m i l a r t o t h e PbTe elements.

The other

samples d i d not shear b u t chipped along t h e l o n g i t u d i n a l axis.

The im-

p o r t a n t items t o note are t h e much higher compressive s t r e n g t h s of both c a s t and hot pressed TAGS compared t o 2P and 3P PbTe and t h e s h a t t e r i n g type f a i l u r e of t h e l e a d t e l l u r i d e s t o a more gradual edge-chipping f a i l u r e as shown i n Fig. 111-40. The semi-quantitative conclusions that can be drawn are as follows: (1) Hot pressed

TAGS-85 samples a r e mechanically superior t o

t h e c a s t TAGS-85 used i n t h e s e t e s t s .

However, new c a s t i n g

techniques a r e producing elements with mechanical p r o p e r t i e s

similar t o h o t pressed elements.

111-72

h i


.....

4RR

TABLE 111-14

COMPRESSION TESTS OF TAGS AND LEAD TELLURIDE

Pressure (PSI) Material

TAGS-85

Diameter Length (in.) (in.) 0.373

0.445

(hot pressed )

2P PbTe

0.380

0.445

a

Edge Chipping

Failure

19,692

24,361

29,174

35,009

25,294

28,186

14,454

20,236

---

12,351

14,070

14,774

J 3

a

20,959

J io, 610

J

10,280 io,610

Average

10,850

INSD-2650-29

111-73


TAGS-85 Hot Pressed

2P PbTe Pressed and Sintered

TAGS-85 Cast

i

I

FIG.

111-40.

APPEARANCE O F COMPRESSION T E S T S A M P L E S


a (2) Both hot pressed and cast TAGS-85are superior to cold pressed and sintered 2P and 3P lead telluride.

3

5.

Sublimation The relatively high vapor pressures with attendant high sub-

limation rates of thermoelectric materials is a well-known phenomenon. Table 111-15 provides a comparison of the vapor pressures of some thermo-

1

electric materials of interest.

J 3

J J J

a a 1

a

TABLE 111-15 VAPOR PRESSUIES OF SOME THERM0EL;ECTRICMATERIALS

T/E Materials

Vapor pressure, torr 45OoC 51OoC 600'~

2N*

2.1

3 p

5.7 x 10-7 6 x 10-

SnTe*

5.8 x

TAGS-85

10-6 1.9

---

Backgrov nd Pressure torr

10-5 1 x1~-3

2.4 x

7.6 x 10-5 2 x 1013 9

10-5

---

2 x 10-5 2

x 10-5

2 x 10-51x

+s~-m-68-500 (Sandia Corporation), Vaporization of Some Commercial Telluride Thermoelements, D.A. Northrop. Inspection of the data shows that thermoelectric materials occupy a rather small vapor pressure range which is much higher than the ordinary struc-

tural materials at comparable temperatures. The vapor pressure of TAGS-

85 was determined at a higher background pressure than the other materials shown. A determination at the equivalent background pressure would show a somewhat higher vapor pressure for TAGS-85. Control of sublimation can be effected by appropriate generator

1

design considerations.

Specifically with a minimum internal pressure

INSD-2650-29 111-75


of 2 psia maintained throughout the life of the mission and with Microquartz or a similar material packed loosely into the annular spacing between the elements and the a n - K 1301 insulation, sublimation of the thermoelectric materials may be reduced to insignificance. Atoms of the inert cover gas interfere physically with molecules escaping from the surfaces of the thermoelements by reducing %he mean free path of the molecules.

The path reduction is a function of the cover

gas pressure and to some extent its atomic weight.

c G

A notable example

of the effectiveness of this means of reducing material losses under similar conditions is the krypton filled high wattage incandescent lamp. Such a lamp h a s a life of about 2500 hours compared to minutes if the

filament was exposed to a hard vacuum.

A second and more effective

means of reducing sublimation is obtained by filling the annular space with loosely packed Microquartz.

In this instance a physical inter-

ference is again provided, but in this particular case the subliming species thermselves as well as the fibrous Microquartz combine to reduce the mean free path of the subliming molecules to a much greater degree than the cover gas alone. Evidence which provides strong support for the above conclusions comes from the operational experience of the SNAP 9A generator on-board the Transit Satellite L363-49B launched December

1963. Data from this satel-

lite indicate the internal pressure to be a "hard vacuum" after two years in orbit.

There is, however, reason to suspect that a low internal

pressure still exists.

Therefore, with a slight internal pressure and,

INSD-2650- 29

111-76

c c


significantly, no fibrous material packed into the annular space between the elements and the f i n - K 1301 insulation, the generator remains operational 5-l/2 years after launch, despite 3-l/2 Fars at a minimal

(<I psi) internal pressure.

3

a

k INSD- 2650- 29

111-.77


C.

FILL GAS MANAGEMENT

This section presents the results of generator fill gas management studies for a Viking RTG.

675 watts

Fuel inventories of 625 and

were considered in conjunction with helium release rates ranging from 0 to 100% of the theoretical value generated during fuel decay.

Generator seal tightness* was also varied from 0.01 x scc/sec.

l o m 5to

This tightness range encompasses actual SNAP

erator hardware ( ranges from 1 to 5 x

5.0 x

19 gen-

) as well as alternate

generator designs with twin seals, metal seals or welded housings.

c

The model used for these analyses is discussed below. 1.

Analytical Model The multi-gas permeability model used for generator pressure and

gas composition predictions is illustrated in Fig. 111-41. For a given generator void volume and average gas temperature ( VG and TG), c

the relationship between partial pressure of the gases considered and permeation rates across soft seals c a n be expressed as

Q + B x C X AP

where

Q = rate of a permeating gas, scc/sec

B

=

permeability constant for the gas at the seal temperature considered

*

C

=

installation tightness factor of the seal, cm

LP

=

partial pressure difference across the seal, atm.

F

Referenced to 3 3 p F seal temperature and one atmosphere differential argon pressure.

INSD- 2650-29

111-78

..

.

.

-.~

. . .

--


I

Q Q Q

1

One elastomer s e a l i s located i n 'the cover ( l a r g e a r e a a t each end of t h e generator) ; two smaller s e a l s are located i n t o p cover e l e c t r i c a l feedthrough connectors.

QA FIG. 111-41. MULTIGAS PERMEATION MODEL

INSD 2650-29

111-79


The resistance of the elastomer to gases is defined by B whereas the configuration match between seals and grooves is given by C. and

AP are time variables.

Only Q

The partial pressure inventories are

calculated and adjusted with time by iterative computer routines which also consider the replenishment of helium gas in the generator through gas release from the fuel capsule. Possible sources of error lie in assumed relationships between permeability constants of the varigus gases with seal temperature, in the measurement of the seal installation tightness factor (determined from hot argon leakage tests on previous SNAP 19 generators), and in the assumption that helium discharge f r o m the capsule is essentially constant through the mission period. Seal permeability and seal tightness are discussed in Ref.

TII-1. The excellent correlation of generator S/N 20 pressures by this model (using initial conditions of leak rate and pressure and the timetemperature-ambient gas history of the seal as input) is illustrated in Fig. 111-42.

Generator S / N 20 is currently undergoing thermal

vacuum tests at Jet Propulsion Laboratory. In general, correlation between data and prediction has been good for both the SNAP 19 endurance test generators (where helium leakage was not a factor) and for the flight SNAP 19 Nimbus I11 generators.

In the

latter case, good agreement was obtained by assuming a helium release rate of 96% of the theoretical helium production rate as illustrated in Fig. TII-43.

The 50% release rate pressure profile is also shown.

Tf the lower release rate ( 5 0 % ) had been experienced, pressure would

., .-,. ..x

.

..

INSD-2650-29 111-80


Q

P U

INSD- 2650- 89 111- 8 1


111- 8 2

INSD-2650- 29 -

.

-

-

-_


e have been 2 psi lower than measured in the 5-l/2 months between fueling and launch. Additional support that the helium release rate of

u

Nimbus I11 generators is near the 100% generation level is derived from performance considerations. The 96% level of release corresponds to a partial pressure of

4.8 psia helium in generators at the time of

launch. A 50% release level would have resulted in a helium pressure of only 2.5 psia and a generator power output nearly

P

at the system level than measured.

3 watts greater

It is this type of success in

correlation of pressure and power performance which justifies the detail of pressure work presented below.

2.

B P

Results a.

Pressure Profiles vs Seal Tightness and Helium Release

Pressure results for the parametric study in seal installation tightness and helium release rate from the fuel are plotted in Figs. 111-44 through

111-48for five periods during mission life, including

five years after launch. The plots are applicable to the proposed argon filled generators for the Viking mission which operate at a

1 I

a a

fin root temperature of about 33OoF.

Equivalent data for RTG's

initially filled w i t h helium are presented in Table 111-16. The allowable pressure range for RTG operation is as follows. At

the high limit, increasing parasitic thermal insulation losses restrict the maximum pressure to somewhat above t w o atmospheres.

lower

pressure limit is similarly fixed by performance considerations since the thermoelectrics experience accelerated losses by sublimation at

1 INSD-2650- 29

111-83


.

..

. ., .~ . .. . .~

'

! . . ...

.

LI.

-

..

i78 -111 62 - O S 9 Z -CISNI

.

.

.


P

P

P

INSD-2650-29 Ill-85


INSD- 2650- 29 111- 8 6


n

n I

J 3

1

J 111- 8 7

INSD- 2650- 29


. ,

,

- ,

.

. . . -

1

... ,.

I

INSD-2650-29 111- 88

.,

.. .

.--..- . .

I

.

.. .

-

I

i i

i

c

E

E

E

w 0

_-


...

TABLE 111-16 SNAP 1-9 HELIUM FILLED GENERATOR INTERNAL pRESSWU3S (PSIA) FOR COMBINATIONS OF SEAL TIGH'ITESS AND HELIUM RELEASE RATES F I N ROOT TEMP. 33OOF. INITIAL FUEL INVENTORY 675 WATTS Seal Tightne s s Factor scc/sec 1.0

10-7

Helium

Time from Fueling 0

scc/sec

0 2.80 x 10-6

7.0 14.0 21.0

25.2 28 .o 1.0 x 10-6

14.15 20.32 29.63 45.15 60.67 69 98 76 19

13-29 13.82 14.61 15-92 17-23

11.86 15.41

10.02

12.50 13.01 13.77 15.04 16.31 17-07 17.58

11.68 13.22 15-53 19 37 23.22 25 53 27 * 07

21.0

25.2 28.0 0

11.57

2.80 x 10-6 7.0 14 .O 21.0

25.2 28.0

1.5 x 10-5

14 39 18.36 24.34 34-30 44.27 50.25 54.24

18.54

7.0 14.0

0

10.71

14.02 15.24 1 5 97 16.46

16.21 19 63

73

11.20 11.go

13.08 14.26 14.96 15.43

21 .o 25.2

28.0

-.

9.34

12.06 12.80

10

2.8 x 10-6 7.0 14 .O

21 33 -

14.64 16.37 18 95 23 27 27.59 20.19 31 92

28.0 0

10-5

Months

13.48 14.01 14.80 16.12 17.44 18.24 18.77

18.02

2.80 x 10-6

1.0

9

25.2

21.0

x

3

14.04 15 73 18.26 22.49 26.72 29.25 30 95

0

2.80 x 10-6 7.0 14.0

5.0

-

Release Rate

.

1NS~-2650-29 111-89

12.77 21.70

23 *07 7.52 8.75

10.60 13.68 16.76 18.61 19.84

20.71 29 59

38.46 43 79 47.34 5.08 7.30 10.62

15-15 21.69 25.02 27.24 1.91 3.28 5-33 8.74 12.15 14.20 15.56 91

1.85 3-25 5.58 7.92 9.32 10.25

15* 1 3 22.78 35.58 48.38 56.06 61.18

69

13.45 26.01

44.98 76.60 108.22

127.20 139 85 6.04 14.54 27.24 48.51 69 79 82 55 91.06

2.25 4.76 8.52 14.78 21.04 24.80

2.98 7.05 13.82 20.57 24.64

27.31

27.34

-53

17 1.54 3.58 6.99 10.40 12.44

1.90 3 *94 7.35 10.76 12.80

14.16 35 1.26 2.63 4.91 7.18 8.55 9.46

0.28

13.80 .20 1.11

2.47 4.75 7.02

8.38 9.29


TABLE

Seal Tightness Factor

scc/sec 2.0

10-5

111-16(Continued)

Time from Fueling

Helium Release Rate scc/sec

9.95 10.40 11.08

2.8 x'io6 7.0 14.O

12.21

13 34 14.02 14.48

21.0

25.2

28.0 3.0 x 10-5

8.57 8.99 9.62 10.67 11.72

2.8 x'io6 7.0 14.O 21.0 25.2

12.35

28.0 4.0 10-5

12

0

2.8 x 7.0 14.0 21.0 25.2

28.0 5.0

10-5

2.8 x io-6 7.0 14.0 21.0 25.2

28.0

9

3

13.50' 1.3 50 13 50 13.50

77

6.12 7.23 8.90 11.68 14.45 16.12 17.23 4.24 5 -15

6.53 8.81 11.10

12.47 13 38

7.41 7.80 8.39 9-37 10.34 10 93 11.23

3.17 3.94 5-09 7.01 8.92 10.07 10.84

6.44 6.80 7.35 8.26 9.17 9.71

2.61 3.26 4.24 5.87 7-50 8.48 9.14

10.08

Q -

Months

21

.64 1.33 2.38 4.12 5.86 6.90 7.60 .63 1.09 1.77 2.91 4.06 4.74 5.20 *7 3 1.07 1.59 2.44 3.29 3.80 4.. 14

33 -

.38

I.06

2.08 3.79 5.49 6.52 7.20 .46 -91 1.60

2.73 3.87 4.55 5 .oo 50

.21

90 1.92 3.62 5.32 6.35 7-03 .20 * 65 1.34 2.47 3.61 4.29 4.74

.17

-

.84 1.36

51 1.02

2.21

1.87 2.72

3.06 3.57 3 991

.82

-52

1.10

.79

1.50

1.20

2.19 2.87 3.28 3.55

1.88 2.56 2.97 3.24

z

69

3.23

3.57 13 .40 .81 1.49 2.17 2.58 2.85

INSD-2650- 29

111-90

_

.

-


ia

u Q

very low pressures.

h a a c t lower limit has not been established, but

pressures greater than 2 psia are considered acceptable for stable per-

formance

.

As shown in the previous section, evidence from Nimbus I11 is that helium Pelease from the fuel approaches 100%. This value corresponds to a helium production rate of 2.81 x

scc/sec for a

675 watt source,

a quantity not to be confused with seal tightness rates. As shown in

a a

Fig. 111-4.6 for the expected seal tightness range of 1 x t.o 5 x lO-â&#x20AC;&#x2122;(the

scc/sec

present RTG design with one Viton O-ring for each

sealed surface has produced seal tightnesses within this range on all generatos built to date), the generator pressure at the end of the Viking mission will be approximately 2 psia or greater even for 0%

a J 3

a

release from the fuel. tightness ( -3

x

For the expected release ( . c r l O O % ) and seal scc/sec) the pressure will be approximately

7 psia. At 2 years after launch (Fig. 111-47) the expected pressure is approximately

5 psia.

For ease of interpretation of figs. 111-44 through

111-48,

pressure profiles of the initdal argon filled RTG are presented in Fig. 111-49 for various seal tightnesses and for helium pelease rates of 100% and SO$. b.

Effect of Launch Scheduling on RTG Performance

Post fueling generator test and storage periods of 9 months were assumed for the pressure and performance analyses previously presented.

An extension or abbreviation of the nine-month prelaunch period w i l l have minimal effects on pressure levels at encounter but w i l l generally

INSD-2650-29 111-91


AI

,

. .

e

,.

I. . . .

,

.

..

E

.

.

...

.

c

dE INSD-2650-29 111-92


be evident at launch. Each combimtion of seal tightness and release rate is unique.

This is illustrated in Fig. 111-50 f o r test and stor-

age periods of 3 and 1 5 months and release rates of 100% and 50%. seal tightness of 3 x

A

scc/sec was assumed. Excessive internal

pressure rise in generators with tight seals and maximum helium rates

a a 1

a a

during prolonged storage periods can effectively be halted by raising seal temperatures above the nominal storage value (230°F to 2 6 0 ~ ~ ) by insulating the fins.

Impact on initial power performance can

thereby be minimized. Figure 111751 shows the effect of the prelaunch storage and hold period for fueling-to-launch-periodsof

3 months and 15 months (at

100% and 50% release rates and a seal tightness of 3 x

scc/sec).

The variation in performance at end-of-mission (EOM) is approximately one watt for a given helium release rate, with the shorter storage time resulting in the higher power output due to the lower helium

a 1

pressure in the RTG. c.

Helium Gas Concentration

The change in percentage of helium gas at known pressures during

mission life is of prime interest for performance predictions.

A

rising helium content will depress generator thermal efficiencies after fueling by raising the effective thermal insulation conductivity. Seal tightness and helium release rates were again combined in parametric studies versus helium concentration at various mission times.

3

Results are presented in the carpet plots of Figs. 111-52 through

I

111-56. It is of interest that for mission durations up to 2 y e a r s

1

INSD-2650-29 111-93


3

a 3

a 3

111- 9 5

INSD- 2650- 29


111-96

INSD- 2650- 29

D E L

r f


INSD- 2650- 29 111- 9 8 .

.,

-- .

.r

..

"....-

. -. .

. -. .

.~~

.

.


111-99

INSD- 2650- 29


-. -.

.

-

--

...................

.....................


...

.

....

.

..

.

.... ..

..

-

.

.

.

.

the seal tightness has a rather minor influence on helium concentration. d.

Effect of fuel inventory and temperature

Figure 111-57 presents the total pressure for a Nimbus I11 configuration generator with a

625 watt thermal inventory and

350'~ fin

root (seal) operating temperatwe. At the EOM, the generator nominal pressure is ,5

psia.

Thus, the effect of the higher fin root tempera-

ture and lower fuel inventory is to reduce the EOM pressure 2 psia below the 675 watt, 33OoF design previously discussed. e.

Generators with initial helium fill

Pressure changes with time for a helium filled generator are illustrated as a function of seal tightness and helium release rate in Table 111-16. The current elastomeric seals are more permeable to helium than to argon.

To retain a supporting pressure in the generator,

the helium filled unit must rely on gas replenishment from the caps.ule

to a greater extent than the argon filled unit.

Pressures approach

levels dictated by release rates soon after launch whereas equilibrium values are approached more gradually by the argon filled generator.

Representative pressure time profiles are presented in

Fig. 111-58. In general, the helium filled RTG will produce a somewhat flatter power profile over the mission than the argon RTG but somewhat less integrated power.

Either RTG w i l l meet the Viking requirements, al-

though the argon RTG has been selected f o r emphasis in this study because of the greater power output available at EOM.

INSD-

2650-29

III- 101


111- 102

INSD- 2 6 5 0 - 29


3.

Viton 0-Ring Summary Pressure p r o f i l e s and

%

helium concentration as presented above

a r e supported by s i x years of elastomeric s e a l evaluation on SNAP programs.

The Viton A compound

V77-545 O-rings as procured from

Parker Seal Company are c a r e f u l l y inspected f o r compliance with d i mensional and i n t e g r i t y s p e c i f i c a t i o n s .

Similar inspections a r e

made on generator s e a l i n g and groove surfaces p r i o r t o i n s t a l l a t i o n . The s e a l s a r e held t i g h t by r i g i d covers and flanges which v i r t u a l l y eliminate s e a l motion during v i b r a t i o n or handling of t h e generators. The i n t e g r i t y of s e a l s i s determined through helium and hot argon leakage tests of completed assemblies. The performance of t h e i n s t a l l e d seal depends on i t s r e s i s t a n c e t o surface d e t e r i o r a t i o n and on t h e memory of t h e s e a l compound a t time and temperature.

Viton A unlike some elastomers, i s superior i n

- r e s i s t a n c e t o surface cracking and sublimation under high vacuum ( R e f . 111-2).

The memory of t h e compound, which r e l a t e s t o t h e c a p a b i l i t y

t o maintain a s e a l by compressive f o r c e s i s good a t sustained temperat u r e s up t o 375OF.

L i t e r a t d e c l a i m c a p a b i l i t y t o withstand short-time

exposures t o 1 0 0 0 ~ ~ . Viton A has widely been used i n s e a l i n g a p p l i c a t i o n s for space.

It has been operated a t temperatures i n excess of 380°F during t h e q u a l i f i c a t i o n cycle of SNAP t h e SNAP

9A S/N 4 generator

19 generators

and has s u c c e s s f u l l y sealed

which has o r b i t e d s i n c e

1963.

Moreover t h e

s e a l s i n t h e c u r r e n t configuration have been temperature cycled from

INSD-2650-29 111-104


7

350.F to room temperature at various occasions during the SNAP 19 endurance test generator program.

Some of these RTG tests are still

in progress. A summary of Viton A seal applications with SNAP RTG's is presented in Table 111-17. Seal performance in all instances has

been satisfactory.

3 3

a 1 INSD-2650-29 III-105


TABLE 111-17

VITON SEAL USAGE SUMMARY I N ISOTOPES PROGRAMS

Generator Program

Serial Number

Oper a t i n g Time ( hrs

-

SNAP

19

w

Nominal Seal Temp.

(as o f )

( OF>

Status

1

4,000

9/66

230-270

Terminated

2

4,000

7/67

230-270

Terminated

3

3, 8Qo

1965

230-350

Terminated

4

3,

1965

230-350

Terminated

E

E

4A

10,400

1/1/68

350

Terminated

5

11,900

12/67

350

Terminated

6

13,400

12/67

350

Terminated

7

4,300

6/8/67

270-350

Terminated

8

4,400

6/8/67

270 - 350

Terminated

9

800

5/68

230

A t Goddard f o r Magnetic Testing

10

800

5/68

230

A t Goddard f o r Magnetic Testing

11,llA

5, 700

i/io/68

270-350

Terminated

12,12A

5,700

i/io/68

270-350

Terminated

13

4,600

1/16/68

270-350

Terminated

14

4,600

1/16/68

270-350

Terminated

15

1,000

5/68

230

Terminated

16

1,000

5/68

230

Terminated

17

5,800

12/67

350

Terminated

18

9,000

4/68

350

On t e s t a t NSRDC

16,000

4/69

350

F i r s t 5460 hours a t Isotopes; remainder a t NASA-MSFC; t e s t continuing.

19

'

.

INSD-2650-29

111-106 . _ _ _

- - -


- -

TABLE 111-17 (Continued)

19

Generator Program - SNAP Serial

Number

Operating T i me (hrs)

io, 150

20

(Cont'd)

(as of)

Nominal S e a l Temp. ( OF>

2/69

350

A t JPL

Status

21

8,300

3/5/69

350

A t JPL

22A 23A

5 ,000 5,000

5/68 5/68

240-365 240-365

Numbus B F l i g h t System 8 A Generators - Terminated

24

5,500

25

5, 500

6/69 6/69

240-375 240-375

Nimbus 111 F l i g h t System Generators - i n Orbit

26

I,200

6/69

330

Life Test

27

200

6/69

330

Awaiting D i spos it ion

28

200

6/69

330

Awaiting Disposition

Generator Program Q/N- IM Q/N-

3

Q/N- 4

SNAP 11

15,000

9/68

250-400

T e s t continuing a t Sandia

17,000

2/69

250-400

Test continuing a t JPL

8,35?.

12/20/67 250-400 Generator Program

Transit

-

> 6,000

Unknown h i s t o r y a t WC (Houston)

-

SNAP 9A

6/1/64

275

S a t e l l i t e memory f a i l u r e termina t e d generator data; pressure w a s 6 psia.

41,000

8/68

275

On 3/66 (20,000 hours) i n d i c a t e d p r e s s u r e reached zero; RTG i s s t i l l functioning.

S/N 2

20,300

--

215-275

Terminated

S/N 3

io,800

--

215-275

Terminated

S/N 5

13,900

--

215-275

Terminated

5BN-1 (S/N 6) Trahsit 5 ~ ~ - 2

(S/N 4)

INSD-2650-29 III-107


E E F E

E

dL


/ 7

I

-

Q

\

IV.

HEAT SOURCE

The design of a n I n t a c t Impact Heat Source (IIHS) f o r an RTG i s most s t r o n g l y influenced by nuclear s a f e t y requirements.

Q

Accord-

ingly, analyses of t h e heat source reported i n t h i s chapter a r e developed

i n a way t h a t f a c i l i t a t e s a preliminary s a f e t y evaluation and permits r a t i o n a l decisions with regard t o design s e l e c t i o n , A.

GENERAL

The primary purpose of nuclear s a f e t y a n a l y s i s i s t o provide component design c r i t e r i a f o r t h e nuclear system.

The o b j e c t i v e s of these

c r i t e r i a a r e twofold:

3 zi 3

1.

t h a t t h e system w i l l s a t i s f y t o t a l s a f e t y requirements, and

2.

t h a t t h e r e s u l t i n g design w i l l be balanced so t h a t no compon-

e n t i s u n i n t e n t i o n a l l y over-designed. The s a f e t y function must be comprehensive t o be u s e f u l .

Failure

of t h e f u e l containment i s a matter of concern, whatever t h e cause. Although t h e magnitude of t h e nuclear assessment v a r i e s with conditions under which f u e l r e l e a s e occurs, that i s no s i n g l e event, such as t h e

a

r e - e n t r y process, which s o dominates with r e s p e c t t o nuclear s a f e t y s i g n i f i c a n c e t h a t a l l o t h e r considerations may be suppressed i n designing t h e nuclear system. Preliminary analyses were performed of a l l p o s t u l a t e d malfunctions which have a dominant influence on t h e design of t h e heat source.

Two

candidate h e a t s h i e l d s f o r a n i n t a c t impact h e a t source were examined

INSD -26 50 -29 IV-1


i n t h i s study. B.

The design options a r e defined and discussed i n Section

Equilibrium temperature p r o f i l e s f o r t h e s e configurations are pre-

sented i n Section C. Malfunction a n a l y s i s was separated i n t o t h r e e c a t e g o r i e s

--

launch

pad and f l i g h t a b o r t s (Sections D and E ) and impact-post impact containI n each case, environments consequent t o t h e malfunc-

ment (Section F).

t i o n were considered as w e l l as t h e response of each candidate configuration.

The s p e c i f i c o b j e c t i v e of t h i s a n a l y s i s i s t o estimate f u e l r e l e a s e

p r o b a b i l i t i e s for each environment and candidate heat s h i e l d . I n t h e f i n a l s e c t i o n (Section G) release p r o b a b i l i t i e s a r e combined w i t h indices which a r e measure o f t h e r e s u l t a n t nuclear safety assess-

ment and summed f o r each candidate h e a t s h i e l d .

When t h i s i s accomplished,

it i s p o s s i b l e t o s e l e c t t h e safest design and deduce f u r t h e r modificat i o n s , i f required. B.

INTACT

IMPACT

HEAT SOURCE (IIHS)

DESIGN

The u l t i m a t e requirements of t h e h e a t source a r e t h a t it maintain

fuel containment under a l l normal conditions of t r a n s p o r t a t i o n , handling, checkout, launch, r e - e n t r y and Earth impact as well as c e r t a i n abnormal ( b u t c r e d i b l e ) p o s t u l a t e d conditions which may occur due t o t e r r e s t r i a l accidents or launch or f l i g h t a b o r t s .

The f u e l capsule must be made of

m a t e r i a l s that a r e chemically compatible with t h e f'uel, r e s i s t oxidation

and seawater corrosion, and have s u f f i c i e n t s t r e n g t h and d u c t i l i t y t o

r e s i s t shock, v i b r a t i o n , shrapnel p e n e t r a t i o n , and E a r t h impact a t component temperatures a s s o c i a t e d with t h e s e environments.

.. INSD-2650-29 IV-2

I .

I


Two types of capsule f a b r i c a t i o n have been adequately explored i n p a s t programs and a r e d i r e c t l y a p p l i c a b l e f o r Viking use.

The f i r s t

concept a v a i l a b l e i s one u t i l i z i n g t h e combination of r e f r a c t o r y metals f o r s t r e n g t h and impact r e s i s t a n c e and refractory/noble metals f o r f u e l

8

compatibility, corrosion and oxidation r e s i s t a n c e .

The s i g n i f i c a n t

advantages of t h i s type of design are that steady s t a t e capsule operat i o n temperatures of 20000F may be t o l e r a t e d .

This approach has been

used i n SNAP 11 and i s c u r r e n t l y being developed f o r SNAP

19 type

RTG's.

I

The second concept uses a superalloy capsule having t h e c a p a b i l i t y of operating a t steady s t a t e temperatures of 1400°F or s l i g h t l y higher. This type of f u e l capsule i s w e l l developed and w a s used s u c c e s s f u l l y i n SNAP 3B, gA, 19, and 27.

A l l of t h e s e RTG systems except SNAP 27

were developed by Isotopes, Nuclear Systems Division. Design of t h e SNAP

a i

a

19 h e a t source f o r i n t a c t r e - e n t r y and i n t a c t

impact i s constrained by t h e hexagonally-shaped c e n t r a l c a v i t y of t h e thermoelectric core.

On t h e

SNAP 19 Nimbus program, t h e s i x thermo-

e l e c t r i c modules were supported around t h i s c a v i t y by a g r a p h i t e heat accumulator block that accepted h e a t from t h e c e n t r a l l y l o c a t e d h e a t source.

I t has been found t h a t by u t i l i z i n g t h i s e x t e r n a l l y configured

hexagonal g r a p h i t e heat accumulator block as an i n t e g r a l a d d i t i o n t o t h e previously developed c y l i n d r i c a l SNAP 19 g r a p h i t e heat s h i e l d , t h r e e d i r e c t improvement e f f e c t s on r e - e n t r y are obtained:

1.

i

a

Additional g r a p h i t e thickness on t h e order of i s provided

for a b l a t i o n r e s i s t a n c e ,

INSD-2650-29

IV-3

160 m i l s


2.

Hypersonic b a l l i s t i c drag c o e f f i c i e n t (W/C A) i s decreased D 2 from 37 t o 34 l b / f t due t o t h e l a r g e r e f f e c t i v e radius and corners, thereby reducing stagnation aerodynamic heating.

3.

Subsonic drag c o e f f i c i e n t i s increased from 0.34 t o 0.67 t h e r e f y reducing t h e c a l c u l a t e d impact v e l o c i t y from 335 t o 245 f p s .

These e f f e c t s p l a y a l a r g e p a r t i n a s s e s s i n g p o s t u l a t e d f u e l release probabilities.

They a f f e c t s a f e t y margins with respect t o h e a t

source melt and a b l a t i o n temperatures during p o t e n t i a l r e -entry and capsule rupture a t E a r t h impact, two of t h e primary modes of p o t e n t i a l f u e l release. 1.

Capsule Design A s i n d i c a t e d i n Fig. IV-1, t h e high temperature f u e l capsule has

t h e configuration of a s h o r t c y l i n d e r with hemispherical ends.

It is

f a b r i c a t e d i n two l a y e r s , a n i n n e r l a y e r (25 m i l s t h i c k ) made of molybdenum 50 weight percent rhenium a l l o y which c o n s t i t u t e s t h e primary containment v e s s e l f o r t h e r a d i o i s o t o p i c f u e l , and an outer l a y e r (nominally

60 m i l s ) of TZM which provides r e s i s t a n c e t o any mechanical

environments (impact, b l a s t , shrapnel, crushing, e t c .) which r e s u l t

from p o t e n t i a l t r a n s p o r t a t i o n , handling or f l i g h t accident, or from E a r t h impact a f t e r re-entry.

The TZM s t r e n g t h member may be coated

with rhodium, i f required, t o minimize t h e e f f e c t of oxidation a f t e r such a n u n l i k e l y a c c i d e n t ,

The capsule i s designed t o contain up t o

755 watts(t) of PU-238 O2 a t a bulk power d e n s i t y of 2.7 watts per

INSD-2650-29 IV-4


la

P

J J 13

J

cd

- -

-.

.

.

.. INSD-2650-29 IV-5

.-


cubic centimeter, although only 675 watts are required f o r t h e Viking RTG design.

a.

Primary container ( l i n e r )

The primary container ( l i n e r ) has a major function, t h a t of f u e l containment under a l l normal operating conditions without undue i n t e r a c t i o n with o r reduction of t h e f u e l form.

Other important character-

i s t i c s of t h e capsule l i n e r which bear on material evaluation and form t h e bases f o r s e l e c t i o n trade s t u d i e s include: (1) F a b r i c a b i l i t y

(2)

Mechanical p r o p e r t i e s

(3) Melting temperature

( 4 ) A v a i l a b i l i t y and c o s t ( 5)

Oxidation/barrier coating a p p l i c a b i l i t y

The l i n e r c o n s i s t s of two spun molybdenum 50% rhenium, 25 m i l t h i c k cups, joined by a c i r c u m f e r e n t i a l weld j o i n t .

Machined f u e l i n g p o r t and

f i l t e r - v e n t housing d e t a i l s , as shown i n F i . I V - 1 ,

are welded a t

opposite i n t e r n a l l o c a t i o n s i n t h e closed cup ends p r i o r t o making t h e c e n t r a l weld j o i n t closure. Fuel loading with PU-238 02 microspheres i s f a c i l i t a t e d by t h e dual l i n e r c l o s u r e a t t h e f u e l i n g p o r t so that complete decontamination of t h e l i n e r can be done a f t e r t h e i n i t i a l seal weld a n d p r i o r t o performance of t h e o u t e r c l o s u r e cap s t r u c t u r a l TIG weld. concept was u t i l i z e d i n SNAP

This design

19 IRHS ( 8 capsules of t h i s design have

previously been fueled) i n which a small amount of ZrO2 p a r t i c l e s was

INSD-2650-29

IV-6


It was

poured i n over t h e f i e 1 t o occupy t h e remaining void volume.

found t h a t by performing t h e f i r s t closure weld over t h e Z r O p a r t i c l e s , 2 w e l d contamination w a s minimal and decontamination p r i o r t o t h e o u t e r

Y

s t r u c t u r a l weld w a s made with ease.

0

Thin gauge m a t e r i a l s a r e used t o construct t h e l i n e r , and i t i s a n t i c i p a t e d t h a t s p e c i a l handling t o o l s w i l l be required during t h e Similar handling t o o l s were r e c e n t l y designed

f u e l loading operation, and b u i l t f o r t h e SNAP

19 IRHS

program, although they were u t i l i z e d

with a heavier gauge superalloy capsule.

3

With proper planning of

handling operations and appropriate t o o l and f i x t u r e design, capsule loading problems can be reduced t o a minimum.

I t should be pointed out

t h a t t h e TIG welds a t t h e two f u e l i n g p o r t closures (seal weld followed by s t r u c t u r a l weld) a r e the only welds required during t h e f u e l loading

1 and heat source assembly operation.

J 3

a

b.

Strength member

The o u t e r s t r e n g t h member i s machined from s t r e s s - r e l i e v e d TZM rod t o obtain a hard, r u p t u r e - r e s i s t a n t s h e l l around t h e l i n e r .

To

f a c i l i t a t e assembly a t t h e m e l i n g f a c i l i t y , a threaded, two-piece

assembly i s used.

Strength m e m b e r s using t h i s technique have been

successfully f a b r i c a t e d and t e s t e d including impact (16 t e s t s )

, plasma

a r c ( 7 t e s t s ) , and s o l i d propellant contact f i r e t e s t s ( 2 t e s t s ) .

J 3

recent s e r i e s of impact and plasma a r c t e s t results are given i n Tables I V - 1 and IV-2, r e s p e c t i v e l y . f o r o r b i t a l decay and

The plasma a r c t e s t simulations

3 6 f~t / s e c s u p e r o r b i t re-entry a r e given i n

Figs. IV-2 and I V - 3 .

i

A

INSD-2650-29 IV-7


TBLE -

Iv-1

IIHS IMPACT TEST PROGRAM* Test No.

Date -

Confirmration

Velocity (Ft/Sec)

Thread-2.47" dia.

1

6/12/69

Straight

2

6/16/69

Taper

3

6/19/69

Taper

4

6/23/69

Taper Thread-2.56" dia.

5

6/27/69

Plasma Arc Tested Taper

Thread-2.47''dia.

*

-2.56" dia. ~hread-2.56"dia.

Impact Angle

Result

301

60'

Female End

Failure i n male thread f u e l released

299

60'

Female End

Successful

297

60'

Female

Successful

302

60'

Female a d

Successful

298

60° Female md

Successful

0

relief;

6

6/27/69

Taper Thread-2.56" dia.

297

10 Female a d

Successful

7

6/27/69

Taper Thread-2.56" dia.

301

90'

Female

Successful

8

6/28/69

Plasma Arc Tested Taper ~hread-2.56"dia.

351

60'

Female End

Successful

9

6/28/69

Taper

298

30'

Male

Successful

lo

6/28/69

Plasma Arc Tested Taper Oxidized (2 times)

352

60'

Female End

Successful; 2 f i n e cracks in TZM; f u e l retained

i1

6/3/69

Bare

301

45O Female End

Successful

12

6/3/69

Taper Thread-2.56'' dia.

352

60'

Successful

13

6130169

Taper

Thread-2.56" dia.

349

10' Male

Successful

14

6130169

394

60'

Female End

Successful

15

7/1/69

Taper Thread-2.56"dia. w %re Taper Thread-2.56"dia.

350

60'

Female E$d

Successful; one f i n e crack i n TZM; f u e l retained

Thread-2.56" dia.

** Taper

(3 times)

%read-2.56"

Thread-2.56" dia.

dia.

oxLdized

*

All samples preheated t o 1900% p r i o r t o impact

**

No threads,using interference f i t

*+

Impacted Without graphite heat shield

Female End


TABLE I V - 2 IIHS PLASMA A R C TEST SIMULATING RE-ENTRY FROM EARTH ORBIT

TEST SUMMARY

Test

No. 1

Test

Date -

6/24/69

Heat Shield

Configuration Hex Prism

Test

Dimensions (in.)

Test Attitude

Condition

6-75 long x 3.5 between

Corner facing

Superorbit 36 K ft/sec

flats

2

6/24/69

Hex Prim

6.75 long x 3.5 between

flow

Yi

Corner facing

6/25/69

Hex P r i s m

6.75 long x 3.5 between

6/26/69

Hex P r i s m

6.75 long x 3.5 between

(lb)

Recession (in.)

NASA-MSC

--

.083

Failure in heat shield thread relief due t o over torquing of threaded end plug; capsule undamaged

Test F a c i l i t y

- Go

Results

Superorbit 36 K f t / s e c 7, - 6'

NASA-MSC

.048

.O83

Successful

Corner facing

Superorbit 36 K f t k e c

NASA-MSC

.046

.Ow

Successful

flow

Yi

flats

4

Maximum Surface

flow

flats

3

Simulation

Weight Loss

-6

Corner facing flow

Orbital &cay

Martin-Baltimore

.16

.077

Successful

Corner facing flow

O r b i t a l Decay

Martin-Baltimore

.16

.082

Successful

Corner facing flow

O r b i t a l Decay

Martin-Baltimore

.16

.082

SuccessfliL

Plats

5

6/27/69

Hex P r i s m

6.75 long x 3.5 between flats

6

6/28/69

Hex Prim

6.75 long x 3.5 between flats


0

8

8'

-

4

P

.I

IV-10


IT - A I

6a-04ga-as~1

c

D


The TZM s t r e n g t h member i s composed of two mating cups which are machined a t t h e open ends with a tapered threaded j o i n t .

Assembly i s

completed by merely i n s e r t i n g t h e fueled l i n e r i n t o t h e female s t r e n g t h member cup and then threading t h e male component i n t o f i n a l engagement. The completed capsule w i l l then be ready f o r i n s t a l l a t i o n i n t o t h e heat shield. There i s always some concern over whether threaded p a r t s may be cross-threaded during assembly.

Cross-threading i s caused by improper

alignment a t the s t a r t of the threading operation by a t l e a s t one-half

a t h r e a d lead.

I n t h e present design, as t h e male half of t h e s t r e n g t h

member i s lowered over the f u e l capsule l i n e r , it w i l l slide over t h e c y l i n d r i c a l surface of t h e l i n e r a d i s t a n c e of approximately 1 . 5 inches before contact with t h e mating t h r e a d s i s made.

If it i s assumed t h a t

a maximum diametrical tolerance of 0.005 inch i s provided, t h i s w i l l r e s u l t i n a maximum angular misalignment of 0.0033 inch or 0.19 degree. This corresponds t o less t h a n h a l f a t h r e a d lead.

This i l l u s t r a t e s

that cross-threading of t h e TZM s t r e n g t h member i s not p o s s i b l e because

of t h e alignment provided by t h e i n t e r n a l capsule l i n e r .

Because of

tapered t h r e a d torquing and t h e preload a p p l i e d t o t h e capsule during assembly i n t o t h e h e a t s h i e l d , t h e threads w i l l not back off as a r e s u l t of v i b r a t i o n loading. c.

Filter

A one-stage ceramic vent i s provided a t one end of t h e l i n e r t o

allow t h e helium generated within t h e capsule t o escape.

INSD-2650-29 IV-12

The vent

c


k3

serves t o contain the f u e l p a r t i c l e s within t h e capsule and gives low pressure drop t o t h e flow of helium.

This s t a g e i s similar t o materials

and i d e n t i c a l i n construction t o that used i n SNAP

19 IRHS,

f o r which

f u e l compatibility and f i l t r a t i o n p r o p e r t i e s with production grade

0

pu-238 O2 microspheres have already been t e s t e d and proven. An a d d i t i o n a l flow-regulating f e a t u r e can b e obtained by t h e a p p l i c a t i o n of a high density, v i t r e o u s oxide f i l t e r - v e n t t o d i r e c t l y replace t h e lower density m a t e r i a l as used i n t h e SNAP

19 IRHS

capsule.

The use of t h e higher density material w i l l provide a c o n t r o l l e d pressure drop across t h e f i l t e r - v e n t on t h e order of 10 p s i a t t h e nominal helium production rate of t h e decaying PU-238 O2 f u e l .

A s a r e s u l t of

n

t h e maintained helium gas pressure i n t h e f u e l cavity, f u e l c e n t e r l i n e

J J J

temperatures w i l l be minimized. Assembly of t h e f i l t e r element i n t o t h e vent r e c e p t a c l e w i l l be performed i n a manner similar t o t h a t developed f o r SNAP

a a I

It

involves t h e process of platinum p l a t i n g t h e o u t e r c y l i n d r i c a l surface of t h e ceramic f i l t e r element, then machining and i n s t a l l i n g it i n t o the machined filter housing receptacle and diffusion bonding t h e element

i n t o place.

1

19 IRHS.

Evidence from impact t e s t s performed w i t h SNAP

19 IRHS

i n d i c a t e s t h a t t h i s vent configuration i s impact r e s i s t a n t without swaging or staking t h e element i n t o place.

A pierce disc is u t i l i z e d

outside t h e vent t o permit complete decontamination subsequent t o f u e l i n g as well as t o p r o t e c t t h e vent element from t h e i n f l u x of d i r t o r cleaning s o l u t i o n .

The d i s c i s p i e r c e d p r i o r t o i n s t a l l i n g t h e l i n e r

i n t o t h e s t r e n g t h member t o a c t i v a t e f i l t e r venting of helium gas.

INSD-2650-29

IV-13


d.

Coatings

A zirconium oxide coating w i l l be required on t h e e x t e r n a l hemi-

s p h e r i c a l surfaces of t h e TZM s t r u c t u r a l member t o provide a d i f f u s i o n

barrier t o t h e t a n t a l u m f e l t c m p l i a n c e pad a t both ends when i n s t a l l e d i n t h e graphite heat s h i e l d .

Application of coating on t h e e x t e r n a l

surface of t h e TZM s t r e n g t h member i s performed as a r o u t i n e process p r i o r t o fueling.

It should be noted t h a t compatibility between t h e TZM s t r e n g t h member and t h e g r a p h i t e heat s h i e l d i s no problem s i n c e t h e m a t e r i a l s do not contact. Heat S h i e l d Design

2.

..

The g r a p h i t e heat s h i e l d , when f i t s within t h e e x i s t i n g RTG heat

source c a v i t y , has an e x t e r n a l configuration of a r i g h t hexagonal prism, 6.75 inches long and 3.50 Inches across opposing f l a t s .

Fabrica-

t i o n of t h e heat s h i e l d i s accomplished by boring out a Carb-I-Tex 502 tube and machining t h e outside t o t h e hexagonal shape.

Two threaded

end caps of Carb-I-Tex 500 are f a b r i c a t e d with spanner wrench holes t o permit ease of i n s t a l l a t i o n (Fig. IV-1).

3.

Capsule S u i t a b i l i t y t o S o l i d Fuel Forms TO

f a c i l i t a t e f u e l i n g operations with s o l i d h e 1 forms, design of

t h e l i n e r f o r edge-type TIG weld c l o s u r e a t t h e hemispherical cap can be provided (Fig. I V - 4 ) .

This l i n e r configuration permits t h e f i n a l

weld t o be performed a t a remote l o c a t i o n from t h e f u e l .

I t would a l s o

allow t h e use of r i g h t c y l i n d r i c a l p e l l e t s having one diameter only, with t h e exception of t h e hemispherical ends,

INSD-2650-29 IV-14


FIG. IV-4. WELD.CLOSuREDETAIL

INSD-2650-29 IV-15


C.

STEADY STATE TEMPERATURF: PROFILES

The b a s e l i n e heat source design employs a Carb-I -Tex 500/502

g r a p h i t e heat s h i e l d .

Alternate designs have been considered which

employ a superalloy, v i b r a t i o n - i s o l a t i o n metal sleeve between t h e heat s h i e l d and thermoelectrics and which consider POCO g r a p h i t e f o r t h e heat s h i e l d .

These various combinations have been thermally analyzed

f o r both beginning-of - l i f e (BOL) normal operation and launch time

( 9 months after BOL) normal operation a t steady s t a t e conditions i n s i d e I n addition, t h e launch phase was s t u d i e d f o r two malfunction

t h e RTG. conditions

-

open c i r c u i t and normal l o a d e l e c t r i c a l operation, both

with a n RTG i n t e r n a l vacuum t o simulate l o s s of f i l l gas a t launch. An

85 node thermal model of t h e h e a t source was constructed, as

shown i n Fig. I V - 5 , model i s a 1 0 '

f o r s o l u t i o n with a d i g i t a l computer code.

The

p i e - s l i c e from one-half of a h e a t source, s i n c e circum-

f e r e n t i a l and a x i a l symmetry can be assumed. Material p r o p e r t i e s are handled by t h e code as constants; values used are given i n Table IV-3.

675 watts per RTG.

The a c t u a l thermal f u e l loading w i l l be

For conservatism, t h i s a n a l y s i s assumed complete

loading of t h e capsule

- 755 watts.

Results of t h e code runs are summarized i n Table I V - 4 .

The use

of a metal sleeve around t h e h e a t s h i e l d adds 50 t o 150°F t o a l l h e a t source temperatures, and a Carb-I-Tex h e a t s h i e l d r e s u l t s i n temperat u r e s about 70°F higher than with a POCO g r a p h i t e s h i e l d .

The conduc-

t i v i t y of gas contained within t h e generator changes with t i m e due


FIG. I V - 5 .

VIKING T H m L MODEL


MATERIAL PROPERTIES OF VIKING HEAT SOURCE

Component

Material

Conductivity (Btu/Hr-Ft- OF)

erature

Ehissivity (With Coatings)

Sleeve

Superalloy

12.

1300.

0.8

Heat Shield

POCO Graphite

35

1100.

0.8

Carb-I-Tex Graphite

3.5

lloo.

0.8

mP

Tantalum

42.

1600.

0.8

Compliance Member

Tantalum F e l t

0.06

1400.

0.8

Strength Member

TZM

85

1600.

0.8

Liner

Mo 50 w/o Re

42.

1300.

0.8

G a s Fill

Argon

0.025

1300.

35% H e l i u m , 65Q Argon

0.086

1300.

Fuel

Pu23% crospheres Argon Gas F i l l

35% Helium- 65% A r g o n

0.73

1700. t o 2600.

0.65 t o 1.25

1700. t o 2600.

0.30 t o Fill


TABU I V - 4 MAXIMUM STEADY STATE TEMPERATURES OF VIKING

Configuration

Time

Cup

Compliance Member

Strength Member

Liner

Fuel

(9) (OF)

(OF)

(OF)

(OF>

(OF>

(OF)

1114 1065

1231 1188

1204 1158

1434 1365

1708 1615

1814 1662

2776 2618

---

1129 1085

1108 1061

1355 1279

1642 1539

1733 1588

2721

1177 1134

1512 1396

1311 1204

1554 1438

1833 go6

1911

2903

1332 1295

1163 1115

1424 1358

17-17 1634

1806 1682

2801

1441

1259

1489

1757

1889

2891

1677

1492

1699

1948

2048

3042

Sleeve

Poco heat s h i e l d with sleeve

BOL Launch

Poco heat s h i e l d without sleeve

BOL Launch

Carb-I-Tex heat s h i e l d with sleeve

BOL Launch

Carb-I-Tex heat s h i e l d without sleeve

BOL Launch

Carb-I-Tex heat s h i e l d without sleeve; vacuum gaps; normal e l e c . load

Launch

----

Carb-I-Tex heat shield without sleeve; vacuum gaps; open c i r c u i t

Launch

--

Heat Shield

HEAT SOURCE

1753

2548 2726

2658


t o t h e a d d i t i o n of helium from f u e l decay.

This change i n gas Composi-

t i o n r e s u l t s i n lower temperature drops a c r o s s gaps a t launch t i m e than e x i s t immediately after f u e l i n g . A l l r e s u l t a n t component temperatures for normal generator opera-

t i o n a r e acceptable f o r a n IIHS.

The vacuum malfunction cases appear

t o be acceptable, even f o r s u p e r o r b i t a l re-entry, but f u r t h e r a n a l y s i s and t e s t a r e necessary f o r confirmation.

Fuel c e n t e r l i n e data f o r t h e

vacuum cases presumes t h e use of a f i n i t e pressure drop f i l t e r i n t h e l i n e r i n order t o r e t a i n a gas environment i n t h e fuel. D.

LAUNCH PAD ABORT STUDY

An a n a l y s i s w a s performed t o evaluate t h e response o f t h e nuclear system t o t h e more severe environments r e s u l t i n g from accidents occur-

ring a f t e r booster f u e l i n g and RTG payload i n s t a l l a t i o n . Centaur Vehicle was assumed f o r t h e a n a l y s i s .

A Titan I I I C /

A representative f a i l u r e

p r o b a b i l i t y of 1/100 was assumed f o r those f a i l u r e modes leading t o p r o p e l l a n t explosion and a subsequent f i r e b a l l .

The o b j e c t i v e was t o

determine t h e t o t a l p r o b a b i l i t y of f u e l release i n t h e launch pad environs, taking i n t o account t h e more severe environments t h r e a t e n i n g

f u e l containment and t h e c a p a b i l i t y of t h e system t o withstand t h e s e environments. The a n a l y s i s employed techniques developed during t h e SNAP 29 program and u t i l i z e d i n t h e PSAR (Ref. IV-1) f o r t h a t system. Included i n t h i s study were t h e e f f e c t s of b l a s t and shrapnel environments,

INSD-2650-29

IV-20

I


U impact of fuel-bearing systems on the surface of the pad, and the thermal environment. The analyses were separated into two phases: environment characterization and system response. 1.

D

Environment Characterization Each environment considered depends, in part, upon one or more of

the vehicle parameters including propelhnt weight, propellant type, and the separation distance between payload and fuel-oxidizer inter-

3

faces.

These parameters are shown in Table IV-5.

TABLE IV-5

VEHICLE PARAMETERS-TITAN IIIC/CENTAUR

a a

Vehicle Titan

J

Distance-RTG to Fuel/Oxidizer Interface( ft)

Stage -

Propellant

0

UTP-3001(solid)

843,000

83

86

1

N2H4/N2H2/UDMH

251,000

73

96

66,coo

23.3

50

29,858

30

25

2

7

Stage Propel-lant Dimension Weight ( l b s.) (ft> -

Centaur

-

P/L

-

I1

20

The mechanical and thermal environments are treated chronologically in the following sections. a.

Blast

Reference IV-1 presents a review and statistical analysis of data pertaining to blast overpressures, includfing dependence on fuel weight, separation distance, and the relationship between rocket propellants

INSD-2650 -29 IV-21


and TIVT,

An empirical equation f o r TIVT explosions was determined r e l a -

t i n g maximum overpressure ( p s i ) t o d i s t a n c e and explosive weight of the form:

where Wf = weight of TNT ( l b s ) and R = separation d i s t a n c e ( f t )

It was f u r t h e r determined t h a t t h e same equation could be employed t o describe overpressures from detonation of l i q u i d rocket p r o p e l l a n t s provided that Wf i s i n t e r p r e t e d as t h e weight equivalent of TNT f o r each type of p r o p e l l a n t ,

S t a t i s t i c a l treatment of t e s t data showed

mean weight equivalency f a c t o r s of f o r hypergolics.

4.5% f o r cryogenic f u e l s and 0.51%

S t a t i s t i c s were c o n s i s t e n t with "one-sided" normal

d i s t r i b u t i o n s with standard d e v i a t i o n s of

5.6 and 1.2$, r e s p e c t i v e l y .

These equivalence f a c t o r s and variances were employed t o d e t e r mine peak overpressure values a t t h e payload f o r t h e appropriate s t a g e s

of t h e reference vehicle. and

Table

IV-6 shows t h e mean, 90 p e r c e n t i l e

99 p e r c e n t i l e TNT equivalence f o r each s t a g e .

Table I V - 7 p r e s e n t s

t h e c a l c u l a t e d overpressures, t h e p e r c e n t i l e values here i n d i c a t i n g t h e p r o b a b i l i t y of not exceeding t h e t a b u l a t e d overpressure.

It i s seen t h a t t h e Centaur s t a g e y i e l d s t h e predominant overpressures because of i t s cryogenic f u e l and proximity t o t h e payload region.

The p e r c e n t i l e s i n d i c a t e t h e p r o b a b i l i t y of not exceeding a

given overpressure.

Thus t h e p r o b a b i l i t y of not exceeding 500 p s i

within t h e payload from a Centaur explosion i s 0.99.

INSD-2650-29 IV-22

I


3

TABLE I V - 6 TNT EQUIVALENTS FOR VEEICULAR STAGES

a J

a

Til" Equivalence ( l b s , )

Stage

Mean

9%

9%

Centaur

1340

3500

5280

Titan ( 2 )

337

1317

2122

T i t a n (1)

1280

5005

8060

a s

a

TABLE

IV-7

OVERPRESSURFS AT PAYLOAD ( p s i ) Stage

Mean -

90%

9%

C e n ta u r

200

386

500

Titan (2)

20

49

68

T i t a n (1)

13

32

45

a

1

I

-

INSD 26 50 -2 9 IV-23


b.

Shrapnel

The shrapnel environment during a launch pad explosion i s not w e l l defined.

An estimate of t h e environment was obtained (Ref. IV-1) from

data derived from t h e command d e s t r u c t of 2 s o l i d rocket motors.

The

environment w a s represented by about 230 pieces of aluminum about 1 f o o t

1/4" t h i c k w i t h a speed of about 300 f t / s e c , each piece of shrapnel t h e r e f o r e having a k i n e t i c energy of about 5 x 103 f t - l b a t square by

impact

.

The p r o b a b i l i t y of shrapnel impacting a given configuration (RTG, h e a t source, o r capsule) was determined by assuming t h e shrapnel t o be distributed isotopically in space f r o m the point of explosion.

The

impact p r o b a b i l i t y then, i s given by:

where

and

N

= number of fragments

R

= d i s t a n c e from source of explosion

A

= i n t e r a c t i o n area f o r a p a r t i c u l a r configuration; i . e . , t h e t o t a l area such t h a t i f t h e c e n t e r of mass of a fragment p e n e t r a t e s t h i s a r e a , impact with t h e t a r g e t w i l l occur.

The number of fragments, N, f o r t h e Centaur vehicle was assumed t o be

about 70.

Under t h e above assumptions, t h e Centaur vehicle contributed

t h e l a r g e s t shrapnel impact p r o b a b i l i t y i n t h e payload region. c.

Impact

It i s shown i n t h e next s e c t i o n t h a t t h e most l i k e l y r e s u l t of a p o s t u l a t e d b l a s t with overpressure as described above would be t o free

INSD-2650-29 IV-24


J 1

73

t h e h e a t source, leading t o subsequent impact i n t h e launch pad v i c i n i t y . An a n a l y s i s based on a survey of l o c a t i o n s of t h e payload a f t e r 30 launch

-l

d

a a

a b o r t s was coupled with estimates of t h e r e l a t i v e f r a c t i o n s of these a r e a s comprising sand, concrete and s t r u c t u r e .

The r e s u l t i n g p r o b a b i l i t y

estimates (Ref. I V - 1 ) were 0.65, 0 . 3 and 0.05, r e s p e c t i v e l y , f o r t h e s e impact media.

The impact speed w a s c a l c u l a t e d t o be about 100 f t / s e c .

The c a l c u l a t e d f a l l t i m e was about 2.8 seconds.

Therefore, f o r any

case i n which t h e f u e l capsule i s p r o t e c t e d by a heat s h i e l d u n t i l

a

impact, t h e capsule i s assumed t o be exposed t o t h e thermal environment

a t 2.8 seconds from t h e f i r e b a l l i n i t i a t i o n . d.

Thermal environment

The thermal environment considered i n t h i s study comprises t h r e e

3

p a r t s , t h e f i r e b a l l contributed by 1iquj.d p r o p e l l a n t s and t o a l e s s e r e x t e n t by s o l i d s , t h e r e s i d u a l s o l i d f i r e c h a r a c t e r i z e d by l o c a l i z e d

r-7

chunks of burning p r o p e l l a n t , and t h e r e s i d u a l l i q u i d f i r e s f e d by pools of l i q u i d p r o p e l l a n t not consumed i n t h e f i r e b a l l .

D 3

The distri-

bution of thermal f l u x e s i n t i m e f o r each source i s discussed below. (1) F i r e b a l l

K i t e and Bader ( R e f . I V - 2 ) have presented a thermal f l u x

model as a f u n c t i o n of a dimensionless temporal argument s c a l e d accord-

u P

ing t o t h e cube r o o t of t h e p r o p e l l a n t weight involved.

Their r e s u l t s

are given i n Table IV-8. The model on which t h i s h e a t f l u x d i s t r i b u t i o n was derived assumed t h a t a l l a v a i l a b l e energy i n t h e p r o p e l l a n t quantity, W, i s

INSD-2650-29 IV-25


TABLE IV-8

H U T FLUX MODEL

22 t (sec)

t* =

w

4

2

(t*) B t u l f t - s e c

( I b s ) 1/3

0

400

0.2

3 32

0.4

2 90

0.6

267

0.8

247

0.9

242

1.o

225

1.2

197

1.5-

165

1.5+

43

1.7

26

1.9

13

-

INSD -2650 2 9 IV-26 -.

.

._

-

-

- - -


i(rs

U D

consumed.

It, t h e r e f o r e , r e p r e s e n t s a maximum h e a t f l u x d i s t r i b u t i o n

f o r a given amount o f p r o p e l l a n t ,

Results of r e c e n t tests completed

under t h e PYRO Program ( t o be r e p o r t e d by Bader) t e n d t o confirm t h e maximal n a t u r e of t h i s thermal d i s t r i b u t i o n .

P

Based on a q u a l i t a t i v e review of t h e s e r e s u l t s we have assumed

and t h e d e n s i t y f o r f i s uniformly d i s t r i b u t e d between 0.2 and 1.0.

P

According t o t h i s assumption t h e h e a t f l u x d i s t r i b u t i o n corresponding t o any p e r c e n t i l e thermal p r o f i l e decreases more r a p i d l y i n t i m e t h a n t h e maximal p r o f i l e . The c o n t r i b u t i o n of s o l i d p r o p e l l a n t s t o t h e f i r e b a l l has

been t r e a t e d previously by assuming a n equivalence of s o l i d and l i q u i d

P

propellants,

A recent analysis ( R e f . IV-3)

of data developed under

P r o j e c t SOPHY y i e l d e d an equivalence f a c t o r , derived from t h e e f f e c t of s o l i d p r o p e l l a n t s on f i r e b a l l duration, of 0.25.

W e have used t h i s

equivalency i n p r e d i c t i n g h e a t f l u x d i s t r i b u t i o n s c h a r a c t e r i z i n g t h e T i t a n IIIC/Centaur f i r e b a l l . (2)

Residual l i q u i d f i r e s

The a n a l y t i c a l model i n c o r p o r a t e s a constant h e a t f l u x followi n g t h e conclusion of t h e f i r e b a l l of

13 B t u / f t 3 -sec f o r

a 30 minute

interval.

(3) Residual s o l i d f i r e s

P

0

S o l i d p r o p e l l a n t i g n i t i o n t e s t s conducted under t h e SNAP 29 Program y i e l d e d h e a t f l u x e s o f about 50 Btu/ft2-sec.

INSD-2650-29 IV-27

i n c i d e n t on t y p i c a l


This value a p p l i e s t o t h e heat

h e a t source and capsule configurations.

t r a n s f e r r e d through t h e surface of specimens, which surfaces were observed t o be coated with varying thicknesses of aluminum oxide.

Burn

times f o r t h e one cubic f o o t p r o p e l l a n t samples were q u i t e uniform a t about

4 minutes. While t h e p r o p e l l a n t used i n t h e s e t e s t s w a s A l G O 1 - I I B (Scout

vehicle p r o p e l l a n t ) , t h e T i t a n s o l i d (UTP-3001) i s not expected t o present a more severe environment. by about

40$, however,

The heat f l u x i s expected t o increase

t h e burning rate i s about 3 t i m e s as g r e a t .

Thus

t h e i n t e g r a t e d h e a t f o r t h e same s i z e p r o p e l l a n t block would be reduced for the T i t a n fuel.

From an a n a l y s i s of p r o p e l l a n t fragments from an SRM detonat i o n (described i n Ref. I V - 1 )

, the

most probable p r o p e l l a n t thickness

t o be encountered ( i f a t a l l ) by a h e a t source f a l l i n g t o launch pad would be about one f o o t , t h e same thickness used i n t h e SNAP 29 t e s t s . T h e fragmentation a n a l y s i s was a l s o used t o estimate t h e

p r o b a b i l i t y that a f a l l i n g h e a t source o r capsule would impact i n c l o s e proximity t o a burning fragment.

The d i s t r i b u t i o n of fragments

was represented by a s p a t i a l d i s t r i b u t i o n of t h e form r e-CYr2,where

r i s t h e distance measured from t h e base of t h e vehicle. C Ywas ,

The parameter,

determined from t h e expected s e p a r a t i o n distance f o r a f r a g -

ment of a given mass. This d i s t r i b u t i o n was then combined with t h e aforementioned information on payload depositions following a b o r t s t o determine t h e

INSD-2650 -29

1v-28


‘ 7

proximity p r o b a b i l i t y .

The r e s u l t was 0.1, which d i d not depend

s e n s i t i v e l y on the choice of t h e parameter a

II

.

I t was f u r t h e r estimated

t h a t i n only one case out of t e n t h e impacted fragment would be i g n i t e d . The t o t a l p r o b a b i l i t y of depositing a heat source on a burning p r o p e l l a n t

fragment w a s t h e r e f o r e about 1/100.

u

Based on t h e s e arguments t h e thermal environment t o which a capsule or h e a t source would be exposed i f it impacted on o r n e a r a chunk of s o l i d p r o p e l l a n t w a s assumed t o be a h e a t f l u x of 50 B t u / f t 2 -see f o r a f o u r minute i n t e r v a l . The t o t a l thermal environment i s summarized i n Fig.

IV-6.

The f i r e b a l l heat f l u x e s occurring l e s s than about f i v e seconds a f t e r

1

detonation are shown as 50 and

99%

p e r c e n t i l e curves.

The

99 p e r c e n t i l e

curve w a s formed from t h e “maximum” h e a t f l u x curve described by Kite and Bader using t h e 55$ equivalence evaluated f o r s o l i d p r o p e l l a n t s 1

contributing t o the f i r e b a l l . l e d t o i t s being considered

2.

a

The variance i n t h i s equivalence f a c t o r

99 p e r c e n t i l e ( r a t h e r than 100 p e r c e n t i l e ) .

Heat Source Response T h e reference system s e l e c t e d f o r a n a l y s i s i n response t o t h e

aforementioned environments c o n s i s t s o f a f u e l capsule enclosed i n a hexagonal prism of Carb-I-Tex 500/502 g r a p h i t e and i n s e r t e d i n t h e

a

RTG.

The f u e l capsule i s a c y l i n d r i c a l element

3.16 inches long with

an OD of 2.4 inches, and hemispherical end caps 2.4 inches i n OD. hexagonal h e a t source has a p r o j e c t e d area of 3.5 x 6.75 i n . 2 p r o j e c t e d a r e a f o r t h e generator i s about

INSD-2650-29

IV-29

6.5 x 10 i n . 2

.

.

The

The


INSD-2650-29 IV-30


Table IV-9 lists materials and approximate thicknesses used in evaluating resistance to mechanical environments, particularly shrapnel

u

impact

. TABLE

IV-9

RTG COMPONENTS CHARACTERISTICS

n 3

a

Material

Component

of Modulus Elasticit ._

(lb/in )

Psi x 10-

Liner

Mo 50 w/o Re

0.030

0.494

40.0

Strength Member

TZM

0.080 (avg.)

0.37

32.0

Heat Shield

0.051

1.9

Hot Shoes

Graphite 0.450 Carb-I-Tex 500/502 Ni with Fe insert 0.050

0.321

30.0

T/E elements

TAGS/PbTe

0.500

0.252

10.0 (avg.)

0.80

0.098

10.0

0.1

0.064

6.2

Cold Ehd Hardware A 1 Housing Mg

II I

Thickness (in)

Densits

a.

Blast response

To evaluate structural response to a blast loading, the pressure

history must be defined. From the detonation, a shock wave is formed consisting of a mass of compressed air propagating toward the target

3

or structural configuration. When the blast wave encounters the structure, the wave is diffracted. The pressure loadings developed by reflected shock waves during this diffraction phase are considerably

1

higher than the side-on overpressure. The time duration of the diffraction phase is very short and often is a cause of failure in high frequency structures.

INSD-2650-29

IV-31


A f t e r t h e b l a s t wave has enclosed t h e s t r u c t u r e , t h e pressure d i f f e r e n t i a l between f r o n t and back faces of t h e s t r u c t u r e decreases considerably.

This loading i s a f u n c t i o n of t h e dynamic pressure and

i s termed t h e drag phase.

The t i m e d u r a t i o n i s r e l a t i v e l y long and

damage i s p r i m a r i l y t o low frequency s t r u c t u r e s .

For t h i s study, i n t e r e s t i s d i r e c t e d toward p o t e n t i a l b l a s t e f f e c t s on t h e generator, f u e l block and capsule.

Table IV-10 constain t h e

considered overpressures and t h e i r p r o b a b i l i t i e s of not being exceeded.

TABLE I V - 1 0 SELECTED OVERPRESSURES Side-on Overpressure, ps (Psi)

Probab ilit y

(9)

200

50

386

90

500

99

By applying t h e conservation of mass, energy, and momentum a t t h e shock f r o n t , t h e r e f l e c t e d p r e s s u r e , p,,

and dynamic p r e s s u r e , q, may be

obtained as a f u n c t i o n of side-on overpressure, ps, and ambient p r e s s u r e , po.

7 Po 7 Po

r

+ 4P,

+ P,

2

=

-3-[

1 7p0

+

ps

INSD-2650 -2 9 IV-32

1


n

s i n c e we a r e i n t e r e s t e d i n launch pad a c c i d e n t s , po = 14.7 p s i .

IV-11 below p r e s e n t s p

r

Table

and q f o r each overpressure of i n t e r e s t . TABm I V - 1 1

EFLECTED AND DYNAMIC PRESSURES

200

386

2601

500

2488 To determine whether or not s t r u c t u r a l damage i s expected from a

b l a s t loading, Fig. I V - 7 from R e f .

IV-4 i s used.

It i s seen t h a t

p o t e n t i a l s t r u c t u r a l damage i s a f u n c t i o n of t h e s t r u c t u r a l frequency, f , depth o r l a t e r a l dimension of t h e s t r u c t u r e , D,

PS ,

and s t r u c t u r a l d e n s i t y , p

f , D, and p

.

side-on overpressure,

From the s t r u c t u r e of i n t e r e s t with

known, t h e c h a r t i n d i c a t e s t h e m i n i m u m side-on overpres-

sure t o i n i t i a t e f a i l u r e .

Table IV-12 p r e s e n t s p r o p e r t i e s of t h e

components of interest.

From Fig. I V - 7 and t h e c a l c u l a t e d frequencies and c h a r a c t e r i s t i c l a t e r a l dimensions, a l l t h r e e components r e p r e s e n t d i f f r a c t i o n t a r g e t s ; t h a t i s , p o t e n t i a l f a i l u r e would be due t o t h e r e f l e c t e d p r e s s u r e .

For

t h e h e a t s h i e l d it can be seen that t h e c r i t i c a l side-on overpressure i s about 150 p s i .

R

The housing p r o p e r t i e s y i e l d a c r i t i c a l side-on

overpressure of approximately 70 p s i .

INSD-2650-29

IV-33


I

TABLE 1v-12 PROPERTIES FOR BLAST W G E DETERMINATIO'N

Item

Material

Frequency f

(CPS) Capsule Strength

TZM

908

Heat Shield

Carb-I-Tex 500/502

5930

Housing

MgTh

Characteristic f-D Dimension Product D fD

(ft)

Structural D e n s i t y

(Wt/Volume Envelope)

(fP4

(#/f%)

170

.35

0.5

2965

39.8

0.66

6790

6.53

1875

10-3

Member

10,300

TRANSITION PROBABILITIES FOR BLAST ENVIROrJMENT

Initial- Sta t e Heat Source

m

0

- 0.99

F i n a l State Bare Camule

-

0.01

Fuel Release - 0


n

/'

P

a a 1 INSD-2650- 29

IV-35


n It does not appear that t h i s c r i t e r i o n of b l a s t damage i s a p p l i cable f o r the capsule s t r e n g t h member.

The c a l c u l a t e d n a t u r a l frequency

i s assumed very conservative (low) s i n c e t h e l i n e r enhances t h e capsule r i g i d i t y and t h e f u e l provides a highly e f f e c t i v e e l a s t i c foundation t o r e s t r i c t v i b r a t i o n a l response,

Therefore, a more r e a l i s t i c approach

involves estimation of material y i e l d i n g .

The TZM s h e l l a t temperature

has a y i e l d s t r e n g t h of approximately 40000 p s i and t h e maximum R/t of the shell i s

14.5.

Thus, t h e a p p l i e d pressure p r i o r t o y i e l d i n g i s

P =

4oOo0 14.5

= 2760 p s i

The maximum amplitude f o r a t r i a n g u l a r acceleration-time pulse i s about

1.4.

Therefore, t h e c r i t i c a l r e f l e c t e d pressure i s g r e a t e r than

2760/1.4 o r 1970 p s i . of 321 p s i .

T h i s value corresponds t o a side-on overpressure

This value i s conservative (low) f o r t h r e e important

reasons : (1) The r a t i o of model frequency t o t h e impulse frequency

must be a unique value t o warrant t h e maximum amplitude response (2)

.

The material has s u f f i c i e n t d u c t i l i t y t o i n d i c a t e s i g n i f i c a n t p l a s t i c flow or deformation p r i o r t o failure.

(3)

The f u e l will respond t o t h e loading imposed.

From these c a l c u l a t i o n s we conclude that the r e s u l t of exposing the i n i t i a l l y i n t a c t RTG would be t o f r a c t u r e the e x t e r n a l housing.

INSD-2650-29

IV-36

P


..

.

It i s f u r t h e r assumed t h a t by v i r t u e of t h e laminar construction of t h e Carb-I-Tex 500/502 heat s h i e l d member, and t h e expected a t t e n u a t i o n of t h e shock pulse, t h a t t h i s member would remain i n t a c t a f t e r t h e b l a s t except i n about 1% of t h e p o t e n t i a l b l a s t magnitudes.

The capsule was

assumed t o have a n i l p r o b a b i l i t y of f a i l u r e from b l a s t e f f e c t s .

From

t h e s e arguments the t r a n s i t i o n - p r o b a b i l i t y matrix f o r t h e b l a s t environment has t h e form o f Table I V - 1 3 . b.

Shrapnel p e n e t r a t i o n

Shrapnel p e n e t r a t i o n p r o b a b i l i t i e s were determined using t h e model developed i n R e f . IV-1. Fragments s e l e c t e d as t h e most l i k e l y t o cause p e n e t r a t i o n by v i r t u e of t h e i r number and k i n e t i c energy were 12 x 12 x

1/4 inches

3 i 3

a

i n s i z e and were assumed t o be aluminum.

Two impact o r i e n t a t i o n s were considered, "wedge" a n d "edge, f i n e d by sketch i n Fig. I V - 8 . o r i e n t a t i o n s such t h a t 8 = l e s s than 22.5'

de-

Wedge impact was defined f o r impact

45' +

Edge impact corresponds t o 8

22.5'.

and g r e a t e r than 67.5O.

The c r i t i c a l impact angle

assumed f o r p e n e t r a t i o n of the heat source and capsule were l i m i t e d t o p l a n a r angles,

9

, within

22.5O o f

goo.

A l l shrapnel o r i e n t a t i o n a t

impact were assumed equally l i k e l y so t h a t both wedge and edge impact had equal p r o b a b i l i t i e s a t about 1/8 (1/2 f o r e i t h e r edge or wedge and

1/4 for

c r i t i c a l angles) i f t h e fragment's c e n t e r of mass were d i r e c t e d

1

within t h e p r o j e c t e d a r e a of t h e t a r g e t .

An a d d i t i o n a l i n t e r a c t i o n

a r e a w a s defined as shown i n Fig. IV-9 so t h a t , i f t h e c e n t e r of mass of t h e p r o j e c t i l e were d i r e c t e d within t h i s area, proper impact f o r p e n e t r a t i o n would occur i n about

1/16

of t h e cases.

INSD-2650-29

IV-37


FIG. IV-8.

IMPACT ORIENTATIONS

6 in.

_--

-1

I

I 6 in.

c.t-C

. I I .

1 I I

I

I

I

Pro j ect e d t a r g e t area

I

I

b Additional

I

I

i n t e r a c t i o n area

L---------

FIG. IV-9.

TARGET AND INTERACTION AREAS

IV- 38


From these hypotheses and t h e 70 shrapnel fragments assumed f o r t h e Centaur v e h i c l e , t h e p r o b a b i l i t y of impacting t h e capsule or heat source w i t h proper o r i e n t a t i o n f o r p e n e t r a t i o n i s about 1 . 0 x The p r o b a b i l i t i e s f o r generator and heat source impact without regard

to fragment o r i e n t a t i o n are about 0.03 and 0 . 0 2 , r e s p e c t i v e l y . Energy requirements f o r p e n e t r a t i n g the m a t e r i a l s comprising each configuration of i n t e r e s t (RTG, h e a t source and capsule) were d e t e r mined using t h e following expressions f o r impact p e n e t r a t i o n for t h e

wedge and edge o r i e n t a t i o n s from Ref. IV-1:

a 3 i

where

P = p e n e t r a t i o n (in3

t = p r o j e c t i l e thickness (in9 = 0.25 E = k i n e t i c energy of p r o j e c t i l e ( f t - l b ) =

J

p

P

= p r o j e c t i l e density (g/cc)

-

5 x 103

2.7

Pt = t a r g e t d e n s i t y A = length of p r o j e c t i l e edge x thickness ( i n2 ) = 0.6

%

= modulus of e l a s t i c i t y of t a r g e t ( p s i )

Equivalent g r a p h i t e thicknesses f o r each component were obtained by s c a l i n g a c t u a l component thicknesses by d e n s i t y and e l a s t i c modulus

correction ratios.

The equivalent g r a p h i t e thicknesses and energy

i

INSD-2650-29

IV-39


requirements t o p e n e t r a t e t h e s e v e r a l configurations are given i n Table I V - 1 4 . TABLF: I V - 1 4 ENERGY REQUIREMENTS FOR RTG COMPONENT PENETRATION

Component

P Equivalent ( i n ) Wedge me

Capsule

0.79

5.6

3.2 x

Heat Shield

0.45

0.45

1.0 x 10

5.4 x 103

RTG (without Heat

4.26

14. 1

9.2 x i o

1.7 x

Energy Required ( f t - l b ) Wedge

me

io3

6.7 x

i o4. -

105

Source) From t h e s e r e s u l t s and t h e c h a r a c t e r i s t i c shrapnel energy of

5 x lo3

f t - l b s , it w a s concluded t h a t capsule or heat source impact i n

t h e wedge o r i e n t a t i o n would most l i k e l y cause f u e l r e l e a s e , whereas a n i n i t i a l l y i n t a c t RTG would provide ample p r o t e c t i o n for t h e capsule and heat s h i e l d .

Impact a t any o r i e n t a t i o n on t h e RTG w a s assumed t o

r e l e a s e t h e i n t a c t h e a t source. These r e s u l t s were used t o determine t h e t r a n s i t i o n p r o b a b i l i t y matrix f o r shrapnel impact, given i n Table I V - 1 5 . TABLE I V - 1 5 TRANSITION MATRIX

Initial State

m

IirPG

0.97

Heat Source

-

Final State Capsule

Heat Source

0.03 0.98

INSD-2650 -29 IV-40

F u e l Release

-0

0.02

-1

Capsule

\

SHRAPNEL

0

1 1 x 10-3


n

c.

Impact response

No analyses were performed f o r impact response of t h e p o s s i b l e configurations.

I t i s believed, however, t h a t t h e following conclu-

s i o n s a r e i n t u i t i v e l y reasonable,

0

If t h e generator impacts any of t h e t a r g e t media ( sand, concrete,

or s t r u c t u r e ) , it i s assumed that t h e housing w 11 f a 1 and deform, exposing t h e heat source. t o remain i n t a c t .

The Carb-I-Tex 500/502 heat s h i e l d i s assumed

From t h e r e s u l t s of t h e preceding shrapnel penetra-

t i o n a n a l y s i s , t h e s e same r e s u l t s would apply even i f t h e RTG were t o

J 1

impact sharp s t r u c t u r e . If t h e heat source impacts sand o r concrete, it is assumed t h a t the

Carb-I-Tex 500/502 heat s h i e l d w i l l survive t h e impact except i n one out

J 3

of 10 cases.

F a i l u r e of t h e capsule i s assumed i f it encounters sharply

pointed s t r u c t u r e .

This p r o b a b i l i t y was estimated a t one i n a thousand,

using t h e t o t a l s t r u c t u r a l impact p r o b a b i l i t y of f i v e i n one hundred and assuming a one i n f i f t y chance f o r impact on sharp s t r u c t u r e a t o t h e r than glancing o r i e n t a t i o n s . Impact o f t h e bare capsule would produce f a i l u r e only i n t h e event

of impacting sharp s t r u c t u r e ; i t s p r o b a b i l i t y i s the same as t h a t f o r

a a

t h e heat source, 10-3. The t r a n s i t i o n matrix derived from t h e s e considerations i s given i n Table

IV-16.

J

INSD-2650-29 IV-41


TABLE

IV-16

TRANSITION MATRIX

I n i t i a l State

Heat Source

Fuel Relc lse

-0

-1.0 0.9

Heat Source

0.1

-

Capsule

c?.

IMPACT

Final State Capsule

m -0

RTG

-

-0

- 10-3

1.0

T h e r m 1 response

Calculations have been performed f o r t h e bare capsule and heat source configurations. was a combination of t h e

The thermal p r o f i l e used i n the c a l c u l a t i o n s

99 p e r c e n t i l e

f i r e b a l l , r e s i d u a l s o l i d pro-

p e l l a n t burning and t h e r e s i d u a l l i q u i d p r o p e l l a n t mvironment.

Accord-

i n g t o t h e s t a t i s t i c a l model, a more severe environment t h a n t h i s would occur i n only 1 out of 10,000 cases. The r e s u l t s of c a l c u l a t i o n s f o r t h e unprotected capsule a r e shown i n Fig. I V - 1 0 .

The maximum l i n e r temperature was about 3250°F, occur-

r i n g s h o r t l y after t h e f i r e b a l l i g n i t i o n .

The e f f e c t of t h e s o l i d pro-

p e l l a n t environment was t o produce a peak l i n e r tem2erature of about 2750â&#x20AC;&#x2122;F.

The equilibrium l i n e r temperature during t h e r e s i d u a l l i q u i d

f u e l f i r e was about 2050%. Calculations performed for t h e same environment but with a CarbI-Tex shielded capsule r e s u l t e d i n a maximum l i n e r temperature of

2640% following t h e s o l i d p r o p e l l a n t environment and an equilibrium temperature of 21409 i n t h e l i q u i d p r o p e l l a n t r e s i d u a l f i r e .

INSD-2650-29 IV-42


1

1 1 1 1

1

1

1

1

k

a,

E

P

: k

I !

%

.8

t-4

Pi

CI

rn

c

a, k

In M

IV-43

INSD-2650- 29

m N

E

k a, Q

2

0 0 N

0 N


3.

Composite Launch Pad Abort P r o b a b i l i t i e s

a.

Summary r e s u l t s

The t r a n s i t i o n matrices may be combined m u l t i p l i c a t i v e l y t o d e t e r mine t h e p r o b a b i l i t i e s f o r each p o s s i b l e state; i . e . , RTG, heat source, capsule, o r f u e l r e l e a s e .

The formalism employed f o r any s e q u e n t i a l

arrangement of environments i s :

(1) Determine p r o b a b i l i t i e s f o r each f i n a l s t a t e from t h e previous environment. (2)

Calculate p r o b a b i l i t i e s f o r f i n a l states of t h e present environment according t o : i

i-1 Ti

P. = P J k

kj

where Ps i s t h e p r o b a b i l i t y of f i n a l configuration, j, from environJ i -1 ment i , P i s t h e p r o b a b i l i t y of f i n a l state k from the p r e c e d i x k i i s t h e t r a n s i t i o n p r o b a b i l i t y from configuraenvironment (i-1)and T kj t i o n s k t o j during environment. Applying t h i s formalism and including t h e launch pad a b o r t p r o b a b i l i t y of

(1) The p r o b a b i l i t y of exposing a h e a t s h i e l d capsule t o t h e f u l l launch pad thermal environment i s

8 (2)

10-3.

The p r o b a b i l i t y of exposing a b a r e capsule t o t h e f u l l thermal environment i s

(3)

-4

3 x 10 ,

The p r o b a b i l i t y o f exposing a b a r e capsule t o t h e post impact thermal environment only ( a f t e r 2.8 seconds) i s 1 x

INSD-2650-29 IV-44


1 1

y3

(4)

a

The p r o b a b i l i t y of f u e l r e l e a s e p r i o r t o t h e thermal environment i s about 2 x loe5.

(5)

The p r o b a b i l i t y of f u e l r e l e a s e r e s u l t i n g

s o l e l y from t h e thermal environment i s negligible. b.

Discussion

The t o t a l f u e l r e l e a s e p r o b a b i l i t y r e s u l t i n g f r o m t h e launch pad a b o r t environments w a s c a l c u l a t e d t o be about 2 x 10-5

.

This r e s u l t e d

from shrapnel p e n e t r a t i o n considerations and estimates of f a i l u r e due t o impacting sharp s t r u c t u r e .

The models used t o derive t h e s e r e s u l t s

are b e l i e v e d t o be conservative.

0

However, p a r t i c u l a r a t t e n t i o n should

be devoted t o improving models f o r shrapnel f l u , s i z e and shape d i s t r i b u t i o n s and p e n e t r a t i o n c h a r a c t e r i s t i c s at t h e r e l a t i v e l y s m a l l impact speeds of 300 t o 600 f e e t per second. B l a s t e f f e c t s will s i m i l a r l y r e q u i r e i n v e s t i g a t i o n .

However, it

i s believed t h a t a t t e n u a t i o n afforded by s t r u c t u r e surrounding t h e

u

RTG, ignored i n t h i s study, tends t o make t h e s e r e s u l t s conservative. There a p p a r e n t l y i s l i t t l e t h r e a t t o f u e l capsule i n t e g r i t y imposed

by thermal environments based on the results of c a l c u l a t i o n s under environments more severe than are a n t i c i p a t e d for t h i s vehicle.

E f f e c t s of

impairment of t h e mechanical p r o p e r t i e s of t h e f u e l capsule by t h e

u

b l a s t , shrapnel and impact environments p r i o r t o thermal exposure and chemical e f f e c t s w i l l r e q u i r e f u r t h e r study.

INSD-2650-29

IV-45


E.

FLIGHT ABORT STUDY

The e f f e c t s of a b o r t f l i g h t c o n d i t i o n s on t h e RTG environment and subsequent heat source performance a r e evaluated i n t h i s s e c t i o n t o i n s u r e h e a t source i n t e g r i t y .

The p r e s e n t a t i o n i s i n chronological

event o r d e r , so t h a t f i r s t t h e Viking f l i g h t conditions a r e studied t o d e f i n e t h e range of a b o r t r e - e n t r y angles and v e l o c i t i e s , and subsequently t h e r e - e n t r y environment and thermal e f f e c t s on the RTG a r e analyzed over t h e a p p r o p r i a t e parametric f i e l d s . 1.

F l i g h t P r o f i l e Analysis The most severe f l i g h t environment (not including the l a u n c h pad

a b o r t ) f o r heat source i n t e g r i t y c o n s i d e r a t i o n i s t h r u s t misalignment during t r a n s f e r from Earth o r b i t t o i n t e r p l a n e t a r y t r a j e c t o r y .

Thus,

a range of a l t i t u d e s and t h r u s t misalignments a r e applied t o t h e

escape" rocket s t a g e s .

II

These c a l c u l a t i o n s y i e l d t h e r e - e n t r y a n g l e s

and v e l o c i t i e s f o r h e a t source environment e v a l u a t i o n s . a.

Launch v e h i c l e s , events and conditions

A T i t a n IIIC/Centaur launch v e h i c l e c o n f i g u r a t i o n i s t o be used

f o r t h e Mars/Viking mission.

The Centaur v e h i c l e i s an improved

v e r s i o n with two s e p a r a t e burns being planned.

The T i t a n I I I C along

with the f i r s t burn of t h e Centaur w i l l be used t o p l a c e t h e v e h i c l e i n t o an e a r t h c i r c u l a r o r b i t .

The second burn of the Centaur then

s u p p l i e s t h e necessary v e l o c i t y increment t o escape the g r a v i t a t i o n a l p u l l of Earth.

INSD - 2650-2 9 IV-46


For these studies, the following initial conditions for the second Centaur burn have been assumed: Circular Orbit Altitude, h = 559,506 feet Inertial Velocity, Vi = 25,597 fps Inertial Flight Path Angle,

<i

=

0 degree

Weight, W = 34,350 lbs. 0

In addition, the following thrust characteristics are assumed. Thrust, T = 30,000 lbs. Specific Impulse, I

= 442 sec.

SP Fuel Mass Flow Rate, M =-2.11 slugs per sec. Burn Time, T = 322 seconds b The above values of initial condition and thrust characteristics are believed to be typical of current launch plans. A Centaur vehicle drag

a I J J

area (C S) of 157 square feet representing a nose on orientation was D used for all re-entry trajectory computations. 5.

Thrust misalignment analysis

T h r u s t misalignment failures such as considered here may be largely

prevented by launch constraints which prevent continuation of the mission when abnormal conditions are detected. An example of such control might be a case where the second Centaur burn is not initiated unless the vehicle is aligned within certain angular limits. However, the availability and/or reliability of such a launch program is not known at this time, and thus an investigantion of superorbital re-entry is pertinent.

INSD- 2 650- 2 9 IV-47


This study i s limited i n scope and thus should be t r e a t e d only a s a preliminary survey.

An i n i t i a l c i r c u l a r o r b i t i n a d i r e c t i o n due

e a s t about t h e E a r t h ' s equator i s assumed.

I n a complete a n a l y s i s , a

range of o r b i t a l d i r e c t i o n s and l o c a t i o n s would be i n v e s t i g a t e d f o r an oblate earth.

Other simplifying assumptions include the p o s i t i o n i n g

of t h e t h r u s t misalignment angle i n a plane normal t o t h e E a r t h ' s s u r f a c e , a s w e l l a s t h e passing of the t h r u s t vector through t h e v e h i c l e c e n t e r of g r a v i t y .

Variations t o such c o n s t r a i n t s can produce s i g n i f i -

c a n t changes i n t h e r e s u l t s and might r e s u l t i n conditions more severe than those c u r r e n t l y being obtained. Figures IV-11 and I V - 1 2 present r e - e n t r y t r a j e c t o r y values r e s u l t ing from t h e occurrence of v a r i o u s t h r u s t misalignments during a Centaur second burn.

R e l a t i v e r e - e n t r y v e l o c i t i e s and f l i g h t p a t h

angles a r e described f o r a l t i t u d e s ranging from 400,000 f e e t through 100,000 f e e t .

The t h r u s t misalignment angles a r e measured r e l a t i v e t o

t h e body x - a x i s which i s i n i t i a l l y aligned with t h e zero degree f l i g h t p a t h angle a x i s .

Negative values of t h r u s t misalignment i n d i c a t e

d i r e c t i o n s towards t h e E a r t h ' s s u r f a c e while p o s i t i v e v a l u e s i n d i c a t e d i r e c t i o n s away from t h e E a r t h ' s surface.

The d a t a of Fig. I V - 1 1 ,

which a r e f o r negative v a l u e s of t h r u s t misalignment, occur while t h e Centaur i s s t i l l t h r u s t i n g .

For a l l c a l c u l a t i o n s , the heat source i s

assumed s t i l l attached t o t h e Centaur v e h i c l e .

Figure I V - 1 2 p r e s e n t s

t r a j e c t o r y d a t a f o r p o s i t i v e v a l u e s of t h r u s t misalignment angle. These d a t a d e s c r i b e a c o a s t r e - e n t r y following a n e l l i p t i c a l f l i g h t path about t h e Earth.

INSD-2650-29 I V - 48


a a

3 J

INSD-2650-29 IV-49


INSD- 2 6 50 - 29

IV-50


2.

Re-entry T r a j e c t o r y Evaluation The r e - e n t r y s t u d i e s include the range of f l i g h t conditions caused

by o r b i t a l decay, t h r u s t misalignment, and an e s t i m a t e of terminal impact v e l o c i t y .

Since d e t a i l e d i n v e s t i g a t i o n of the s e p a r a t i o n of t h e

RTG from t h e v e h i c l e and subsequent exposure of t h e h e a t source w a s not

performed, t h e heat source i s assumed t o r e - e n t e r b a r e and unattached t o o t h e r s p a c e c r a f t components.

This assumption i s believed conserva-

t i v e , b u t confirming c a l c u l a t i o n s a r e required. a.

O r b i t a l decay t r a j e c t o r y

O r b i t a l decay t r a j e c t o r y r e s u l t s a r e shown i n Fig. IV-13.

The

i n i t i a l conditions a t 400K f e e t were: i n e r t i a l v e l o c i t y = 25,690 f p s , i n e r t i a l f l i g h t p a t h angle = 0 . 1 degree, and i n e r t i a l heading degrees (measured clockwise from t h e North Pole).

I .

120

The beginning of

continuum flow regime f o r t h e hexagonal heat source i s shown t o occur a t a n a l t i t u d e of approximately 200K f e e t and w a s based on t h e c r i t e r i o n described i n Ref. V - 2 :

i s t h e f r e e s t r e a m Knudsen number and P / p i s the d e n s i t y Kn s m r a t i o a c r o s s t h e shock wave. Since maximum heating r a t e occurs a t

where N

approximately 200K f e e t , t h e IIHS spends 1850 seconds out of 1950 seconds

3

of heating pulse i n f r e e molecule and t r a n s i t i o n a l flow regimes.

a I

1.3 .

a

.-

-.

.

INSD-2650-29 IV-51

.

._

.-


a . Histories of h, VR, YR and M

150'

, R' 100

0

200

I

400

600

800

1000

1200

1400

1600

1800

Trajectory Time from 400K Feet ( s e c )

Continuum flow regime, begins b.

Histories of qm, qs and qs Orbital Decay

250

-

.

N

* *

2

I

200

a, 3

v1

a, m

a

150

2 C

0"

2

100

4

j; W

G.

8'

50

Lr

0 11

0

Trajectory Time from 400K Feet ( s e d

FIG.

IV-13.

TRAJECTORYHISTORIES

FOR

INSD-2650- 29 IV-52

AN ORBITAL DECAY

2000


b.

Superorbital re-entry

Because of time c o n s t r a i n t s , the s u p e r o r b i t a l r e - e n t r y s t u d i e s were performed concurrently with f l i g h t p r o f i l e p r e d i c t i o n s .

Thus, parametric

i n v e s t i g a t i o n s of t r a j e c t o r y v a r i a b l e s were made t o include t h e values obtained from t h r u s t misalignment conditions.

The i n i t i a l r e l a t i v e

v e l o c i t i e s (V ) and r e l a t i v e f l i g h t path angles ( Y ) ranged between R

R

26,000 and 48,000 f t / s e c and -10 and - 3 degrees, r e s p e c t i v e l y .

(It

should be noted t h a t t h e s e r e p r e s e n t t h e most severe r e - e n t r y c o n d i t i o n s . ) A l l t r a j e c t o r i e s were i n i t i a t e d a t 400,000 f e e t and zero l a t i t u d e and

3

longitude with a heading of due East. The hypersonic b a l l i s t i c c o e f f i c i e n t (W/C A ) f o r the SNAP 19 heat

D

source i n side-on s t a b l e a t t i t u d e i s 34 psf and t h i s value w a s held constant throughout t h e t r a j e c t o r y down t o a Mach number o f 2 . 0 where aerodynamic h e a t i n g becomes small. Figure I V - 1 4 shows t h e t r a j e c t o r i e s i n v e s t i g a t e d and those t h a t re-entered d i r e c t l y . studied f u r t h e r .

The t r a j e c t o r i e s t h a t skipped out were not

It should be noted t h a t , although only d i s c r e t e p o i n t s

were s t u d i e d , l i m i t a t i o n s on t h e skip-out and d i r e c t t r a j e c t o r i e s may

be deduced. The maximum convective h e a t i n g r a t e s and t h e t o t a l s t a g n a t i o n convective energies f o r t h e parametric range a r e depicted i n F i g s . IV-15aY 15b and 15c.

The unusual boundaries (cusps) on t h e t o t a l

heating curves a r e a r e s u l t of t h e skip-out c u t o f f conditions.

It i s

apparent t h a t the maximum t o t a l heat supplied i s between 60,000 and

INSD-2650-29 IV-53


IV-54


7

3

IV-55

INSD-2650-29


a n

l.r3

P U 1

Q P IJ

a J

a J i 3 il 1

t

INSD-2650-29 IV-57


. 80,000 B t u / f t

2 while t h e maximum r a t e reached about 2000 B t u / f t 2 -sec.

A b r i e f i n v e s t i g a t i o n of r a d i a t i o n heating divulged t h a t l e s s than 10%

of t h e t o t a l heat i s contributed by r a d i a t i o n .

This s m a l l magnitude

together with time c o n s t r a i n t s j u s t i f i e d omission of f u r t h e r study of r a d i a t i o n heating and i t s e f f e c t on heat source temperature during re-entry. c.

Terminal impact v e l o c i t y

Re-evaluation of the subsonic drag c o e f f i c i e n t f o r t h e h e a t source y i e l d s an estimated impact speed, 245 f t / s e c .

The technique involved

c i r c u l a r c y l i n d r i c a l d a t a w i t h a s u r f a c e roughness f a c t o r t o a d j u s t f o r t h e hexagonal c r o s s s e c t i o n .

Although t h i s technique i s q u i t e

approximate, r e c e n t preliminary drop t e s t s conducted a t Tonapah Test Range by Sandia Corporation i n d i c a t e a terminal v e l o c i t y q u i t e c l o s e t o the c a l c u l a t e d value. 3.

Heat Source Thermal Response a.

Analysis

The two-dimensional heat source thermal model used i n t h i s study i s shown i n Fig. ID-16.

The model i s s u i t e d f o r s t u d i e s i n which t h e

r e - e n t r y mode i s side-on-spin o r side-on-stable edge leading.

For

conservatism, only t h e s t a b l e mode w a s evaluated s i n c e t h i s mode y i e l d s higher temperatures and termperature g r a d i e n t s than t h e spinning mode. Two d i f f e r e n t thermal s h i e l d m a t e r i a l s , POCO, type AXF (AS)-Q1,

and

Carb-I-Tex, type 5 0 0 , w e r e evaluated.

INSD -2650-29 IV-58

I


The thermal model was programmed for the TAP I11 computer code and solutions were obtained using an IBM 360-50 computer. The initial temperature varied from 1050°F at the external surface to 1825'F fuel core (Node 100).

at the

Surface recession rates caused by oxidation were

evaluated but the resultant dimensional change of the configuration was not factored into the temperature distribution calculations. Further study is required to determine the effects of surface recession on predicted temperatures. The heating distribution for the side-on stable attitude of the heat source was obtained from the re-entry trajectory studies described above together with test data from the SNAP 2 9 program. Figure IV-17 shows this distribution. b.

Results

Temperature histories of heat source components during an orbital decay re-entry (-0.1';

25,690 ftlsec) are shown in Figs. IV-18 and IV-19

for POCO and Carb-I-Tex 500/502 thermal shields, respectively. For either shield material, the fuel capsule liner is maintained below the melting temperature. Hence, either thermal shield material is considered adequate for orbital decay re-entries. Peak liner temperatures as a function of initial path angle and velocity for superorbital re-entry are presented in Figs. IV-20 and IV-21, respectively, for POCO and Carb-I-Tex heat shields.

It is clear

by comparing these figures that a Carb-I-Tex thermal shield provides substantially lower liner temperatures than a POCO shield.

In fact,

INSD- 2650- 2 9 IV-6 0 .-

-..-"-.

-

. . .

...

. .

.

.

.

.

-


P P

INSD-2650-29 IV-61


IV-62 ~.

. .

...

.

~

. . .. - . . ~ -


a 3 3

INSD- 2650- 29

IV-63


IV-64

INSD-2650- 29


u

a 3

INSD-2650-29 IV-65


the Mo-50% Re capsule liner has no apparent temperature restrictions for any of the orbital and superorbital re-entry conditions investigated if a Carb-I-Tex thermal shield is used (Fig. IV-21). Temperature distributions through Carb-I-Tex and POCO thermal shields are presented in Figs. IV-22 and IV-23 for a superorbital trajectory (initial flight path angle

=

-loo; initial velocity

=

30,000

ft/sec). 4.

Heat Shield Thermal Stresses A s a result of the geometric complexity associated with a hexagonal

heat shield and the thermal gradients in both the radial and circumferential directions, t h e precise determination of thermal stresses resulting from re-entry heating requires development of a digital computer program. However, by assuming a circular cross-section and approximating the temperature profile, an estimate of the maximum thermal stress is obtainable from simpler analytic methods. The hexagonal heat shield investigated has a distance across the flats and an inside diameter of 3.5 inches and 2.37 inches, respectively. The superorbital re-entry condition assumed for evaluation is based upon an initial velocity of 30,000 ft/sec., and an initial flight path angle of -loo at an altitude of 400,000 feet (the same conditions used in Figs. IV-22 and IV-23).

The orientation is related to a mode of

side-on-stable. Two graphitic heat shield materials were investigated:

POCO (type AXF), and Carb-I-Tex 500.

The solid lines in Fig. IV-24

present the inside and outside surface temperature profiles at the

INSD-2650-29 IV-66 - .

- --

- _ - __

.

-

-

__

-.

-

.

.


INSD-2650- 29


k.

,'

i

, , . ...

',


D

P

3

a

I

I

INSD-2650- 29 IV- 69

I

I


time of peak surface temperatures for the POCO graphite as determined from the results in the previous section. The surface temperatures for the structural model used here are also represented in the figure by the dashed lines. Note that temperature and geometric symmetry exists about the center line in the direction of the velocity vector. Upon investigating the two temperature profiles shown in Fig. IV-24

(i.e., solid and dashed lines), the major difference is seen to exist at the shaded areas. Fortunately, these two locations are at the corners where: a.

Surface recession will round off the corners, hence reducing the slopes of the temperature gradients

presented for these regions. b.

The peak temperatures are very local in nature and should not have any significant thermal stress effect.

The maximum thermal stress may be obtained for a cylindrical shell where the temperature varies around the section and through the thickness. Goodierâ&#x20AC;&#x2122;s approach(Ref. I V - 5 ) is to represent %h-e s u r f a c e %empemtures by simple Fourier series where the temperatures are,

T

=

1

T 0

A

0

= A/ 0

+ A

cos8 1

+ A/

1

cos e

+B +

sine

1

Bâ&#x20AC;&#x2122; sin 1

e

05 e 5

7~

(IV-1)

05e l n

(IV-2)

where T and T are the inside and outside temperatures (OF), respectively. 1 0 The extreme circumferential bending stress (psi) is given by the equation,

INSD-2650-29 IV-70


From the approximated temperature distributions shown in Fig. IV-24, the mathematical representation becomes,

P

T 1

P

u

= 2450

+

3900 + 2500

T

-750 sin 8

1150 c o s 8 COS

8

-1900 sin 8

0

Substituting into equation (IV-3),

Y

o = +

Ea 2 (1-v)

E1450

-

3

1350 cos0 -k 1150 s i n e

For POCO AXF at a mean temperature of -5000°F,

E = 0.9 x 106 psi = 4.8 x

1

in/in.-OF

v = 0.3 The maximum stress occurs at 6 =0' where the stress magnitude is 8640 psi.

J J

This stress value is of similar magnitude with the anticipated flexural capability of the material at elevated temperature. For example, the ultimate tensile strength at 6500OF is reported to be approximately 4000 psi, and the ultimate flexural strength of POCO graphite is usually about 2 or 3 times greater. A similar temperature profile was developed for a heat shield

manufactured from Carb-I-Tex 500 with the same initial conditions of

1 J

INSD -2650-2 9 IV-71


30,000 ft/sec. velocity and -10'

path angle. The comparison of front

and back face temperatures is given in Table IV-17. TABLE IV-17 FRONT AND BACK FACE TEMPERATURES OF HEAT SHIELDS POCO -

Carb-I-Tex Outside-Front

7900°F

7500°F

Inside-Front

4750°F

3600°F

Outside-Back

1600°F

1400°F

Inside-Back

1230°F

1300°F

From the temperatures shown, the AT profile is more severe for the Carb-I-Tex with a mean front face temperature for Carb-I-Tex of 6325'F as opposed to 5550°F for POCO AXF-Q1. At room temperature, E

cy

is (2.7 x lo6) (.61 x

or 1.65 psi/OF

for Carb-I-Tex versus approximately 7.7 psi/OF for POCO. Although values will be significantly lower at the temperatures of interest, equation

IV-3 indicates that the stress levels should be considerably lower in

.

the Carb-I-Tex

There are many forms of Carb-I-Tex where the cloth layup may have various orientations. Additional flexure capability is conceivable by including cross-grain needles to increase the ability to transfer shear across the laminates in flexure. From the preliminary calculations provided, it appears that the heat shield of either material must be further investigated since little

INSD - 2650- 2 9 IV-72 . .

...

.

. ...

.

..

.

-~ -

- .


Irs

is known of the mechanical properties of the two materials at elevated temperature. Thus, test programs are necessary to establish thermal shock capability of the configuration and determine elevated temperature strength. 5.

Flight Abort Event Probabilities The previous sections contain parametric data on the response of

the heat source to orbital and superorbital aborts.

Suborbital aborts

were not investigated owing to: (1)

R B

the lesser severity of the re-entry heating pulse compared with superorbital entry, and

(2)

the small probability of impacting land after exposure to re-entry heating.

a.

Orbital decay

For orbital decay conditions, the temperature histories contained

a

in Figs. IV-18 and IV-19 may be utilized to estimate failure probabilities. It should be noted that the conservative assumption implicit in using

these results is that stable, corner-on entry occurs. This is believed

a a 3

not to be the likely case; however, a detailed study beyond the present scope is required to prove this point. ability was taken as

Orbital decay occurrence prob-

lo'*.

The maximum temperatures of the capsule liner during entry are about 3400°F and 2900°F for the POCO and Carb-I-Tex heat shields. Use

a

of the error analysis results presented in Chapter V of Ref. IV-6 yields

I

capsule shielded by POCO AXF and Carb-I-Tex 500 graphite.

1

negligible failure probabilities for the TZM/molybdenum 50 w/o rhenium

INSD-2650-29 IV-73


The c a l c u l a t e d thermal stress i n the g r a p h i t e f o r e i t h e r h e a t s h i e l d i s about a f a c t o r of 3 o r 4 lower than t h e estimated f a i l u r e s t r e s s (10 t o 1 2 thousand p s i ) .

It i s concluded, t h e r e f o r e , t h a t d i s -

i n t e g r a t i o n of the h e a t s h i e l d , p o t e n t i a l l y permitting t h e r e l e a s e of f u e l a t high a l t i t u d e , i s q u i t e u n l i k e l y under o r b i t a l decay r e -

e n t r y conditions. The p r o b a b i l i t y of mechanical f a i l u r e of t h e POCO heat s h i e l d w a s taken a s

i n accordance w i t h t h e r e s u l t s of t h e IRHS a n a l y s i s

(Ref. I V - 7 ) .

For Carb-I-Tex, a n improvement f a c t o r of 10 ( f a i l u r e

probability =

w a s assumed owing t o i t s b e t t e r mechanical pro-

perties. b.

Superorbital re-entry

S u p e r o r b i t a l e n t r y c o n d i t i o n s expose t h e h e a t source t o more severe thermal/mechanical environments.

From a n a l y s i s of p h y s i c a l l y

p o s s i b l e r e - e n t r y modes, i t i s apparent t h a t r i s k assessment depends very s t r o n g l y on occurrence p r o b a b i l i t i e s f o r t h e necessary combinat i o n s of speed and r e - e n t r y angle t o produce t h e s e severe c o n d i t i o n s . Again, a l l c a l c u l a t i o n s were done f o r a s t a b l e e n t r y mode (corner-on) so t h a t r e s u l t s , p a r t i c u l a r l y f o r thermal s t r e s s e s i n t h e h e a t s h i e l d ,

a r e conservative. Figure I V - 1 8 p r e s e n t s r e s u l t s f o r t h e heat source containing a s h i e l d of POCO g r a p h i t e .

For r e - e n t r y speeds i n excess of 38,000 f p s

and r e - e n t r y angles s t e e p e r than about -7 degrees, the Mo-Re m e l t temperature i s approached.

These l a t t e r conditions a l s o cause tempera-

t u r e s c l o s e t o the melt p o i n t of t h e TZM s t r e n g t h member.

INSD - 2 650- 2 9 IV-74

Use of t h e


Carb-I-Tex heat shield, however, limits the TZM strength member and the Mo-Re liner temperatures such that little probability of approaching melting conditions occurs. Another requirement for additional research pertains to the response

P

of graphite materials to thermal stresses. Data plotted in Figs. IV-22

and IV-23 indicate temperature gradients of more than 5,000 and 4,000°F for the Carb-I-Tex and POCO heat shields, respectively, and the stress calculations are in the marginal range for each material. While this condition occurs for the conservative assumption of

J

a 3

stable, corner-on entry, the re-entry speed and angle combination to produce these temperature differences was not maximal.

accurate, quantitative evaluation of the superorbital re-entry condition depends upon: (1)

i

J

Therefore

the probability of achieving superorbital re-entry conditions,

(2)

the probability of stable or near-stable re-entry motion, and

(3)

the response of the heat shield to stresses originating from radial temperature differences

-l

J

ranging from 4,000 to more than 5,000°F. A final consideration is the high surface temperature of the heat shield.

The requirement exists to verify calculated recession rates for POCO 0

and Carb-I-Tex graphites at surface temperatures approaching 8,000 F.

1 INSD-2650- 2 9 IV-75


c.

Summary

The flight abort results reviewed above may now be summarized in terms of probabilities of occurrence and system response leading primarily

to

one of two situations: (1)

heat shield failure leading to release of fuel at high altitudes, and

(2)

loss of strength member prior to impact.

The results are presented in Table IV-18 as the estimated negative logarithms (base 10) of probabilities.

Therefore a large, positive

number indicates an estimated small probability. The quantity N represents the negative logarithm of the unknown 1

probability of achieving the combination of re-entry speed and flight path angle designated as the super orbital re-entry range. N

2

represents

the event of exceeding the melt point of the Mo-Re liner. It is believed that the numbers N and N are very large compared 1 2 with unity (small probabilities), based on the previous discussion of flight mechanics.

This contention must be verified by a detailed

study of abort modes for the launch vehicle and their consequences. The composite failure probahilities according to a given mode for a specified system is lO'(Q

+-

R, where Q and R are the negative logarithms

of occurrence and response probabilities.

INSD-2650-29

IV-76

1


TABLE I V - 1 8 IN-FLIGHT ABORT PROBABILITIES Negative Logarithm (base 10) of P r o b a b i l i t y

F a i l u r e Mode

H

Environment

F l i g h t Condition

Occurrence

POCO -

Carb-I-Tex

(2)

5 (3)

Loss of h e a t s h i e l d

Mechanical

Orbital

2 (1)

Loss of h e a t s h i e l d

Thermomechanica 1

Superorbital

N1

0.3

0.3

Loss of s t r e n g t h member

Thermal

Superorbi t a 1

N

N2

' N 2

I

7: ucn

uo I

1

h,

W

NOTES:

-2

(1) An occurrence p r o b a b i l i t y of 10

i s assigned f o r a b o r t s leading t o s h o r t o r b i t s .

(2) Estimated f a i l u r e p r o b a b i l i t y f o r IRHS heat s h i e l d (Ref. IV-7).

(3) The mechanical f a i l u r e p r o b a b i l i t y f o r t h e Carb-I-Tex h e a t s h i e l d w a s assumed t o be smaller by a decade than t h a t f o r POCO, by v i r t u e of i t s b e t t e r s t r e n g t h properties.


F.

1.

IMPACT AND POST-IMPACT STUDY

Impact Containment Impact s u r v i v a l c h a r a c t e r i s t i c s a r e n o t , a t the present time, p r e -

d i c t a b l e on a s t r i c t l y t h e o r e t i c a l b a s i s .

Therefore, no attempt was

made t o determine impact f a i l u r e p r o b a b i l i t y .

However, h e a t sources

comprising a TZM s t r e n g t h member, l i n e r and f u e l simulators and a POCO g r a p h i t e s h i e l d have been impacted a t speeds of 297 t o 394 f p s without f a i l u r e of t h e s e l e c t e d s t r e n g t h member (Table I V - 1 ) . The predicted terminal speed a t impact i s about 245 f p s so t h a t i f s u r v i v a l of t h e heat source i s demonstrated a t 300 f p s , t h e f a i l u r e p r o b a b i l i t y would be small, indeed. f a i l u r e p r o b a b i l i t y of (a)

lo'*

W e have assumed a s t r e n g t h member

which t a k e s i n t o account i n a general sense:

The l i k e l i h o o d of exceeding 245 f p s on impact owing t o corner r e c e s s i o n i n t h e hex h e a t s h i e l d .

(b)

Impact under c o n d i t i o n s more severe than t e s t e d (sharp o b j e c t s , e.g.)

(c)

V a r i a b i l i t y i n h e r e n t i n m a t e r i a l s and capsule f a b r i c a t i o n process

The l i n e r i s assumed t o f a i l i f t h e s t r e n g t h member s u f f e r s breakup o r piercing. The l i k e l i h o o d of f u e l r e l e a s e following b u r i a l of t h e heat source must a l s o be considered.

To t h i s end w e have assumed t h a t b u r i a l t o

a depth s u f f i c i e n t t o m e l t t h e encapsulating m a t e r i a l s occurs i n one

out of t e n random impacts.

A s w i l l be seen below, w e assume a f u r t h e r

INSD-2650-29 IV-78


0

m i t i g a t i o n of t h e subsequent hazard owing t o the immobilization of t h e fuel a f t e r i t s release. 2.

Post-Impact Containment The successful design of an I I H S implies that t h e most l i k e l y

0

happening following an a b o r t and c a p t u r e of t h e nuclear system i n the E a r t h ' s g r a v i t a t i o n a l f i e l d i s i n t a c t d e p o s i t i o n of t h e h e a t source e i t h e r on land o r i n t h e ocean.

3

The advantageaccruing t o an i n t a c t

impact heat source i s thus i n t r i n s i c a l l y a s s o c i a t e d with the time period during which i t remains i n t a c t a f t e r impacting land. I f we assume conservatively t h a t t h e p r o b a b i l i t y of recovery by

a search team following land impact i s small (0.1 o r l e s s ) , then t h e

J

a

value a s s o c i a t e d with a given containment d u r a t i o n i s t h a t of "accid e n t a l l y " f i n d i n g and recovering t h e h e a t source i f t h e r e a r e people i n t h e r e g i o n of d e p o s i t i o n , o r diminished hazard i f the h e a t source i s not found by v i r t u e of only a few people being i n t h e region.

Evaluation of the s a f e t y advantage derived from use of t h e I I H S a s opposed t o t h e IRHS i s performed h e r e i n by computing a p r o b a b i l i s t i c hazard measure.

The model used takes i n t o account random discovery

of the heat source (accidental recovery).

The advantage f a c t o r i s

evaluated a s a f u n c t i o n of containment t i m e ; i . e . ,

t h e advantage f a c t o r

i s evaluated under t h e c o n d i t i o n t h a t t h e heat source remains i n t a c t

f o r a time T . It i s found t h a t , while t h e recovery p r o b a b i l i t y is increased only

by a f a c t o r of abouttwo f o r 100-day containment as compared

-

INSD- 2 650 2 9 IV-79


t o instantaneous r e l e a s e , t h e hazard measure i s improved by n e a r l y

2 o r d e r s of magnitude. a.

Model d e s c r i p t i o n

Assume d e p o s i t i o n of t h e h e a t source occurs i n a region having population d e n s i t y P ( p e r s o n s 1 f t 2 ). I t is f u r t h e r assumed t h a t recovery i s accomplished i f any person approaches t h e h e a t source w i t h i n a d i s t a n c e a/2.

I f t h e average speed of persons located

w i t h i n t h e d e p o s i t i o n region i s

v ( f e e t l d a y ) , then t h e r a t e of

sweeping out area by each person i s a v ( f t 2 /day). Now the a r e a "reserved" f o r each person i n t h e region i s 1 / P . Therefore, the incremental p r o b a b i l i t y of f i n d i n g the heat source i n t i m e increment d t i s av d t = Pav d t .

11P

The d i f f e r e n t i a l equation s a t i s -

f i e d by t h e recovery p r o b a b i l i t y , p, i s t h e r e f o r e yielding p = 1

-

e

-Pav t

* dt

= (1

-

p) P a v

Thus, t h e non-recovery p r o b a b i l i t y f o r a r e g i o n having d e n s i t y p a n d elapsed time T i s t h e simple exponential:

Pm = e

-

Pav T

Now i f t h e population d e n s i t y spectrum i s defined by g ( P ) such that 0

l:

( P ) d P = 1.0, then t h e t o t a l p r o b a b i l i t y of not f i n d i n g

t h e source i n t i m e T is

INSD-2650-29 IV-80


This probability is not the sole measure of safety in that a nonrecovered source which ultimately releases fuel constitutes no hazard if there are no people in the region. Previous analyses have shown that a reasonable risk measure is proportional to the population

I

a

density; therefore, an appropriate risk index can be calculated in the form

In previous studies (Ref. IV-8) we have derived an exponential representation for the probability density function of worldwide population density based on Edringtonâ&#x20AC;&#x2122;s data (Ref. IV--9).

It has the

form: 3

i = l where the parameters A

i

and a i are given in Table IV-19.

Substituting the exponential representation of g (P) into the expressions for P (T) and R (T) we find NR 3

INSD-2650-29 IV-81


TABLE IV-19

POPULATION DENSITY FUNCTION PARAMETERS

-i

A,

Ut2)

ai

-

lo7

1

7 1 . 7 3 x 10

2.9

2

1.98

5 7 . 0 8 x 10

3

105 4 1 . 2 2 x 10

4 8 . 7 8 x 10

c


~~

U ia

.

.....

.

. .

.

....

.

.-

=I 3

and

R(T)

Ai 2 CY

i

i - 1 b.

Results

NR and R f o r 1, 10, 100,

Approximate r e s u l t s were c a l c u l a t e d f o r P and 1000 days.

These a r e given i n t h e TableIV-20.

The average

speed v was taken a s about 1 milelday and t h e combination a v was taken 4 2 as 5 x 10 f t /day.

This, i f one can be assured of containment f o r 10 days following impact, t h e estimated advantage i s a reduction of a f a c t o r of 10 i n t h e r i s k magnitude.

For each decade increase i n containment time,

t h e r e a f t e r , t h e hazard decreases by about one decade.

J

f o r long containment times, t h e r i s k decreases l i k e T

Eventually, -2

.

The r e s u l t s a r e s e n s i t i v e t o t h e q u a n t i t i e s a and v only i n t h e combination avT.

Thus, i f v were taken t o be 1/10 of i t s present

value while a and T were f i x e d , t h e r e s u l t would be t h e same a s evaluated a t T ' = 0.1T: R (T, 0.1 a , v) = R (0.1 T , a , v)

While t h e values assumed f o r a and v a r e only e s t i m a t e s , i t seems unreasonable t h a t t h e product av would d i f f e r by a s much a s a f a c t o r of 10 from i t s assumed value.

J

INSD-2650-29 IV-83


TABLE I V - 2 0

EFFECT OF CONTAINMENT TIME ON FUEL RELEASE PROBABILITY AND RISK

Containment Time, T (days)

Total Non-Recovery Probability PNR (TI

Risk Index R (T)

Normalized Risk Index R (T) /R (0)

1

0.93

1.0 x

0.50

10

0.75

1.9

0.095

100

0.53

2.08 x

1.04 x

1000

0.22

2.79

1.39

INSD -2650-29

IV-84

lo-*


.. .

1

We conclude from this approximate analysis that design of an IIHS heat source for 100-days' containment affords a risk reduction of

li U

about 2 orders of magnitude.

If the source is likely to be recovered

shortly after deposition owing either to the efforts of

a

search team

or from sighting during re-entry, the relative advantage as a function of time remains because of the non-zero probability of non-recovery. However, if early recovery is very probable, non-recovery obviously poses a smaller risk and other hazards become more dominant in in-

3

fluencing design.

a

I

r"

R,i

:

INSD- 2 65 0-2 9 IV-85


G.

DESIGN CONSIDERATIONS

The r e s u l t s of the preceding s e c t i o n s of t h i s chapter may now be summarized.

The r e s u l t i n g p r o b a b i l i t i e s of malfunctions and system

response a r e considered t h e b e s t a v a i l a b l e f o r d e c i s i o n making with r e s p e c t t o the design of t h e heat source.

Where occurrence probabi-

l i t i e s have been roughly estimated based upon p a s t experience, e.g., i n the case of o r b i t a l a b o r t frequencies o r impact containment of the capsule and l i n e r , i t should be recognized t h a t s u b s t a n t i a l e f f o r t i s required t o provide s i g n i f i c a n t l y improved estimates. The r e s u l t s , shown i n T a b l e IV-21,

a r e grouped according t o

consequences of malfunctions and system f a i l u r e - - h i g h a l t i t u d e r e l e a s e and ground r e l e a s e a t t i m e s varying from t h e i n s t a n t of impact t o 100 days a f t e r impact% Also included a r e hazard i n d i c e s i n d i c a t i n g the r e l a t i v e s e v e r i t y of t h e types of events considering t h e q u a n t i t y and nature of released f u e l and t h e e x t e n t of population exposures

I

(Ref* I V - 7 ) . The hazard i n d i c e s were derived f o r t h e microsphere f u e l form. Release a t high a l t i t u d e was estimated t o cause degradation i n s i z e of t h e o r i g i n a l p a r t i c l e s , increasing i n h a l a t i o n exposures. value i s taken a s 1.0.

Its relative

The remaining i n d i c e s p e r t a i n t o ground r e l e a s e

and vary depending upon l o c a t i o n of t h e r e l e a s e (random o r i n t h e launch pad environs),

Ground r e l e a s e s have smaller i n d i c e s by about

2 o r 3 decades than r e l e a s e a t high a l t i t u d e ,

*loo- day

containment was c o n s e r v a t i v e l y assumed based on Isotopes' and vendor t e s t data a c t u a l containment t i m e would be in excess of this value.

-

INSD-2650-29 IV-86

I

I 1 I


TABLE IV-21

MALFUNCTION (NET RELEASE) PROBABILITIES

F a i l u r e Mode

POCO

Consequence

Loss of Heat Shield

High A l t i t u d e Release

+

Shrapnel P e n e t r a t i o n

Ground Release

2

Post Re-entry Impact Failure

Ground Release

Capsule M e l t (burial)

Ground Release

3

Post-Impact Oxidation

Ground Release

2.7

NOTE:

- (N1+0.3)

10

-5 (1,2) 3 x 10

10-3

Hazard Index

Carb -I-Tex

+

- (N1+0.3)

10

1.0

2

3.6

3

5.2

3

5.2

2.7

5.2

for i n - f l i g h t a b o r t s and 0.3 for land impact. (1) Includes p r o b a b i l i t i e s of The same i s t r u e for a l l e n t r i e s below and t o t h e r i g h t . (2) Ignores s u p e r o r b i t a l melting of s t r e n g t h member.

elements i n t h i s row.

Same i s t r u e for o t h e r


The f i r s t row i n Table

IV-21 i n d i c a t e s t h e estimated p r o b a b i l i t y

and consequences of l o s s of the heat s h i e l d during r e - e n t r y .

The f i r s t

term i n the r e l e a s e p r o b a b i l i t i e s d e r i v e s from mechanical f a i l u r e of t h e g r a p h i t e owing t o a c c e l e r a t i o n s p o s s i b l e p r i o r t o peak heating f o r o r b i t a l aborts.

These include t h e f l i g h t malfunction p r o b a b i l i t y

(lo-*) and the system response p r o b a b i l i t i e s and Carb-I-Tex)

taken from Table I V - 1 8 .

or

f o r POCO

Composite i n d i c e s ( r e f .

Table IV-22), formed by multiplying release p r o b a b i l i t i e s f o r each system and t h e l i s t e d hazard index, a r e on t h e order of

and

loe7

depending upon t h e h e a t s h i e l d m a t e r i a l , provided s u p e r o r b i t a l e n t r y

i s q u i t e improbable ( N l > 7 ) .

I n any case, t h i s f a i l u r e mode y i e l d s

one of t h e l a r g e s t composite hazard i n d i c e s f o r heat sources w i t h POCO heat s h i e l d s . Shrapnel p e n e t r a t i o n p r o b a b i l i t i e s were discussed previously and

was derived.

a r e l e a s e p r o b a b i l i t y on t h e order of 2 x

The

hazard index i n t h i s case p e r t a i n s t o f u e l r e l e a s e i n t h e launch pad environs and i s somewhat smaller than f o r ground r e l e a s e i n uncontrolled areas.

This f a i l u r e mode has a n estimated composite index

which i s smaller than t h e major c o n t r i b u t o r s .

I d e n t i c a l r e s u l t s are

estimated f o r a l l systems by v i r t u e of n e a r l y i d e n t i c a l p e n e t r a t i o n c h a r a c t e r i s t i c s f o r each capsule. Post r e - e n t r y impact f a i l u r e owes t o f a i l u r e of t h e s t r e n g t h member and l i n e r elements of t h e capsule. t o have a s i n g l e f a i l u r e p r o b a b i l i t y

(lo-â&#x20AC;&#x2122;),

INSD-2650-29

IV-88

The capsule w a s assumed r e g a r d l e s s of h e a t

I


-. .

.

.

.

. .

.

. .

.

..

- ..

.

.. .

-

.

TABLE IV-22 COMPOSITE INDICES

POCO

F a i l u r e Mode

a s 9 1

Loss o f Heat Shield

+

Shrapnel P e n e t r a t i o n

7 x

- (N1+0

10

Carb'-I -Tex 3)

+

- (N1+0.3)

10

7 x

Impact F a i l u r e (post r e -entry)

1.5

1.5

Capsule Melt (bur i a 1)

1.5

1.5

Post Impact Oxidation

1.4

1.4

T o t a l (excluding supero r b i t a l thermal s t r e s s f a i l u r e of heat s h i e l d )

-6 1.5 x 10

-7 6 . 1 x 10

INSD-2650-29 IV-89


shield differences.

The composite hazard magnitude of about 10

-7

appears t o be r e a l i s t i c and r e p r e s e n t s t h e lower l i m i t of design capability.

(It would be d i f f i c u l t t o demonstrate an impact f a i l u r e -2

p r o b a b i l i t y much smaller than 10

). A t t h i s l e v e l , assuming a l l

o t h e r f a i l u r e modes gave smaller i n d i c e s , t h e i n t a c t impact design would be s a f e r by about two decades than t h e IRHS (used f o r Nimbus 111). The b u r i a l p r o b a b i l i t y , a s i n d i c a t e d above, includes, i n a d d i t i o n t o abort

and Land impact (0.3), a p r o b a b i l i t y of 0.1 f o r b u r i a l

a t s u f f i c i e n t depth t o m e l t t h e capsule.

The hazard index i s a l s o

mitigated by a f a c t o r of 10 owing t o expected immobilization of t h e f u e l p a r t i c l e s by r e s o l i d i f i e d s o i l and capsule materials.

The compo-

s i t e index i s about t h e same as t h a t f o r impact f a i l u r e . I n e v a l u a t i n g r e l a t i v e hazards from post-impact oxidation of the capsule, t h e r e s u l t s of Section I V - F - 2 were used.

Thus t h e hazard

i n d i c e s f o r ground release of f u e l a f t e r 100 days w a s assumed t o be

-3 smaller than f o r instantaneous r e l e a s e (5.2 x 10 ) by a f a c t o r of 100. The 100-day containment t i m e w a s conservatively assumed based on

I s o t o p e s ' and vendor t e s t d a t a which shows t h e probable containment

t i m e t o be i n excess of this value.

A s containment periods i n excess

of 100 days do not s i g n i f i c a n t l y e f f e c t t h e r i s k index ( s e e Table IV-20), 100 days was conservatively assumed for t h i s preliminary evaluation.

The composite i n d i c e s i n Table IV-22

may be used t o d e r i v e con-

c l u s i o n s from p r e s e n t estimates and t o i n d i c a t e emphasis required i n f u t u r e work.

The conclusions a r e as follows:

INSD- 2 65 0- 2 9 IV-90


r?

.'113

a

1.

The minimum achievable l e v e l of r i s k expressed i n t h e r e l a -7 t i v e u n i t s presented above i s about 10 This level r e -

.

D

s u l t s from f a i l u r e of t h e capsule on impact w i t h an e s t i -

-5

a

mated p r o b a b i l i t y of 3 x 10 (lo-* f o r o r b i t a l a b o r t , 0 . 3 -z f o r land impact, and 10 f o r capsule f a i l u r e ) . 2.

Present estimates i n d i c a t e t h a t b e s t source candidates have comparable r i s k l e v e l s w i t h i n about a f a c t o r of two.

3.

a

Key assumptions made i n d e r i v i n g these r e s u l t s include: a.

For t h e heat s h i e l d s , t h e occurrence p r o b a b i l i t y of s u p e r o r b i t a l conditions and s t a b l e e n t r y and thermal

-7; mechanical

s t r e s s f a i l u r e i s smaller than 10

f a i l u r e of heat s h i e l d during o r b i t a l r e - e n t r y i s

(POCO) o r b.

1

(Carb-I-Tex).

Impact f a i l u r e of s t r e n g t h member and l i n e r m a t e r i a l s has a p r o b a b i l i t y of

c.

lo-'.

A containment period i n excess of 100 days i s expected f o l l o w i n g p o s t - r e - e n t r y impact.

Future work should address i t s e l f t o t h e v a l i d i t y of assumptions a s w e l l a s t h e refinement of a l l s i g n i f i c a n t p r o b a b i l i s t i c models.

4.

With t h e above assumptions, performance approaching t h e

-7 estimated lower l i m i t of about 10 would require demons t r a t i o n of:

INSD-2650-29 IV-91


a.

Greater heat shield integrity during re-entry.

b.

Decreased probability of melting -or burying on impact; somewhat reduced shrapnel penetration probability.

c. 5.

Post-impact capsule integrity.

Finally, and significantly, a risk level decreased by 1 to 2 orders of magnitude, compared with the SNAP 19 IRHS

launched on Nimbus 111, appears achievable. In these considerations, we have ignored compatibility requirements between fuel and liner materials. At the present time it is considered that the molybdenum alloys are sufficiently compatible with Pu02 so that containment is not jeopardized based on tests conducted at Los Alamos Scientific Laboratory, Battelle Memorial Institute,and Mound Laboratory. The results of this study suggest that either candidate may be utilized and provide a significant safety advantage over the flight-qualified SNAP 19 IRHS. We have further assumed in this analysis that the physical characteristics of the fuel are those estimated for Pu02 microspheres. If it is shown that present estimates for microspheres are unduly pessimistic, or if a safer form is utilized, the estimated hazard index will be further reduced. The formulation of safety design criteria for an IIHS was the

indicated goal of the safety evaluation. Achievement of this goal requires:


1.

Finalization of probability models particularly in the areas of blast response, shrapnel environment and penetration,

impact media -and their consequences, capsule burial, super-

U

u

orbital re-entry, and re-entry rotational states. 2.

Definition of the relationship between release probabilities and the materials and dimensions of candidate configurations.

3.

Specification of a criterion for the composite index described above (or as an improvement factor relative to

P

t h e IRHS).

With these requirements satisfied, the establishment of design criteria and specifications follows in a straightforward manner.

a 3

INSD - 2650-29 IV-93


v.

r!

1 J

a

RTG/VLC INTEGRATION CONSIDERATIONS

This chapter presents the spacecraft integration considerations associated with incorporating multiple RTG's as the Viking Lander Capsule (VLC) prime power source. The Martin Denver proposal study, as well as studies presently being performed at that facility, were used to establish baseline VLC configurations. Figure V - 1 shows the Viking Lander Model built by Martin Denver based on the Viking proposal configuration which employed two R E units.

A. MECHANICAL

3 1

1.

Installation

A Viking RTG uni&,may be mounted in any orientation on the VLC by attachment to the VLC structure at either one or both of the RTG end closure flanges. A t preseht the mounting bolt circle pattern in each end flange (and associated end cover) co-isistsof six equally spaced holes on a 7.5-inch diameter circle (see Fig. V-2).

I 3

Attachment to the

structure is with 0.25-inch diameter bolts. For a two RTG unit stacked installation (similar to the existing SNAP 19 Nimbus 111 RTG subsystem configuration), the RTG units are bolted end-to-end at the six attachment holes in the mating flanges.

1

d

J

Again, mounting in any orientation on the VLC may be accomplished at either one or both of the RTG stack outer end cover flanges or at the mating flange interface. The mounting provisions and end cover design shown in Fig. V-2 should be considered as typical and can be altered to the VLC de-

INSD-2650-29 v-1


/-

FIG. V - 1 .

V I K I N G LANDER MODEL

1 meteorology instruments 2 radioisotope thermoelectric generator 3 30-mm steerable S-band high-gain antenna 4 propellant tank 5 reaction control jet s 6 telescoping strut 7 crushable honeycomb landing pad 8 terminal descent engine 9 subsurface probe 10 soil sampler

I N S D - 2 65 0- 2 9

v-2

J-


a 3

J

1 1

J

I

P

I

v- 3

INSD- 26 50 - 29

tz 3

.f

I

<\I

31


signer's specific requirements. This may include a variation in mounting hole geometry as well as an attendant change in the end cover design.

Such changes may be desired to permit mating with VLC thermal

control devices designed to extract heat from the RTG during "cold" periods such as during the Martian night. A specific contact area at the RTG/thermal control device interface may be required to achieve the desired heat flow into the spacecraft. The present Martin Denver design employs the RTG mounted directly to the VLC structure with a thermal control device which maintains contact with the RTG in the area of the lower cover during periods when heat flow to the VLC is desired. When heat is not required the device retracts away from the RTG surface, thus opening the heat transfer path from the RTG. The single RTG unit envelope is 10.75 inches high by across the fin tips.

24.4 inches

For a two unit, stacked configuration the height

is approximately doubled, 21.50 inches, and the dimension across the fin tips increases approximately 2 inches. An increase in fin width over the single unit is required to compensate for (1) loss of heat dump capability from the end cover of one RTG and (2) loss of view factor to the environment due to the presence of both RTG's. 2.

Dynamic Considerations The proposed Viking baseline design is similar to that of SNAP 19

generators S/N and 13,

27 and

S/N

28 which were dynamically tested on June 12

1969. These generators were hard-mounted to the vibration

table and dynamically tested in the heated, operational condition to the levels shown in Table 111-6 of Section 111-B. These dynamic test levels were determined for the RTG from Viking launch environment

INSD-2650-29

v-4


7

d

a a

J

a

specifications. Results of these tests showed no change in power output in either generator. Although preliminary in nature, the results of these tests, in conjunction with the considerable SNAP

Nimbus test experience on similar type units, demonstrate the basic

RTG configuration is capable of satisfying the Viking dynamic requirements.

3 3

a

a a

a

19

INSD- 2650- 29

v-5


B. THERMCIl; The RTG's w i l l be subjected to a wide spectrum of thermal environments during the course of the Viking mission.

Prelaunch

environments include those associated with normal ground handling activities of the RTG's prior to installation on the VLC, and operations subsequent to installation, including VLC sterilization within the sterilization canister. After sterilization, the VLC (with RTG's installed) remains in the sterilization canister during subsequent launch preparation, launch and interplanetary cruise. Just p r i o r to deorbiting for entry into the Martian atmosphere the

sterilization canister is ejected.

Subsequent to Mars landing, RTG

operation will be in the sun-shade cycle characterized by the Martian diurnal cycle.

In this section, the thermal limitations for the RTG are presented and operation in the various environments evaluated in light of these limitations. 1.

RTG Temperature Limitations The RTG temperature limitations may generally be categorized as

either those concerned with the thermoelectric degradation rate or those associated with component failure.

In the RTG, the primary cause of an

increase in the thermoelectric degradation rate is any significant increase in the hot junction temperature above the design value.

The

component failures of concern are the melting of the PbTn solder used for bonding the thermoelectric couples at the cold junction to the copper electrical connector straps, and a relaxation of the O-ring seals which


r 7

'3

could result in increased gas leakage from the generator.

k3 D

2N design are defined as 1 0 6 0 for ~ ~ continuous operation and 1075O~for

u

sently being obtained on the SNAP 19 and 29 programs for generators,

The hot junction t emperature limitations for the baseline TAGS-85/

short-term operation (on order of a month).

Life test data are pre-

0

modules and couples operating in the 1025 to 1075 F temperature range (see S,ection 111-B). Results to date show that long-term operation up

to 1075'F may be allowable; however, for purposes of this study, 1075째F

is assumed to be the upper limit for short-term operation until addition-

a

al data are available.

i

53O0F.

The FbIn solder used to bond at the cold end melts at approximately Since the A T from this solder joint to the radiator fin root

i s approximately

solder melt.

@OF, a fin root temperature of 4 7 0 ' ~ would result in

A fin root temperature of 45OoF has been selected as the

limitation to prevent melting at this joint. Operation of the Viton O-ring seals at a temperature in excess of

3

400째F for extended periods of time could result in excessive set of the

a

V i t o n O-ring seals

J

about 400째F or lower for short-term, high temperature operations

cover gas.

and accelerated leakage of the generator internal

Thus, it is desirable to maintain the seal temperature at

such as sterilization. No extensive data are presently available above 400째F to accurately predict the resulting seal integrity for longterm operation.

However, component testing is scheduled to be initiated

at Isotopes in this calendar year to determine the effect of O-ring

I

i


temperature on the seal tightness. Results of these tests will permit selection of a more accurate temperature limitation than is possible with existing data. Table V-1 summarizes the RTG temperature limitations. Evident is that the O-ring seal is the limiting factor for the fin root temperature. 2.

Viking Study Results During the Viking proposal effort, Martin Denver performed ex-

tensive VLC thermal integration studies for various mission thermal environments.

However, these studies were performed using the existing

Nimbus I11 type generator design parameters 350'F

(625 watts fuel loading,

design fin root temperature) and the two-RTG VLC configuration

shown in Fig. V-1.

Table V-2 presents the results of these evaluations.

The worst case for the RTG fin root temperature, 415'F,

as expected

occurs during sterilization. This temperature is considered marginal for the short term sterilization period, as the 400°F limit is exceeded and must be evaluated further in the aforementioned seal tests to be performed. Since the RTG is on short circuit during this operation, the resulting hot junction temperature, 10lg°F, is well within limits (see Table V-2). The maximum RTG fin root temperature at Mars occurs when the VLC is descending to the Martian surface.

The fin root temperature during

this short time period is 365OF and the corresponding hot junction temperature 1045'F.

This is the worst thermal case for Mars operation

INSD-2650-29

v- 8


since the RTG is on load and at the maximum fin root value. From Table V-2 it is evident that no thermal problems associated with the

RTG are anticipated f o r any operations subsequent to sterilization. Verification of seal integrity at sterilization temperatures requires

u

demonstration testing.

It should be noted that the temperature data given in Table V-2

3

assumes the baseline TAGS-85/2N operational characteristics are similar to those f o r the Nimbus I11 generators for which the analysis

3

was performed.

1

This assumption is permissible for a two-RTG YLC

configuration, but may be in error for the present four-RTG configurati'on due to the higher total thermal inventory present. Martin Denver is presently preparing thermal models for the four-RTG con-

s

figuration, and data will be available shortly to permit evaluation of this VLC concept. TABLE V - 1

RTG ?TEMPERATURF:LIMITATIONS Item

L i m i t ation

Comment

' ' e

f-

Hot junction temperature

Thermoelectric power degradation generally increases at hot junction values significantly in excess of the sgiected design value

PbIn cold end solder

Melts at 53OoF

O-ring seals

Operation at elevated temperature f o r extended periods of time causes relaxation of the O-ring and a potential increase in gas leakage from the RTG

'

1060 (continuous

operation) 107>(short-term operation--up to yone month)

'

L

7

J

Fin root temperature limited to 45OoP+ Fin root temperature limited tc+ 400°F for short-term operation (up to 200 k S ) and 375OF f o r continuous operation*

*The liqita$ion.on fin root tem erature is controlled by the O-ring since its limitation is less than that f o r the solder.

3

INSD-2650-29

v- 9


TABU V-2 PREDICTED RTG FIN ROOT TEMpERAmS FOR VIKING'MISSION ENVIRONMEPJT ( Two-RTG Configuration)

Environment

RTG Fin Root Temp. (OF)*

Load Condition

Hot Junction Operating Temp. (OF)

Sterlization

415

Shorted

1019

Prelaunch and Launch

279

Shorted

863-

Cruise, new-Earth

370

Shorted

954

Cruise, near-Mars

363 365

On Load

1043

On Load

1045

Descent t o Mars Surface

'

Mars Surface Operation Extreme Hot

340

On Load

1020

Extreme Cold

194

On Load

874

*The SNAP 19 Nimbus 111 parameters were assumed i n the Martin Denver proposal s t u d i e s (625 w a t t thermal inventory, 350°F design f i n r o o t temperature, two-RTG VLC configuration).


___--_

C.

---

~

. -.. . ..

. ..

. .........

ELECTRICAL

A s was s t a t e d i n t h e g u i d e l i n e s i n Section I-B, t h e e l e c t r i c a l i n t e r f a c e between t h e RTG and Viking Lander Capsule ( n C ) i s assumed t o be a t t h e RTG e l e c t r i c a l output connector.

This assumption

i s based on t h e recognition t h a t i n p r e s e n t program plans, Martin

Denver w i l l provide the a s s o c i a t e d power system wiring and power cond i t i o n i n g equipment (e.g.,

DC-DC converters and v o l t a g e r e g u l a t o r ) .

The content of t h i s s e c t i o n i s o r i e n t e d p r i m a r i l y toward d e s c r i b i n g the RTG e l e c t r i c a l c h a r a c t e r i s t i c s and t h e i r influence on mission planning.

However, a discussion on r e l a t e d e l e c t r i c a l equipment and instrument a t i o n i s a l s o presented. 1.

Load Condition a.

Normal operation

A s i s t y p i c a l with most power sources, t h e performance c a p a b i l i t y of an RTG i s d i r e c t l y r e l a t e d t o t h e load condition under which it i s

operated.

F i g w e V-3 shows t h e p r e d i c t e d nominal o p e r a t i o n a l charac-

t e r i s t i c s of t h e b a s e l i n e Viking RTG a t various t i m e s during t h e mission.

A f i x e d operating l o a d voltage of 2.9 v o l t s has been s e l e c t e d as t h e design p o i n t s i n c e t h e maximum g e n e r a t o r power a t t h e end-of-mission (one year after launch) i s r e a l i z e d

a t t h i s value.

A s t h e RTG output

i s by i t s e l f unregulated, a v o l t a g e r e g u l a t i o n device i s r e q u i r e d a t

some p o i n t i n t h e c i r c u i t , such as a t t h e RTG output o r a t t h e VLC power

bus, t o maintain t h i s voltage.

pi.

1 ~ ~ ~ 2 629 50v-11


IND- 2650-29 v- 12


a $rpa

iJ

In Fig. V-3 it is seen that the hot junction operating temperature variation with load voltage is virtually linear, increasing with increasing load voltage.

Thus, as the load voltage is decreased toward short circuit

from the 2.9 volt design point the hot junction temperature drops.

Con-

versely, as the load voltage is increased above the design point, the

3

a J 13 J

hot junction increases. Since the fraction of heat converted to electrical energy (RTG conversion efficiency) is relatively low (approximately

6%) the effect of load voltage variation on the cold junction and fin root temperatures is small.

Both temperatures change less than l5'F

from the

nominal design values for the extreme load conditions of open circuit and short circuit. At these loads, the RTG power output is essentially zero and the heat to be radiated to space increases approximately 6% above that at nominal operating conditions. b.

Short circuit operation

Operation of the RTG on short circuit at any time desired during the mission is permissible.

3

Unlike batteries, operation with the output shorted

does not result in a discharged condition and will not significantly effect subsequent performance.

As discussed above, operation at s h o r t circuit

results in a lower hot junction temperature and thus is desirable, though not necessary, when electrical power is not required.

For example, pre-

vious SJ!l.AP 19 units have been operated on short circuit during prelaunch storage periods.

Short circuiting the RTG output reduces the hot junction

temperature approximately 100°F below the nominal load value for the same radiator fin root temperature. The current Martin Denver mission operations call for the R E ' S to


be maintained on short circuit during storage, sterilization and interplanetary cruise. As indicated in Table V-2, sterilization is tile only operation necessitating short circuit operation -to assure subsequent 0

satisfactory operation (temperature would exceed 1100 F if maintained In Mars orbit, prior to Mars entry, the RTG's w i l l be switched

on load).

onlload for VLC activation and checkout, and w i l l remain on l o a d throughout the remainder of the mission. c.

Open circuit operation

&tended

operation of the RTG at the open circuit condition is not

allowable due to the high hot junction operating temperatures incurred.

A temperature increase above the nominal value on the.order of 100°F results at steady state open circuit conditions. However, the RTG output may be open circuited momentarily (on the order of 30 seconds) to obtain instantaneous voltage readings for purposes of RTG performance evaluation

load switching.

The rate of hot junction temperature

increase after open circuiting is closely related to the thermal capacity of the heat source,and this characteristic thermal inertia permits short-term open circuit operation. 2.

Environmental Effects a.

Fin root temperature variation

Operation of the RTG in the Martian environment results in a varying R"G fin root temperature which influences the RTG power output. The maximum fin root temperature range during operation on the Martian swface predicted by W t i n Denver proposal studies was presented in

-3

v- 14

E


Table V-1.

However, as the Martin Denver proposal studies are based

on test results and thermal models employing two RTG's of the SNAP 19 Nimbus I11 design

(625 watt thermal inventory, 35OoF nominal fin root

temperature), the results are not directly applicable to the proposed TAGS-85/2N Viking RTG design (675 watts thermal inventory, 33OoF de-

I

a

sign fin root temperature) and f o u r - R E VLC configuration. Thus, o n l y a qualitative evaluation is possible at present. Figure V-4 shows the effect of fin root temperature variation on power output for the Martin Denver predicted operating temperature

-l

range. The results shown are for the VLC configuration presented in

I

the Martin Denver proposal studies (two Nimbus I11 type RTG's).

On

the Martian surface, a maximum power decrease of approximately 2 watts was predicted from nominal operation to either the coldest or hottest Martian environment. Although no substantiating data are presently available, it is anticipated that for the proposed TAGS-85/2N RTG, the power fluctuation curve will be similar to that shown in the figure, and the maximum power decrease on the order of 2 watts. To quantitatively determine the influence of fin root temperature variation, Martin Denver is presently conducting integration studies on the current VLC concept employing f o u r baseline TAGS-85/2N RTG's.

The

results of these studies, in conjunction with scheduled thermal tests in the near future on this type RTG by Isotopes in which power output as a function of fin root temperature w i l l be determined, will permit a comprehensive evaluation.

INsD-2650-29

v-15


v- 16


b.

Housing thermal gradient

The Martin Denver thermal control concept for the VLC requires extracting heat from the R E at the base mounted to the VLC structure. Thus, during the W t i a n night (and other possible mission environments where heat is required) a thermal gradient is imposed on the RTG due to the axial heat flow to the mounting base.

Martin Denver pro-

posal studies have shown the maximum housing temperature gradient and heat conducted to the VLC to be on the order of 50°F and loo watts respectively. This magnitude gradient is expected to have no significant effect on RTG performance, but may conservatively be assumed to be 1 watt for preliminary studies. Scheduled thermal tests to be performed by Isotopes w i l l result in more accurate design values for power fluctuation. These tests will involve determining the effect on powex- output of extracting up to 200 thermal watts at the RTG mounting base. e.

Emissive coating deterioration

The operating environment on the Martian surface is expected to be hostile toward VLC component coatings and finishes.

The radiator sw-

face of the RTG is especially susceptible since it by necessity is exposed directly to the environment.

Thus, the RTG radiator coatings

could potentially be subjected to severe sandblasting by sandstorms.

As the effect of coating deterioration on RTG power output depends on the aforementioned planned thermal tests that w i l l determine the effect of fin root temperature variation on power output, only a qualitative evaluation is possible in this study.

INSD- 2650 - 29

v-17


Figure V-5 presents t h e r e s u l t s of s t u d i e s performed f o r t h e The VLC configuration presented

Martin Denver Viking proposal e f f o r t .

i n t h e proposal employed two RTG's of t h e SNAP The e m i s s i v i t y

(f t) and

s o l a r a b s o r p t i v i t y (&

19 Nimbus I11 design. S

) values were d e t e r -

mined from preliminary s p e c t r a l r e f l e c t a n c e t e s t s performed a t Isotopes on both coated and bare, sandblasted t e s t samples f a b r i c a t e d from t h e present S M P 19 r a d i a t o r and coating materials, MgTh and zirconium oxide/sodium s i l i c a t e r e s p e c t i v e l y .

c t = 0.83,

cys =

0.20.

As-sprayed coating p r o p e r t i e s were

Value for both b a r e and b a r e sandblasted Mg-

Th samples wereCt = 0.3 t o

0.4 and

CY, = 0.5 t o 0.5 (values f o r both

b a r e and sandblasted samples were v i r t u a l l y i d e n t i d a l ) . i n Fig,

V-5, loss of 50% of t h e coating

(et =

t h e f i n r o o t temperature approximately 659F.

As is shown

0.6, ass 0.4) increases

Thus, t h e 340°F m a x i m u m

Martian s u r f a c e operating temperature shown i n Fig. V-4 would increase

to 405OF, with an a s s o c i a t e d power decrease of approximately 1 w a t t . T h i s decrease i s l e s s than t h e 2 w a t t value previously determined f o r

operation from extreme h o t t o extreme cold. The results above may b e q u a l i t a t i v e l y applied t o t h e TAGS-85/2N

RTG.

As t h e nominal d e s i g n f i n r o o t temperature i s lower, 33OoF as com-

pared t o 35OoF, t h e f i n r o o t temperature f o r a comparable coating cond i t i o n w i l l be lower, although t h e power decrease w i l l be approximately t h a t shown f o r t h e Nimbus I11 configuration.

However, as t h e 2 watt

power decrease due t o f l u c t u a t i o n between t h e a n t i c i p a t e d extreme h o t and cold environments i s g r e a t e r than t h e 1 watt decrease due t o coating

INSD-2650-29

v- 18


a 3

3 3

a a

J

2650- 29

. v-19

INS&


deterioration, only the former value need be considered for determining minimum RTG power requirements. RTG radiator emissive coating evaluations are scheduled to be performed by Isotopes in the near future.

These evaluations will in-

clude tests to determine the effect of sandblasting as well as humidity and other gas environments on the stability and thermal properties of the present as well as other candidate coatings. d.

Results

The Martian environmental effects on the RTG w i l l result in a maximum power decrease of approximately

3 watts from the nominal value.

Two watts are due to fin root temperature variation, (including effects

due to both extreme hot and cold environments and deterioration of the coating), and one watt due to the thermal gradient imposed on the RTG during the VLC thermal control function.

The Martin Denver determined

RTG power requirements for the Viking mission stated elsewhere in this report consider the above effects.

Performance testing on TAGS-85/2N

generators and coating evaluation studies to be performed at Isotopes in this calendar year w i l l provide substantiating data for the performance extrapolations made from the Nimbus 111 RTG test data assumed in the above discussions.

3.

Related Equipment Practical limitations in the mechanical and electrical design of

the RTG dictate that it operate at an inherently low output voltage, and into a load which is sufficient to prevent RTG damage due to thbrmal

stress. These two R'I'G constraints require the employment of power

INSD-2650-29

v- 20


7

-3d

conditioning equipment (PCE) t o convert t h e low voltage RTG output t o a higher and more f u n c t i o n a l l e v e l , and t o maintain a constant l o a d

iJ

on t h e RTG.

The power conditioning equipment c o n s i s t s g e n e r a l l y of a

DC-DC converter ( t o convert t h e v o l t a g e ) and a shunt r e g u l a t o r ( t o

maintain t h e l o a d ) . I n t h i s s e c t i o n a discussion on PCE requirements and t y p i c a l des i g n approaches a r e presented.

S p e c i f i c s of t h e PCE design and i n t e -

g r a t i o n w i l l be performed by Martin Denver. a.

DC-DC converter

The DC-DC converter changes t h e low voltage DC output of t h e generator t o an a l t e r n a t i n g voltage, amplifies it t o t h e l e v e l required f o r t h e system load, and then r e c t i f i e s and f i l t e r s it t o provide t h e r e q u i r e d DC output,

I n general, t h e s t a t i c converter accomplishes t h i s

by u t i l i z i n g t h r e e s u b c i r c u i t s : and a r e c t i f i e r - f i l t e r .

a power stage, an o s c i l l a t o r - d r i v e r ,

The o s c i l l a t o r - d r i v e r i s u t i l i z e d t o c o n t r o l

s o l i d - s t a t e switching devices contained i n t h e power stage, such t h a t t h e y a l t e r n a t e l y d e l i v e r t h e generator voltage t o a step-up transformer

u

a l s o contained i n t h e power stage.

The stepped-up output voltage of

t h e transformer i s then r e c t i f i e d , f & l t e r e d and d e l i v e r e d t o t h e load. In order t o obtain a high conversion e f f i c i e n c y , t h e power s t a g e must operate i n t h e switching mode.

I n t h i s type of operation, t h e

semiconductor device i s switched r e p e t i t i v e l y from t h e s a t u r a t e d ( h i g h current, low v o l t a g e ) condition t o t h e cut-off ( h i g h voltage, low c u r r e n t ) condition.

Switching i s accomplished v e r y r a p i d l y through t h e i n t e r -

'

INSD-2650-29 v- 21


medi te, high-dissipation region, and the

irage power dis iP ,teais

very small compared to the total power switched. The transistor is perfectly suited for htilization in low voltage DC-DC converters in that it approximates the operation of an ideal

switch.

This is a result of the low collector-emitter saturation

resistance and the relatively high blocking capability (collectoremitter voltage) of the transistor. transistors: germanium and silicon.

There a r e two basic types of power The germanium unit has a collector-

emitter saturation resistance lower than that of the silicon, and for this reason it yields the most efficient converter.

The silicon unit,

however, has some advantages over the germanium such as higher temperature ratings, faster switching times, and high collector-emittervoltage capabilities. For the Viking mission silicon transistors will be required due to the high sterilization temperature (e l25OC).

Selection

of high current silicon transistors with a low collector-to-emitter

saturation voltage for tkie converter has resulted in a predicted convsrte r efficiency of approximately 85%for the Viking application with an

RTG input of 2.9 volts. Figure

V-6shows

a typical converter circuit of the type previously

utilized for the Nimbus applicatian which was designed and developed by Isotopes as a part of the SNAP19 RTG Power Supply System. Tne DC-DC converter consists of, in part, a current feedback driven power oscilla, and Q3, feedback transtor represented by switching transistors Q2

former T2, power transformer T1, saturable reactor L3, and the base peaking inductors L1 and L2. The power oscillator chops the RTG low DC voltage into a square wave and tr

'

1NS~-2650-29 v-22

the higher required bus


67,

*

.

9a3

01,-

I

5x3


When Q2 i s reversed biased

v o l t a g e a t t h e secondary of transformer T l .

( o f f ) , Q3 i s forward biased (on) t o s a t u r a t i o n and v i c e versa.

They

t h e r e f o r e a l t e r n a t e l y apply voltages of opposite p o l a r i t y t o the p r i mary windings ( t e r m i n a l s 1, 2, and 3) of t h e power transformer T1. Since t h e switching times a r e of t h e order of a few microseconds, a square wave voltage i s induced m t h e secondary ( t e r m i n a l s 11) windings of T1 a t a stepped-up l e v e l ,

9,

10, and

The t r a n s i s t o r s a r e driven

a l t e r n a t e l y i n t o s a t u r a t i o n by t h e c u r r e n t induced i n t h e base d r i v e windings ( t e r m i n a l s 1, 2, and 3 ) of "2 by t h e feedback winding (terminals

6

and

7) of T2.

The frequency a t which t h e t r a n s i s t o r s are

switched i s c o n t r o l l e d by the saturable reactor, L3.

Efficient switch-

over of t h e t r a n s i s t o r s i s caused by negative voltage feedback from winding

7-8 of T1 which i s a p p l i e d t o winding 4-5 of T2 a f t e r L3

saturates.

The time required f o r L3 t o s a t u r a t e determines t h e switch-

i n g frequency of the converter (for t h e SNAP

19, Nimbus

converter design,

L3 s a t u r a t e s i n 1.7 milliseconds thus giving a switching frequency of approximately 600 h t z ) .

c

Neither T1 o r T2 a r e driven i n t o s a t u r a t i o n ,

which enables them t o operate e f f i c i e n t l y .

Both transformers and t h e

s a t u r a b l e r e a c t o r contain a n i c k e l - i r o n a l l o y t o r o i d a l core which provides high i n i t i a l permeability and low l o s s e s .

The base d r i v e c u r r e n t

i s p r o p o r t i o n a l t o t h e t r a n s i s t o r c o l l e c t o r c u r r e n t by t h e r a t i o of t h e feedback winding t o t h e base d r i v e winding.

The switching t r a n s i s t o r s

a r e t h e r e f o r e always operating a t t h e i r optilpum power loss condition r e g a r d l e s s of load c u r r e n t .

I N S D - S ~ ~29 O-

V-24

I

.


â&#x20AC;&#x2DC;1

L)

u

To insure reliable starting of the converter circuit for any environmental condition, a blocking oscillator consiting of transformer T 3 , transistor Q,l

diodes C R l through CR4, resistors R 1 and R2, and

capacitor C1 is employed.

The blocking oscillator applies a pulse

through C E and R 3 to the base emitter junction of the switching transistor Q2.

ia 1

A bias voltage from a winding (terminals 4, 5, and 6)

of T1 and diodes CR3 and CR4 is applied to the blocking oscillator once the power oscillator operation is initiated. The bias voltage shuts off the blocking oscillator and renders it inoperable as long as the power oscillator is operating.

The blocking oscillator is capable of starting

the power oscillator with a generator voltage as low as 0.8 lpIlC.

1

The base peaking inductors, L1 and L2, are incorporated to decrease the turn-on and turn-off times of the power switching transistors. This reduces their switching loss and thus increases the converter efficiency, The square wave output voltage of transformer T1 is full wave rectified by the diodes CR5 and CR6.

J

(In the Nimbus converter design these

are actually the base-collector junctions of germanium transistors which exhibit a forward voltage drop lower than the conventional silicon rectifier, and thus enhanced efficiency.)

a a

The rectified voltage is filtered

by capacitors C2, C3, and C4 and inductor

a. The capacitors are series

connected in order to divide the voltage imposed on them.

The series

connection enhances the reliability of the circuit since the capacitor failure mode is a short circuit.

The resistors,

R4, R5, and R6 help

maintain equal voltage sharing by the three capacitors.

INSD-2650-29 v-25


b.

Constant voltage r e g u l a t o r

For t h e Viking a p p l i c a t i o n t h e power output from t h e RTG will be regulated a t t h e RTG output t o provide a 28 VDC power bus.

The shunt

r e g u l a t o r schematic shown i n Fig. V-7 i s a t y p i c a l c i r c u i t which might The shunt r e g u l a t o r maintains t h e output voltage of t h e

be employed.

RTG a t t h e d e s i r e d value throughout t h e mission by varying t h e e f f e c t i v e

r e s i s t a n c e of a c u r r e n t c o n t r o l l e d shunting element placed across t h e RTG output.

The shunting element d i s s i p a t e s any power ( i n excess of

t h a t demanded by t h e system l o a d ) necessary t o maintain t h e RTG (and bus) voltage a t t h e f i x e d design values. The shunt r e g u l a t o r shown a c t s as a feedback c o n t r o l system.

d i f f e r e n t i a l amplifier (Ql and

Qz) i s

A

used as a comparator which samples

\

a p o r t i o n of t h e RTG voltage (through r e s i s t o r s Rl and F2) and compares

it t o t h e r e f e r e n c e provided by a zener diode 2 2 .

If t h e output voltage

attempts t o go above t h e design voltage, a d i f f e r e n c e voltage V D i s developed by t h e d i f f e r e n t i a l a m p l i f i e r , which i s converted t o a proportional c u r r e n t I ~ b tyh e t r a n s i s t o r &3.

(Variable r e s i s t o r R1, i s s e t such that

ID i s zero when t h e RTG output i s a t i t s design v a l u e . ) by t r a n s i s t o r s s i s t o r QS.

Q4

ID i s amplified

and Q5 and supplied t o t h e base of t h e shunting t r a n -

The base c u r r e n t i s s u f f i c i e n t t o enable QS t o shunt the

necessary c u r r e n t t o maintain t h e f i x e d RTG output.

Through t h e DC-DC

converter, t h e bus boltage i s thus maintained a t 28 VDC. The shunt r e g u l a t o r concept discussed above should be considered as t y p i c a l s i n c e t h e r e g u l a t o r could a l s o be placed across t h e DC-DC conv e r t e r output.

However, t h e present Martin Denver design approach i s t o

I

INSD-2650-29

v-26

II


D

a 3

a 9

--

---

I FIG. V-7.

SHUlYT REGULATOR CONCEPTUAL SCHEMATIC

1 1~m-2650-29

v- 27


r r e g u l a t e the RTG output p r i o r t o the converter t o decrease t h e c u r r e n t load on t h e converter, thereby enhancing converter r e l i a b i l i t y .

4.

Instrumentation I n order t o determine t h e performance of various subsystems during

t h e mission, t h e VLC w i l l employ a telemetry system which transmits various data, i n d i g i t a l form, t o a ground s t a t i o n .

The telemetry system

w i l l contain an a n a l o g - t o - d i g i t a l converter i n order t o t r a n s b t e t h e data i n t o a t r a n s m i t t i b l e form.

If monitoring of t h e RTG and power

conditioner equipment performance i s desired, a telemetry s i g n a l cond i t i o n e r u n i t (TSCU) w i l l be necessary t o condition t h e measurements of t h e various RTG power supply pwameters i n t o analog signals which a r e

compatible with t h e VLC a n a l o g - t o - d i g i t a l converter. The TSCU would r e c e i v e i t s power from t h e VLC 28 VDC bus and would monitor RTG parameters such as temperature and v o l t a g e a n d power conditioning equipment parameters such as temperature, voltages, and currents.

It would then condition t h e s e measurements t o analog outputs

i n t h e form of voltages within a desired range and impedance l e v e l es d i c t a t e d by t h e a n a l o g - t o - d i g i t a l converter c o n s t r a i n t s .

Each of t h e

analog outputs d e l i v e r e d by t h e TSCU wou3.d have an a s s o c i a t e d c a l i b r a t i o n curve or conversion f a c t o r from which t h e measured parameter represented b y t h e output could be derived.

Figure V-8 i s a t y p i c a l block diagram of t h e functions which would be performed by a TSCU f o r t h e Viking system.

In order t o accomplish

t h e s e functions, t h e TSCU would contain v m i o u s s u b c i r c u i t s as described below.

'

INSD-2650-29 v- 28


a.

Temperature measurement

There are three basic devices which can be utilized for measuring temperature and sending changes; the thermistor, the resistor, and the thermocouple.

For measuring low temperatures ( ck .e., temperatures below

70ooF) the thermistor should be used, in that it provides a very large resistance change for a relatively small change in temperature. The TSCU signal can therefore be provided directly, with a small power drain,

by aonnecting the thermistor in a voltage divider circuit such as the one shown in Fig. V-9. (a). The thermistor cannot be reliably used for measuring temperatures in excess of 70O0F, which precludes their use for measurement of high RTG temperatures such as the thermoelectric hot junction. tures may be sensed by a platinum resistance sensor.

High tempera-

This type of sensor

is accurate, repeatable, and reliable if mounted properly and electric a l l y isolated from its environment. The change in the resistance of a

platinum resistor due to a temperature change is very small in comparison to a thermistor and, therefore, additional circuitry may be required i6: order to obtain a desired TSCU output voltage while maintaining power drain at a minimum.

The resistance change, for example, could be con-

verted to a proportional voltage change by a lower power bridge circuit which would then be amplified by an operational amplifier if a high TSCU voltage output is desired.

Such a circuit is sketched in Fig. V-9 (b).

The bias voltages for the previously mentioned vbltage dividers, amplifiers, etc. will be supplied by the rectified outputs of a low

v- 30

'


19

li

B J J

Er s c u=

FIG. V-g(a).

I J

LOW TEMPERATURE MEASURING AND CONDITIONING CIRCUIT

. . /'

FIG. V-g(b)

. HIGH TEMPERATURF: MEASURING AND

J

J rNS~-2650-29

v- 31

CONDITIONING CIRCUIT


power DC-AC i n v e r t e r ( r e f e r t o Fig.

V-8).

A l t e r n a t e l y t h e thermocouple may be used as a temperature measurement device.

However, it r e q u i r e s a more complex conversion system i n

t h a t a reference temperature or compensating c i r c u i t must be provided.

b.

Voltage measurements

The RTG output v o l t a g e and power conditioning equipment voltages can

he measured and conditioned by a simple voltage d i v i d e r c i r c u i t i f i s o l a t i o n and/or p o l a r i t y r e v e r s a l i s n o t required with r e s p e c t t o t h e YLC I f i s o l a t i o n i s required, t h e s e voltages

a n a l o g - t o - d i g i t a l converter.

may be measured and conditioned by t h e DC-DC converter c i r c u i t sketched i n Fig. V-10.

"he base d r i v e for t r a n s i s t o r s Q and l Q,2 i s maintained a t

a constant frequency and voltage l e v e l by a square wave output from t h e

low power DC-AC i n v e r t e r .

and , $2l The square wave i s coupled t o Q

through transformer E, and is s u f f i c i e n t t o a l t e r n a t e l y d r i v e Q and l Q2 , i n t o s a t u r a t i o n f o r every h a l f - c y c l e of t h e square wave.

R e s i s t o r Rl

The TSCU output v o l t a g e is pro

provides base d r i v e c u r r e n t l i m i t i n g .

F o r t i o n a l t o t h e measured voltages by t h e t u r n s r a t i o of T1 (minus t h e various s e r i e s voltage d r o p s ) .

Diodes CRl and CE?, inductor L1 and

c a p a c i t o r C 1 form a full-wave r e c t i f i e r and f i l t e r c i r c u i t with t h e center-tapped transformer T1. c.

Current measurement

The i n p u t and output c u r r e n t s of t h e conditioning equipment may be measured and conditioned by t h e t r a n s d u c t o r c i r c u i t shown i n Fig. V-11. This c i r c u i t enables isolated c u r r e n t monitoring with v e r y s m a l l power

drain.

Transformers T l and T2 have t o r r o i d a l cores, through which t h e

c u r r e n t c a r r y i n g conductor will pass. .

..

7%-

.b"

-

The AC windings

INSD-2650-29 V-32

'

Nl. and N2 a r e


Q r!

b3

n Gi

45

a a

Erscu

J

FIG. V-10. PCE &, RTG VOLTAGE MEASURImG AND CONDITIONING CIRCUIT


E r*cU

FIG. V-11. PCX INPUT AND OUTPUT CUBRIWJ? MEASURING AND CONDITIONING TRANSDUCTOR CIRCUIT

v- 34


connected i n s e r i e s i n such a way that when t h e AC current,

5

(supplied

by t h e low power DC-AC i n v e r t e r ) assists t h e magnetomotive f o r c e of %he measured current, 11 i n one core, it opposes t h a t magnetomotive f o r c e i n t h e other core.

The AC current, I 2 , i s full-wave r e c t i f i e d and f i l t e r e d

t o produce a TSCU output voltage which w i l l vary l i n e a r l y with t h e DC current being measured. Low power DC-AC i n v e r t e r

d.

The DC bias voltages f o r t h e amplifiers, bridges, and voltage dividers, and t h e AC d r i v i n g voltage f o r t h e voltage and current measuring c i r c u i t s

3 1 J

i s supplied by a low power DC-AC i n v e r t e r such as t h e one shown i n Fig. V-12.

The DC-AC i n v e r t e r operates from t h e 28 VDC bus and d e l i v e r s a

square wave voltage t o each of t h e secondary windfngs (S1 through S7). Windings S1 and S2, f o r example would provide t h e AC voltage f o r t h e curr e n t measuring transductor c i r c u i t s .

Windings S3,

S4,

and S5 would pro-

vide t h e base d r i v e power f o r t h e c i r c u i t s measuring t h e RTG output v o l t ages.

The output of windings

~6 and S7 would be full-wave r e c t i f i e d

and

f i l t e r e d t o provide t h e various necessary b i a s voltages. A conceptual schematic of a 1'SCU operating from t h e

28 VDC bus and

supplying t y p i c a l measurement and conditioning c i r c u i t s i s shown i n Fig. V-13.

1

INSD-2650-29

v-35


.

I-

-.- - - - - - -o 20vdc REGULATED BUS

-1

A

s3

c

.

431

-

-.

4

SI

-.(

4

I

FIG. V-12. LCN POWER DC-AC INVERTER CIRCUIT


r---*

FIG. V-13.

-- - -

--

CONCEPTUAL TSCU SCHEMATIC, SHOWING TYPICAL MEASCONDITIONING C I R C U I T S

. INSD-2650- 29

v-37

AND


D.

NUCL;EAR

RADIATION

This section evaluates unshielded radiation fluxes and biological dose rates at points in the vicinity of a spacecraft employing one or more single generators or double generator (stacked) systems. Flux and dose attenuation data f o r typical shielding materials are included. The flux attenuation data have been used to develop near optimum (with 'respect to weight) shield configurations for attenuation of total flux to some prespecified limit. Total shield weight for gamma flux attenuation is given. For situations where personnel access into the vicinity of the

RTG's is required, access times and temporary shielding requirements may be estimated fromthe data herein. 1.

Unshielded Fluxes Penetrating nuclear radiation from plutonium dioxide fueled RTG's

consists primarily of neutrons m d gammas. (a

, n) reactions in oxygen and

Neutrons are produced from

fuel impurities and from spontaneous

and induced fission. The dominant characteristics of the energy spectrum indicate that the primary source is the 0l8 (

Q!,

Ne21 reaction.

Measured emission rates for smalil samples of Pu02 produced with natural oxygen show a maximum of 2.5 x 104 neutrons/sec-gram of PU-238. Combined with a conservative estinrate of neutron multiplication within a distributed source (keff = O.25),

3.33 x

the corresponding source strength is

10 4 neutrons/ sec-gram of pug238

or a total source of

neutrons/sec for a capsule with a loading of presents the neutronssource spectrum.

ms~-2650-29 v- 38

3.96 x

10 7

675 watts thermal. Table

V-3


Gammas are emitted from the fission process, fission product decay, decay of fuel and impurity isotopes and from decay of compound nuclei formed in the ( u , n) process.

Results presented herein are based on

the twenty energy group spectrum of Table V-3.

This spectrum was

derived from the data of Stoddard and Albenesius.* TABLF--V- 3 GAMMA SOURCE SPECTRUM .

I

--

Energy Range (MeV) 0.0

Gammas/sec-gram of Pu238 1

- 0.044

2.36 5.65 2.56 1.30 2.50 6.04 5.38 3.15 1.86 3.81 1.21 3.19 1.12

0.044 0.2 0.2 - 0.3 0.4 0.3 0.4 0.5 0.5 0.6 0.7 0.6 0.7 0.8 0.8 0.9 0.9 - 1.0 1.0 1.2 4.2 - 1.4 1.4 - 1.6

-

1.6

1.8 2.0

1.8 --- 2.0 3.0

5.0

- 4.0 - 5.0 - 6.0

6.0

7.0

3.0 4.0

x lo8 x 107

104 104 103 x 102 x-102 x 105 x 105 x 102 x 102 x 103 103 x

8.50 x 102

7.84 x 2.29 x 2.30. x 6.97 x 2.33 x 5.85 x

102 103

lo2 101 101 100

*D. H. Stoddard, E.L. Albenesius, "Radiation Properties of for Isotopic Power Generators If, DP-984, July 1965.

INSD-2650-29

v- 39

I

PU-238Produced


TABU V-3 (Continued)

NEUTRON SOURCE SPECTRUM

Point Index, i

A

(e,n) N e u t r o n s

B ( n / s e c - g m of Pu-238) F i s s i o n N e u t r o n s

E n e r g y L e t h a r g y( ') (MeV) U

_c

5. 50 5.49 2.97 9.76 2.05

1

14.09

0

2

10.97

0.25

3

8.55

0.50

4

6.66

0.75

0 0 0 0

5

5.18

1.00

0

6

4.46

1. 15

0.93 x 10

7 8

4.04 3.14

1.25 1.50

2.85

9

2.45

1.75

4 2.10 x 10

10. 11

1.91

2.00

0.80

1.49

2.25

lo4 4 0 . 5 6 x 10

12

1.16

2.50

2.43

13

0.900

2.75

14

0.702

3.00

0.71

15

0.546

3.25

0.39

16

0.331

3.75

0.18

17

0.201

4.25

0.80 x

18

0.122

4.75

0.0449

5.75

19

I

lo3 lo3 lo3 lo2 lo2

6.94 x

14.09 MeV.

.-

lo4 4

lo2

2 3 . 4 4 x 10 1 8 . 1 0 x 10

0

..

9.76 x l o 2 3 2 . 0 5 x 10 3 4 . 0 5 x 10 1.56 lo4 4 3 . 3 7 x 10

0.60 x 10 4 0.62 x 10 4 0.57 x 10 4.99 lo3 3 4.09 x 10 3.23 lo3 3 2.47 x 10 3 1 . 3 5 x 10

e

=:

x lo2 x 1.03

0.52 x 10

lo3 3 1 . 3 2 x 10

tn (Eo/E) where Eo

5 . 5 0 x 10' 1 5 . 4 9 x 10 2 2 . 9 7 x 10

0.37

lo4

0.31 x

x 10' 1 x 10 x lo2

3.12 x 103

3 4 1 . 1 9 x 10

.

Composite Spectrum A +B

2.70 x lo4 4 1 . 4 2 x 10 4 1 . 1 3 x 10 3 7 . 4 2 x 10 3 5 . 4 1 x 10 '3.94 lo3 2.86 x 103 1 . 5 3 x 103 7.74 x

lo2

2 3 . 7 5 x 10 2 8 . 1 0 x 10

'


B

13

'

Neutron and gamma fluxes were evaluated in the vicinity of two

R E I s coupled as shown in Fig. V-14 and for a single generator as shown in Fig. V-15. Fluxes were determined by first evaluating the neutron and gamma dose rates at points along the lines shown in these figures. Dose rates were evaluated by use of a point kernel technique

II

employing cylindrical geometry (computer program SPEND).

A brief

description of this technique and associa-Leduncertainties are given later in the section.

Transport calculations for a spherical source

of puo2 in air provide the appropriate dose rate-to-particle f l u x

conversion factors. These calculations ai-e the results of program

DTF (Sn discrete ordinates technique) for a spherical source of plutonium dioxide (radius

B

=

4.4 cm) surrounded by thin regions of

materials which approximate the actual capsule geometry.

Computed

flux-to-dose rate ratios at various distances in air from the spherical source are given in Table V-4.

The data of this table in-

dicates an approximate constancy in the flux-to-doseratios.

J 1 3

variance in the gamma ratio is, in part, d.ueto the angular nature of the f l u x which effects a hardening of the spectrum at large

distances.

Constant flux-to-dose rate ratios of 8.15 n/cm2 sec/mrem/hr

and 1300 photons cm2 sec/mrem/hr were used to scale the computed dose rates and obtain f l u x .

1 1

The

1Ns~-2650-29 V- 4 1


1 ~ ~ ~ 2 629 50V- 42 ..

. .

-


~~

L7 A

Generator

Axis 1 ,

1

I

!

l-

FIG. v-15.

SINGLE GENERATOR CONFIGURATION

SHOWING

FLUX AND D B E RATES WERE COMPUTED.

INSD-2650-29 v- 43

LINES ALONG WHICH

bl


TABU V 4 FLUX AND DOSE RATES FROM A SPHERICAL SOURCE O F PLUTONIUM DIOXIDE

A.

Neutrons

Radial Distance

30

100

Flux

Dose Rate

4790

590

'447

55

18.7

500

B.

(1024 THERMAL WATTS)

8.12 0

0

8.13

2.29

8.17

1000

4.85

0.591

8.21

3000

0.590

0.0711

8.30

Gamma

Radial Distance

30

100

P a r t i c l e Flux

49900

39.5

1260

'

4640

500

194.0

1000

50.1

3000

Dose Rate

5.99

INSD- 2650-29 v-44 .

3.64 .145

1280

1340

0355

1410

.00381

1572

I


Q Neutron and gamma fluxes are given .i.n Tables

V-5 and V-6 a t

v a r i o u s p o i n t s a l o n g l i n e s L1 t h r u L 7 ( F i g s . V-14 and

V-15).

Since

P 9

t h e c o n f i g u r a t i o n s e x h i b i t n e a r r a d i a l symmetry about t h e i r axes and

n

greater t h a n t h o s e g i v e n i n Tables V-5 and

symmetry about t h e mid-plane,

results may be i n t e r p o l a t e d t o o b t a i n

t h e f l u x e s a t any p o i n t e x t e r i o r t o t h e RTG's.

For d i s t a n c e s

V-6 t h e r2 r e l a t i o n s h i p

may be used; i . e . , 2

r,

Q P

where r

1

i s a p o i n t a t which t h e f l u x i s ' t a b u l a t e d (rl greater than 3.0 f t ) .

a a 3

I

P INSD-2650- 29

v-45


TABLE V-15

NEUTRON AND GAMMA FLUXES I N THE VICINITY OF

A 2-RTO !CANDM CONFIGURATION (See Fig. V-14 f o r Location of Lines ) Distance from Geometric Center of Configuration (ft)

P a r t i c l e Flux ( /cm2-~eC)

Line -

Garmna -

Neutron

1.92 x 104

1.82

105

1.46

105

L1

1.38 x i o4 4 1.14 x 10

1.25

105

0.75

L1

6.63

x 103

7.90

x 104

1.0

L1

4.18

103

5.16

x 104

2.0

L1

1.18

103

3.0

L1

5.0

0.27

L1

0.42

L1

0.50

1.51 x

i o4

5.35 x 102

6.89

103

L1

1.95 x 102

2.50 x 103

8.0

L1

7.60 x 10'

9.72

10.0

L1

4.85 x lG

6.16 x 102

15.0

L1

2.14 x lo1

2.71 x 102

20.0

u

1.19 x 101

1.51 x io2

30.0

L1

5.23 x

0.27

Id-A

3.35

x 10

0.42

L1-A

1.61

x io

0.5

Ll-A

1.19

104

0.75

L1-A

6.19

103

x 102

d

loo

6.88

4

4.17

105

4

2.02

105

x


i?

t3 Ji

TABLE V-5 (continued) Distance from Geometric Center of Configuration

Q

u

J

a J si

Particle

F ~ U ( /cm2-sec)

-

(ft)

Line

Neutron

Gamma

1.0

L1-A

3.86 x 103

4.78 x 104

2.0

L1-A

1.44

1.0

L2

5.29 x 104

3.0

L2

5.30

5.0

L2

1.92 x 102.'

2.41

10.0

L2

4.78 x lo1

5.99 x 1 3

20.0

L2.

1.17 x lo1

1.46 x 102

30.0

L2

5.14 x

loo

6.36 x lo1

1.0

L3

4.57

103

5.58 x 104

3.0

L3

6.07 x 103

5.0

L3

2.17

10.0

L3

5.36 x lo2

20.0

L3

1.12 x

30.0

L3

4.91 x 1 0 '

1.0

L4

5.01

103

5.72 x 104

3.0

L4

4.79 x

lo2

5.11

103

5.0

L4

1.70 x 102

1.81

103

10.0

L4

4.2!0 x lo1

4.43 x lo2

20.0

L4

1.04 x lo1

1.08 x 102

30.0

L4

4.52 x 1 0 '

6.67 x lo3

x lo2

lo1

INSD-2650-29

v-47

103

103

1.31 x 102 5.71 x lo1

4.72 x lo1 . .

1

104


TABU V- 5 (continued) Distance from Geometric Center of Configuration

Neutron

(ft)

Galllma

1.0

5.35

103

5.34

104

3.0

4.35 x lo2

4.04

103

5 .o

1.52 x 102

1.40

103

10.0

3.74 x lo1

3.43 x 102

20.0

9.21 x 1 0 '

8.37 x 1 0 '

30.0

4.02 x 1 0 '

3.64 x lol 5.01 x io4

103

1.0

5.50

3.0

3.68 x 102

2.89

5.0

1.26 x io2

9.57 x 102

10.0

2.26 x lo2

0.76

3.05 x lo1 1.95 x 104

0.83

1.28

1.0

6.30 x lo3

1.38 x 105 6.10 x i o4

8.72 x 102

6.69 x 103

3.0

3.36 x 102

2.41

103

5.0

1.09 x 102

7.37

x 102

8.0

4.05 x lo1

2.63 x 102

2.54 x 10'

1.63

X

15 .O

1.10 x lo1

6.94

x 1 0 '

20.0

6.08 x 1 0 '

3.78

30.0

2.63 x

2.0

10.0

,

.

INSD-2650-29 v-48

104

loo

103

2.24

.

1.63

105

' .

102

lol x 1 0 '

I


TABLE V-6

NEUTRON

D R P

3

AND GAMMA FLUXES

IN THE VICINITY OF A

SINGLE GENERATOR CONFIGURATION (See Fig. V-15 for Location of Lines) Distance from Geometric Center of Configuration (ft)

Particle nux ( /cm2-sec)

Line -

Neutron

GaDltEL

0.27

L1

3.08 x LO4

3.97 x

105

0.42

L1

1.36 x

1.79

105

0.50

L1

9.58 x 1103

1.26 x 105

0.75

L1

4.33

LO^

5.68 x io4

1.0

L1

2.45 x 1.03

3.20

2.0

L1

6.13 x 1 - 0 ~

7.95 x 103

3.0

L1

2.72 x 1.02

3.51 x 103

5.0

L1

97.6

1.25 x 103

8.0

L1

37.9

4.85 x

10.0

L1

24.2

.3.09 x lo2

15.0

L1

10.7

1.35 x lo2

20.0

L1

5.95

75.4

30.0

L1

2.60

32.8

1.0

L2

2.41 x 103

3.07 x

3.0

L2

2.67 x lo2

3.37 x 103

5.0

L2

96.0

1.20 x 103

10.0

L2

23.8

2.99 x lo2

INSD-2650-29

v-49

:LO4

104

lo2

104


TABLE V-6 (Continued) Distance from Geometric Center of Configuration

P a r t i c l e nux ( /cm2-sec) Line -

Neutron

GaSrma

20.0

L2

5.86

72.9

30.0

L2

2.56

31.7

1.0

L3

2.30 x io3

4 2.69 x io

3.0

L3

2.54 x 102

2.99 x

5.0

L3

91.3

1.08 x lo3

10.0

L3

22.7

2.67 x io2

20.0

L3

5.58

65.3

30.0

L3

2.44

28.3

1.0

L4

2.12 x 103

2.22 x 10

3.0

L4

2.35 x 102

5 -0

L4

84.5

10.0

L4

21.0

2.50 x io3 8.98 x 102 2 2.22 x 10

20.0

L4

5.17

54.7

1.0

L5

1.98 x i o3

1.89 x i o4

3.0

L5

2.16 x io2

5 00

L5

77.5

2.05 x 103 2

10.0

L5

19.2

1.82 )r 1 0 '

20.0

L5

4.73

44.6

30.0

L5

2.07

19.3

(ft)

INSD- 2650-29

V- 50

103

4

7.35 x

10


TABLE

Distance from Geometric Center of Configuration (ft)

P

J

a J I

V-6 (Continued) Particle lux ( /cm2 -see)

Line _._

Neutron

Gamma

1.0

~6

2.11 x 103

1.89 x 104

3 *O

~6

2.15 x lo2

1.82

103

5.0

~6

76.4

6.38 x

10'

10.0

~6

18.8

1.56 x lo2

20.0

~6

4.63

37.8

30.0

~6

2.02

16.5

0.38

L7

1.91 x 104

2.24 x 105

0.45

L7

1.25

1.38 x 105

0.62

L7

6.04 x :LO3

1.62

L7

7.86 x

2.62

L7.

2.94 x :LO2

2.41 x 103

4.62

L7

93.1

7.36 x lo2

7.62

L7

33.9

2.63 x lo2

9.62

L7

21.2

1.63 x lo2

14.62

L7

9.07

69.2

19.62

L7

4.99

37.8

29.62

L7

2.15

16.1

INSD- 2650- 29

v- 51

104

LO^

6.10

104

6.70 x 103


2.

Shield Attenuation and Weight Optimization In the final analysis, reduction of flux levels, if required, must

consider the total weight associated with placement of the RTG's and shielding within .the configuration. The flux levels could obviously be reduced by the use of booms to increase the source-to-detector separation distance. Another option would be the placement of "shadow" shielding between the RTG sources and the areas where radiation sensitive equipment may be located. "he problem of optimized shielding (minimum total system weight) would be solved by consideration of a l l factors such as boom weight and shield weight (including canning and support structure).

This section discusses shield attenuation for

reduction of the total flux or gamma flux. For a fixed configuration, and given a f l u x design criteria, the data of this section provide the nearly optimum shielding required. Neutron and gamma fLux attenuation data were developed by use of the previously mentioned spherical source of Pu02. The DTF code was used to compute the neutron and gamma fluxes and dose rates at various distances in air beyond the source surrounded with varying thicknesses of typical neukrons and garmna shielding materials which are suitable for space applications.

The gamma and neutron source spectra used in these

calculations is that of Table V-3.

Fluxes and dose rates at varying

distances in air from 3Oto 3000 dm from the center of the source were compared to those computed for the unshielded source to obtain attenuation factors.

v- 52


For distances greater than 3.0 ft from the source, attenuation The attenuation factors

factors were found to be nearly constant. are shown in Figs.

v-16 through V-19. These results may be considered

u

applicable for points of interest greater than 3 feet from the RTG

P

between shielded and unshielded configurations are due to differences

sources.

Differences in attenuation factors as a function of distance

in the angular spectral distribution for shielded and unshielded cases.

(At distances less thaa 3 feet, the attenuation factors would be somewhat less than those shown.)

Shield material densities a r e as given

below:

D

Shield Material

( g/cm2 1

Density

( 1b/in3)

Neutron :

r3

J J J

a li

Polyethylene

0.92

0.0332

Lithium Hydride

0.82

0.0296

11.4

0.412

15.3

0

21.5

0.777

Gamnta:

Lead Tungsten Platinum

(no)

553

For a given 'flux criterion, the attenuation data of Fig. v-16 through V-19 coupled with the unshielded flux at the point of interest (obtained from the data of Tables V-5 and V-6)may be used to determine a near optimum shield configuration.

INSD- 2650- 29

v- 53

Two types of criterion


,

1DLs~-2650-29 V- 5b , ”

I


P P

8

P

la P

INSD-2650 - 29

v-55


1

3

2

I

I

I

1

I

l

I

I

I

I