Aether Group Report June 2009
Aether is an extremely long endurance multi-role unmanned aerial vehicle designed to operate with a 60kg powered payload carried in two pods below the wing. Applications include aerial reconnaissance, communications relay and atmospheric sampling. This report summarizes and explains the design of the aircraft and describes techniques that can be used to recycle the vast majority of the aircraft components.
Aether Engineers Configuration definition and CAD Craig McPherson Adam Omar Aerodynamics Andrew Carmichael Nicholas Thornton Structural analysis Abiel Neo Ronald Uzande Weights and payloads Michael Chow Shu Zhang Aircraft performance Laurence Lai Tom O'Connor Flight control system Manu Goel Peter Shone Aircraft systems Jose Arizaga Andrea Carrara Operations and costs Charlie Burnett Valerio D'Alessandro Power systems Rene Du Cauze De Nazelle Chi Leung Ho Information management and technical coordination Nikhil Chandaria Thomas Robinson
Contents Configuration and CAD .................................................................................................................................. 4 Aerodynamics .............................................................................................................................................. 10 Structures .................................................................................................................................................... 16 Weights & Payloads ..................................................................................................................................... 20 Aircraft Performance ................................................................................................................................... 26 Control Systems ........................................................................................................................................... 31 Aircraft Systems........................................................................................................................................... 36 Power & Propulsion ..................................................................................................................................... 42 Life Cycle Analysis ........................................................................................................................................ 49 Bibliography................................................................................................................................................. 54
Configuration and CAD This summary will discuss the finalised configuration of the aircraft along with the finalised Computer Aided Design models of the project
Configuration The initial configuration considers various broad approaches to year long sustainable flight before the design of any aircraft components can be considered. A variety of concepts were considered at this early stage ranging from basic lighter than air vehicles to complex mechanical aircraft to maintain high efficiency. The ultimate approach taken follows the path of a large solar powered wing to generate power throughout the hours of daylight with the support of batteries for outputting power during the night. The illustration below shows the concept of the chosen design path.
The concept of the aircraft revolves around extreme high efficiency in order to utilise continuous flight with minimal power consumption. Components have therefore been designed with this concept in mind and will be discussed in this section. The aircraft boasts an extremely high aspect ratio of 30 with the main wing spanning 60m with a chord of only 2m. This aims to keep the lift induced drag generated to an absolute minimum, an important focus for an ultra high lift coefficient wing section. Placed at the end of the wing are specially designed wingtips called â€œwinggridsâ€?. These are a new concept designed to reduce wingtip drag (a major component of drag in the aircraft). The aircraft generates its thrust through eight motors powering four 1.5m diameter twin blade propellers. This was designed to provide ideal operating conditions for both the motors powering the aircraft as well as an optimum efficiency for the propellers. Propulsion is the major consumer of power within the aircraft however addition powered is required for avionic systems and the running of the payload. A total calculation of the power required gives a solar panel area of 90m2 and a battery weight of 102kg The aircraft includes a dual fuselage layout with the two fuselages situated at the midpoint of each semi span. The two fuselages reach back approximately 15m from the main wing and both support a vertical Project Aether
and horizontal stabilising tail. This offers the stability required for the aircraft to maintain efficient flight. The twin tail design arose from the need to support the large span of the main wing. With a conventional single tail, the main wing stretches out 30m either side of the fuselage and hence endures large bending moments and torsion loads. The twin tail layout greatly reduces these loads and hence allows a hefty weight reduction from a structural point of view. A major weight of the aircraft lies within the batteries used to store electrical energy for continuous flight throughout the night. With the aircraft designed to fly at +/- 45Â° Latitude throughout every day of the year, a longest night can calculated at 16.5 hours. As mentioned earlier, a total weight of the batteries required can be calculated at 102kg. In order to distribute weight evenly throughout the span wise direction of the aircraft, the batteries are shaped into long thin sections covering a total distance of 16m span wise. This is split into three segments as tabulated below: Segment Battery 1 Battery 2 Battery 3
Position -1m to 1m -22.5m to â€“29.5m 22.5m to 29.5m
Weight 12.5 43.75 43.75
A camera is included in the aircraft placed centrally looking forwards to allow remote control from a ground station. Further avionic systems are repeated in a dual redundant philosophy. These systems are placed outboard of the central battery and are symmetric in layout as to maintain a fair span wise weight distribution. The placement of the avionics system can be seen in the illustration below. Midspan
Upon completion of the major design components, a series of more minor components have been designed that are often over looked in the bigger picture. These include such features as the avionics platform which offers a horizontal support shelf to place the avionics equipment on. With little research required in these components, they have been built around existing modelled parts and hence are heavily designed during the CAD phase of the project. Illustrations of these parts will be explained in later sections
CAD Computed Aided Design is an essential part of modern engineering. The CAD model of the LEEO-V aircraft serves as the primary method of visualisation. The CAD model is also used to provide a platform for component integration, as well as to calculate values required by the other technical teams. These values include calculations of the centre of gravity, moments of inertia, and various masses and volumes.
The CAD model of the aircraft was modelled and assembled in PTCâ€™s ProEngineer Wildfire 3. The majority of components were designed based on data and specifications received from the specific technical teams, however, FATT 01 was responsible for designing certain components not covered by the other technical teams. These include the following components. Tail Boom Joint Wing Boom joint Engine Motor Cowling Horizontal and Vertical Tail Tips Avionics Shelving Images of these components can be seen further below. The final CAD model consists of 935 parts. The large number of parts is partly due to the limitations of the program coupled with the extreme dimensions present on this aircraft. Problems arise with modelling the extremely thin skins and ribs of the wing and tail, which have a thickness of 0.125mm. The 60m span wing skin has been split into sections of 25cm, and similar segmentation has been done elsewhere to avoid software errors. The level of detail of the CAD model extends to the individual components and the internal aerostructure. The aim is to model each component as accurately as possible to the given specifications. Adhesives, bolts, connectors, wires and cables were not modelled due to the time constraints and scope of the project.
CAD Results The following pages contain: 1. A gallery of components designed by FATT 01 2. An orthographic projection of the complete aircraft model 3. A perspective view of the complete aircraft model Colours have been exaggerated for clarity and visibility The external skins are made of a translucent material The solar film on the outboard section of the port wing has been removed to reveal the underlying structure Components designed by FATT 01: Tail-boom joint:
Engine motor cowling (in silver):
Payload Pods (small, medium and large):
Aerodynamics Initial Configuration The first step in the aerodynamic analysis was to look into existing solutions that have similar design aims and performance specifications. This required us to look into solutions such as; solar-powered, longendurance UAVs (the Qinetiq ‘Zephyr’ and the NASA Helios) and high performance gliders (the ‘eta’ glider designed by Flugtechnik & Leichtbau (Flugtechnik & Leightbau, 2003)). By investigating these solutions we could identify the key aerodynamic features; A large aspect ratio, to reduce the induced drag of the aircraft. The maximisation of the aerodynamic performance, which in turn means maximising the ratio of liftto-drag, L/D. The optimisation of the aerofoil design to meet with aerodynamic performance and structural requirements The inclusion of low drag geometries of the other, no lifting components. A awareness of the necessity for the minimum drag of all components.
Aerofoil Selection As with any airplane, the aerofoil selection is one of the key parts of the aerodynamic design of an aircraft. It alone is responsible for providing the lift for the aircraft, and also is an important contributor to the total drag of the aircraft. Our aircraft required a wing section capable of providing a large lift (CL = 0.815) and a very low drag (CD = 0.0101). Due to the low amount of solar energy available at high latitudes in winter, the amount of thrust the engines could provide was relatively low, which in turn meant the total drag of the aircraft had to be low as well. This also meant that the speed at which the plane was able to fly would be much lower than initially estimated. This in turn meant that for a given weight, the lift coefficient of the aerofoil had to be larger than if the plane had been flying at larger speeds. Selecting an aerofoil for our aircraft was a complex process, as it would involve analysing thousands of pre-existing aerofoils to see which one proved to be the most adequate for our needs. Instead, it was decided to make use of one Javafoil’s most powerful features: modifying a pre-existing aerofoil. This could be done in several ways, including modifying the thickness to chord and camber to chord ratios, trailing edge deflection, and most importantly, the pressure distribution along the chord of the section. In essence, Javafoil enabled the aerofoil to be reverse engineered, by selecting the desired characteristics, and then re-shaping the aerofoil to fit the desired performance. This process was done manually and consisted primarily on trial and error, though several underlying ideas were used to obtain the best results possible: A similar approach as that used in the design of supercritical aerofoils was used: instead of having a suction peak, the objective was to create a large suction “plateau”, which would avoid the positive pressure gradient and allow for a larger area of lift producing surface while delaying separation of the flow. This also would also lead to an increase of the area of laminar flow, delaying transition and hence reducing skin friction drag.
The lower surface was designed such that the pressure near it was almost identical to the free stream pressure. This limited the lift produced by the wing, but added to the fact that there was also a near zero pressure gradient along the chord, helped reduce the overall drag significantly by delaying the transition to turbulent flow until the trailing edge. The aerofoil pressure distribution was modified for a given angle of attack; hence performance at other angles of attack was not optimal. This was done due to the large proportion of time the aircraft was to spend at cruise conditions, case for which it was optimized. The reason for not flying with α=0 was that an increased α meant that during winter months at high latitudes, the sunrays would strike the solar panels in a less oblique fashion. The aerofoil selected as a baseline was Selig/Donovan SD7032. The original version was much too thin for the structural requirements of the wing, but roughly provided the qualities our aerofoil needed: Low camber to reduce drag and a high lift coefficient at low angles of attack. The final resulting wing section after it was optimized proved to satisfy both the initial requirements of high lift and very low drag. As will be explained in further detail later on, this aerofoil was used as a constant section throughout the wing. A similar process was used to design the horizontal tailplane. Information on the characteristics of the required section was obtained from FATT 05, who provided a required CL and size of the tailplane. In this case however, the baseline model was a NACA 6-series aerofoil: NACA 63-408. This series was designed specifically for low Reynolds numbers, and hence was ideal for use as a tailplane. Using a cambered aerofoil instead of a simple symmetric aerofoil also allowed us to reduce the drag component of the tail, as it was designed specifically so that the trim angle of the tail was near zero degrees, as opposed to having symmetric aerofoil at large deflection. This provided substantial decrease in pressure drag from the tailplane. More information on the tail design can be found in L.Lai and T. O’Connor.
Drag Estimation After selecting the aerofoil section which suited our needs, producing high enough lift-to-drag ratio (L/D) and the minimum amount of zero-lift drag (CDo), the drag of the other components of the aircraft, and the total drag of the final aircraft configuration had to be estimated. The final drag coefficient of each component was calculated, using various methods including a panel method solver (JavaFoil Applet, 2007), ESDU data sheets (ESDU, 1973) and empirical estimations (Serghides, 2008). The results can be seen below (fig.1):
Component Wing Booms Horizontal Tails Vertical Tails Wing Tips Payload Pods
Zero-Lift Drag Coefficient 0.0055 0.0324 0.0068 0.0060 0.0091 0.0783
Lift-Induced Drag Coefficient 0.0049 0 0.0112 0 0.0122 0
Drag Coefficient 0.0104 0.0342 0.0180 0.0060 0.0213 0.0783 Fig.1 Component Drag Coefficients
Due to the large proportion of lift-induced drag of the main wing, a wing-tip was employed that would reduce this effect. By employing the wing-tips discussed in section 4, the lift-induced drag of the main wing was reduced by 35%. If we take each of these values we can break up the drag of the aircraft into its respective components, as shown below in figure. 2:
Fig.2 Component Break Up
In the above plot the â€˜other componentsâ€™ section accounts for the drag of the hubs, payload pods and the wing tips of the aircraft. The total drag coefficient of the aircraft was estimated using a component build-up method, which uses the contribution of the drag from each of the components to estimate the global drag coefficient. The algorithm used took the form:
By estimating the drag coefficient of each component and then multiplying it by planform area of the component divided by the total planform area of the aircraft, we could correctly weight the drag contribution of each component. Upon calculating this value an error of 5% was added onto the estimate. This was to take account of possible regions where the drag had not been accurately predicted. This includes regions of separated flow, regions of turbulent flow and also to account of the fact that a smooth surface finish had been assumed over each component. The drag coefficient of the aircraft was finalised, using this procedure described above, to be; The drag of the aircraft was then calculated at cruise conditions, with a flight speed of 30.6 m/s , at an altitude of 65000 ft to be 66.40 N. Using this value, and calculating the lift generated by the main wing, we were able to calculate the L/D of the aircraft. This was calculated to be 62.51. This in turn equates to a glide angle of 0.86o, and a sink speed of the aircraft of 0.46m/s.
Wing Tips Selection and Design As the design of the aircraft progressed, it became apparent that the total drag of the aircraft was too large. It shouldn’t have been a major problem, as all that was required was an increase in thrust. This however, would in turn require more solar panels and batteries to power the engines which would increase the overall weight substantially. Instead, several methods were researched aimed at reducing the drag produced by the aircraft. The first method was aimed at achieving an elliptical loading on the wing, as this has been proven to provide the optimal efficiency in terms of drag reduction (Anderson, 2001). The means of achieving this type of distribution are varied, though generally involve adding geometric or aerodynamic twist to the wing or tapering the chord. All these methods aim to achieve a span efficiency (Oswald’s efficiency “e”) of 1, which is taken to be an ideal case. In practise, achieving this lift distribution is very hard, and maximum values of e = 0.9 are typical. In our case however, an efficiency of greater than 1 was required for an appreciable drag reduction, which would in turn lead to a decrease in weight. Hence, non planar features were researched. The optimal design was what is known as Winggrid (La Roche & La Roche, 2004). This method involved placing several smaller winglets in a staggered cascade arrangement at the tips of the main wing.
This arrangement can be considered non-linear, as its performance cannot be accurately described by linear theory of superposition of several configurations. In effect, these winglets serve as a means of reducing the induced drag by breaking up the trailing vortex that would normally form at the main wing tip. This disperses the intensity the main vortex would have, hence reducing the induced drag. This design however, required a rectangular lift loading, which was the main reason for not implementing any sort of twist or taper to the wing.
The benefits of this design are apparent if we compare the Oswald’s efficiency of this design with respect to an elliptical loading and the consequences it has on the drag of the aerofoil. Further technical information on the design of the winggrid can be found in the technical summary in section 6.
Oswaldâ€™s Efficiency Factor Aircraft L/D Total Aircraft Drag
Elliptical Distribution 0.9 53.2 78 N
Winggrid 1.46 62.5 66.4 N Fig.4 Winggrid Comparison table
Wake Analysis After being given the positioning of the booms and the tails (15m from the wing tips and 15 back from the aerodynamic centre of the wing respectively), a downwash analysis had to be performed to calculate the possible downwash from the main wing, not including the wing tips. This was done by performing a Biot-Savart Analysis (Anderson, 2001). The results of this analysis can be seen below (fig.5):
Fig.5 Downwash Analysis
The results show that at the spanwise position of the booms the downwash is minimal, with a velocity of just 0.499 m/s. This results in a velocity seen by the horizontal tail of 30.604 m/s. This interaction between the vortex ejected from the main wing and the tail, can be seen graphically using a Trefftz plane analysis, carried out using Tornado (project redhammer, 2007).
Fig.6 Wake-Tail Interaction
* All values taken for cruise at 65000ft
Summary * Aircraft Speed
Total Lift Total Aircraft Drag CD(a/c) L/D(aircraft) Main Wing Section L/D(section) Horizontal Tail Section
4150.94 N 66.40 N 0.0130 62.51 Selig/Donovan SD7032 (modified) 153.34 NACA 63-408 (modified) Fig.7 Summary of Technical Data
Span (no wingtips) Span (inc. wingtips) Chord length, c Wingtip blade chord, cwt Taper ratio, Λ Wing angle of attack Wingtip section (1st blade) Wingtip section (2nd blade)
60 m 63 m 2m 0.5 m 1 2o NACA 63-108 (modified) NACA 012108 Fig.8 Summary of Wing Dimensions
Fig.9. Main Wing Section- SD7032 (modified)
Thickness-to-chord ratio, (t/c) Maximum thickness location, (x/c)m Camber-to-chord ratio, (f/c) Maximum camber location, (x/c)mc
12.59% 38.90% 3.95% 49.00%
Fig.10. Horizontal Tail Section – NACA 63-408 (modified)
Bending Moments (Nm)
Total Bending Moments 35000 25000 15000 5000 -5000 0
Distance from root (m) Load Moments
The structural design objective was to design an aircraft of minimum weight. This as well as simplicity in analysis is achieved by distributed the loads on the wing such that it is span loaded for level flight.
Torsion The structure would be subjected to torque loads that would result from the chord wise distribution of inertial forces with the lift acting at the aerodynamic centre and the weight at the centre of gravity.
Fig. 1.3.1Torque due to Inertial Loading A wing tubular spar was employed under the assumption that the entire wingâ€™s torsion and bending stiffness is provided by this structure. The design procedure is then dependent largely on the stiffness of the material chosen for the spar.
Gust Loads Analysis was simplified by looking at sharp edge gusts to determine whether or not the structure would able to resists the loads produces under the given design weight.
Sharp edged gust and velocity triangle Change in angle of the wing given by
Additional lift induced on the wing due to the gust is determined by
The structure was found to fail under moderate gusts only being able to withstand only 1.09g manoeuvring loads.
General Rib assembly at wing/boom joint Rib design is carried out using FE Software in order to avoid the complexity and the time constraint of doing this analytically. Different style ribs with varying cut outs are designed to account for the interactions that take place across the span of the wing.
Torque Tube Layup No.of Plies
40 30 20 10 0 0
Distance from root (m)
The design of the torque tube was optimized by tapering the tube so that it is sufficiently stiff in regions where the maximum loads would be encountered. This produced a large weight saving as opposed to having a constant thickness throughout the 60m span wing.
Structural Vibration Vibration analysis was carried out using ht weightless lumped mass concept in order to find the natural resonant frequencies of the wing structure. This is essential before the flutter speeds of the aircraft are found to ensure that these speeds do not coincide with the natural frequencies of the wing.
Weightless beam/Lumped Mass idealisation Equations of motion:
The unit load method is used to find the flexibility matrix where the displacements are caused by bending and torsion so that
f5 0.93 Natural Resonant Frequencies of wing structure Component
Mass (kg) (1 component)
Batteries and Systems
Final Structural Weight breakdown The material properties of the UHM Epoxy/Carbon fibre composite used are given below: 0 Degree Ply Properties E1 (GPa) E2 (GPa) G (GPa) v12 v21 Ply Properties
365 7.14 5 0.23 0.0045
Weights & Payloads Initial Weight estimation The vehicle weight can be divided into 7 main sectors. Systems
Solar Panel, Battery, Engines & Propeller
Main Wing Structure Spar, Ribs, Wing Skin
Tail Plane Structure Horizontal & Vertical Tail, Rod
Payload, Pod Fairing
Undercarriages (zero weight solution)
Avionics System Cabling & Others
Weight budgeting & further Weight Reduction Component Weight
Initial Total Weight
Component Weight Percentage
Initial Target Weight
New Weight Budget
Evaluation of Weight Budgeting Method This method of weight budgeting is ideal in terms of understanding the weight proportion of each components. It gives an idea how far off the percentage difference is compared to similar vehicle such as QinetiQ Zephyr. However, it is also important to look into how weight reduction of each group affects others as well as the realist of the reduction due to technology constrains. This leads to further weight iteration and investigation of possibility of further weight reduction. Increase efficiency in battery and solar panel Battery Energy Density: 260Wh/hr 320Wh/hr Solar Panel Efficiency: 13% 15% No gust design Change in wing skin Material PVC Elastomer Shore 5A (1.1 x 103Kg/m3) PMP which has a density of (840Kg/m3) Weight Contributor Solar Panel
Initial Weight/ Kg 100.50
New Weight/Kg 79.88
Tail Optimisation The distance between the wingâ€™s aerodynamics centre and the tailplane wingâ€™s aerodynamics centre has been sized arbitrary of 15 metres. Investigation was made to reduce the weight of the tail. The tailplane weight is inversely proportional to the increase of boom length due to reduction of tail surface area. Conversely, the boom weight is proportional to the increase of boom length. Hence the combination should produce a minimum mass. The blue line on figure 4shows with boom length of greater than 10m, the reduction of weight is very small but the change of CG is large. But taken into consideration such aft CG, it was decided along with FATT05 team that the boom length of 15m would be the final length.
Weight distribution Initial Idea - Wrap the battery around the pod. HOWEVER: High manufacturing Cost Aerodynamic Inefficiency Improved Idea - distribute the battery throughout in the wing structure Advantages Minimise Bending moment Make full use of internal wing volume Remain slender pod for aerodynamic efficiency. Battery
Major Battery Part
Minor Battery Part
Wing Span = 60m
Centre of Gravity z x
Initial POD Cross-section x
y F IGURE 1 - DEFINITION OF COORDINATES ( ISOMETRIC VIEW LEFT , CROSS SECTIONAL VIEW
The overall centre of gravity (global CG) is calculated by the sum of moment caused by individual mass away from divided by the total weight of the aircraft. (Roskam, 1989) p.117 ;
Where = individual component mass,
= local CG for individual components, W= the total weight
Longitudinal CG Acceptable range where the range will provide a positive stability margin: 1.12 < x < 1.26 Lateral CG The lateral CG is ensured to be equal to zero (YCG = 0), so that it is laterally stable. Local Centre of gravity Local CG was determined either manually calculated or using Pro-Engineer as some objects like ribs are very difficult to calculate the local C.G. Final Value of CG The finalised centre of gravity was determined once the individual weight and the location of local CG were determined. And the table shows the CG of Aether is within the acceptable range. Longitudinal CG, x (m) 1.191
Lateral CG, y (m) 0.000
Vertical CG, z (m) -0.004
Moment of Inertia The moment of inertia and it is a vital property which determines the manoeuvrability of an aircraft in the mechanics of flight. The moment of inertia calculations are dependent on the position of componentâ€™s local centre of gravity and global C.G. The moment of inertia calculations are calculated using the parallel axis theorem. (Roskam, 1989) p.121-122 Local moment of inertia Local mass moment of inertia of individual components is dependent on the shape of the component. The local mass moment of inertia of the component is assumed to be negligible as the order is a lot smaller in comparison and it is extremely difficult to approximate manually, unless the shape is simple. Final Value of Mass Moment of Inertia 4
IXX (m ) 6.648
IYY (m ) 2809.2
IZZ (m ) 2802.5
IXY (m ) 0
Since the aircraft is symmetrical, the inertia products calculations.
IXZ (m ) 54.8
IYZ (m ) 0
are zero which are validated by the
Payload Weight Variation Change of payload weight will vary the performance of the aircraft. For instance, the global CG position will change. Investigation has been made to find out the change of X-CG with change of payload weight. The graph shows by the variation of payload weight have a small change on the position of the longitudinal C.G. However; the C.G values are still within the acceptable range shown by the dotted lines. Hence, using this Project Aether
configuration of the aircraft, the change in payload weight would not destabilise the aircraft. Based on the graph, a linear relationship has been determined, and the empirical relationship is given by: Final (Target) Weight The final target weight has been calculated by the balance of lift and weight during cruise conditions. To calculate the final target weight, the overall lift generated by the aircraft calculated by FATT02 & FATT05 and are summarised in the following table. Lift generating component Wing Wing Grid Tailplane Total
Lift (in N) 4109.0 117.7 144.0 4370.70
Lift (in kg) 418.9 12.0 14.7 445.60
Therefore, the target weight of the vehicle is 445.6kg. In order word, the aircraftâ€™s final weight is 445.60kg. With the spar weight designed to be 71kg, the total weight of the aircraft including the weight of cabling is 416.0kg, which is 29kg less than the target weight, and to achieve the target weight, a method of weight allocation is used. Weight Allocation Weight allocation is a procedure used to either reduce or increase the weight of the component in order to achieve the final target weight. From section 13.1, the aircraft was 29kg under weight. A decision has been made to allocate the 29kg to the spar component. From all of the components, adding structural weight was more beneficial than other components. The benefit of increasing the spar weight is to increase manoeuvrability of the aircraft. With 71kg of spar, the vehicle could only take 1.09G.
Payload Payload Selection Mission Substitution Application
Atmospheric ChemistryAir Quality/Ozone
Surveillance & Reconnaissance
Advanced Very High Resolution Radiometer
PR- 274 Airborne Terminal Eq.
Metrology Research & Earth Observation - MEDUSA (Used in the Mercator-1 UAV) Property
Weight Ground Resolution 2Kg
30cm at 18Km
Power Consumption 50W during the day
Surveillance & Reconnaissance For surveillance and reconnaissance, the basic required equipments are optical camera, infrared camera and Synthetic aperture radar.This equipment consists of 3 units, including 2 directional antennas Units Microwave Modem Assembly (MMA)
Size/m 8.05 x 2.83 x 0.69
Power Consumption +28 VDC, 227 W 208 VAC, 3 phase, 400 Hz, 370 W
Directional Waveguide Antenna (Fore) Directional Parabolic Antenna (Aft)
2.83 x 2.83 x 4.30
Included in MMA
2.59 x 1.58 x 2.74
Included in MMA
Payload Integration The payload is carried by two pods which will be integrated under the wing. The decision to use the pods instead of integrating payloads within the wing was driven by the many benefits, I. Allow different sizes of payload to be used and providing a multi-functional vehicle. E II. Easy accessibility to payloads, easy for maintenance and upgrades. III. Save cabling weight However, although there are a lot of benefits towards using pods to integrate the payload. However, there are disadvantages when using the pod, they include: A. The external pod increases skin friction drag of the vehicle. B. The use of the pod will limit the advantage of distributing the payload along the span to achieve minimal bending moment. It was decided to use the pods as the advantages provide a vast amount of attractive selling points to the customers.
Pod Design The pod design involves connecting the payload directly to the boom. This idea was inspired by researching on the mechanisms used for connecting additional equipments onto fighter jets. The payload would be packed into a frame, and the frame is joined to the boom using integrally machined lugs and is secured by using locking pins and screws. The weight of the frame is included as the payload. Figure 3 & 4 illustrate the design idea.
F IGURE 3 - P OD DESIGN
F IGURE 2 - F RAME J OINT
The aerodynamically shaped fairing provides an external cover for the any arbitrary shaped frame as long as the dimension is smaller the fairing. The benefit of using a frame will provide multi-connection points where the strapping of payload would become very easy. Also, the frame can provide a stable storage space for the payload as the frame is fully locked by the pins. The lugs are designed where one pair would be normal to another pair, so that all degree of freedom will be locked. Furthermore, in order to have a balance aircraft in the lateral direction, the payload must have equal weights on each of the two pods. This will maintain an overall lateral C.G. of zero.
Pod Sizes Like sizes of payload like a LOS Antenna (25x25x25cm) (Global Hawk - Integrated Communications System, 2009) would require a large storage space. Three choices of payload fairing were offered. It offers an extra option for the operator to choose different sizes of payload, which will be an attractive selling point to the seller. Project Aether
Pod Fairing Radius Internal Volume (per pod)
SMALL 13.5 (cm) 54.6 (litres)
MEDIUM 20.0 (cm) 117.3 (litres)
LARGE 27.0 (cm) 208.4 (litres)
*In comparison, a medium sized suitcase has a capacity of 66 litres and large suitcase has 98 litres. However, the increase of size will reduce the value of L/D and consequently reduces the power available to the payload. Pod Fairing Lift to Drag Ratio (L/D) Power constraint
SMALL 62.5 236.8 (W)
MEDIUM 61.1 198.4 (W)
LARGE 59.7 159.0 (W)
Pod Materials Initially, Aluminium SiC was the ideal material, it offers good impact absorption property and supreme low density. However, it absorbs electromagnetic waves, which limits the use of radars or antennas within. Therefore, Polymethylpentene (PMP) was chosen as the material for the pod fairing. PMP has an excellent optical transparency, which allows radar, antenna equipment to be used. And PMP is recyclable too. Pod Fairing Fairing Weight (per pod)
SMALL 430 (grams)
MEDIUM 911 (grams)
LARGE 1278 (gram)
Fairing Integration The integration of the fairing would involve using adhesives, where the upper edges would be glued onto the bottom of the wing. It is a cheap, effortless solution. And also, it is a very light weight solution.
Aircraft Performance Tail Sizing Using the tail volume coefficient method (Raymer, 1965) and typical values for gliders, initial horizontal and vertical tail areas were found. The required inputs were 120m2 wing area, 60m span and 2m chord and the vertical and horizontal tail volume coefficients were 0.02 and 0.5 respectively. The distance between the wing and tail aerodynamic centres was 15m. This was decided taking into account the downwash induced by the wing, the overall weight of the tail/boom configuration and the aircraft centre of gravity position. After completing a Biot-Savart analysis the downwash on the horizontal tail was found and induced a comparatively small force. The variation of weight of the tail-boom configuration with respect to distance from the wing was calculated. At 15m distance the weight was not minimum but chosen as a compromise to achieve the desired centre of gravity, discussed later in longitudinal static stability. Taking into account the twin tail configuration, this total area was halved to find the area and therefore properties of each tail surface. This gave the following results: Table A1.1 - Tail geometry Horizontal Tail 2
Total Area (m ) 2 Individual Area (m ) Aspect Ratio Span (m) Chord (m) Sweep Angle (째) Taper Ratio
8.00 4.00 6 4.90 0.82 0 1
9.60 4.80 1.5 2.68 1.79 0 1
The vertical tail aerofoil section was chosen to be of NACA 0009 to allow sufficient internal space for structure whilst not incurring too much drag. The horizontal tail section was required to trim the aircraft whilst incurring the lowest drag and was therefore designed so that at zero angle of attack it produced this desired lift. The horizontal section was chosen by the aerodynamics team.
Figure A1.1 - Horizontal tail section NACA 63-408
Longitudinal Static Stability Neutral Point The neutral point is the most aft position the centre of gravity of the aircraft could be before it becomes unstable. It is obtained by finding the point where the pitching moment derivative with respect to angle of attack is zero. The pitching moment derivate is affected by wing, horizontal tail, fuselage and engine contributions. The fuselage had a negligible effect due to its relatively small dimensions. The wing and horizontal tail lift curve slopes were provided by aerodynamics as 6.7 per radian. Measured from the leading edge of the wing, the neutral point for this aircraft was calculated to be at 1.35m.
Static Margin A small, positive static margin was desired such that the plane was stable in pitch but not require excessive elevator deflection to maneuver. It was chosen to be 7.0% requiring a centre of gravity position to be 1.20m from the leading edge. However, the final position of the centre of gravity was 1.19m meaning that the static margin became about 8.0% which was within our desired static margin range. Trim For trim conditions the pitching moment coefficient about the centre of gravity C m cg had to be zero for the total cruise lift coefficient. For this to be satisfied the tail had to be set at a certain incidence. To find this incidence a trim plot of total lift coefficient against Cm cg was created. Initially using a symmetric horizontal tail section, the tail angle required to stabilise the plane was found. However the required tailplane angle was very high and meant that there would be excessive drag and that the aircraft would be difficult to control. To solve this, an aerofoil was designed so that it gave the necessary horizontal tail lift coefficient at close to zero degrees.
Figure A2.1 - Longitudinal trim plot It showed that the aircraft angle of attack for trim was about 0.5Â° and required a horizontal tail angle of just 0.1Â°.
Lateral Static Stability When the aircraft is in steady, level flight, no moments are created in the yaw or roll directions due to the symmetry of the plane. However, angle of sideslip derivative could be determined and this analysis was carried out to demonstrate that the aircraft possessed sufficient control power and directional stability. This means that with a perturbation away from steady level flight conditions, the aircraft would return to itâ€™s original state. It involves the closely couple coefficient of yaw and roll derivatives. Determined using two different empirical methods (Raymer, 1965) (Roskam, Airplane Design Part VI: Preliminary Calculation of Aerodynamic, Thrust and Power Characteristics, 1987), the rolling and yawing moment due to sideslip derivatives, Cnb and Clb were found from the wing, fuselage and tail contributions to the moment. Table A3.1 - Lateral-directional stability coefficients
Method 1 Cnb Clb
Typical Stable Values
0.03 – 0.10 -0.02 – 0.08
Comparing the obtained values with typical expectations for stable aircraft, the results indicated that our aircraft was indeed stable. Wing tunnel testing and CFD analysis would be required to validate the results before the aircraft was manufactured.
Control Surface Design Pitch Control The pitch of the aircraft is predominantly controlled by the horizontal tails. Unlike conventional aircraft which employ elevators on the horizontal tail, the Aether aircraft uses an all-moving tailplane. This allows required lift forces to be achieved with much smaller deflections and was easy to achieve since the tail sections were so light. Since the tail was all moving, this meant that the rate of change of coefficient of lift and of pitching moment coefficient were the same as those for the overall section. The following data was supplied by the aerodynamics group. Table A4.1 - Pitching coefficients
Yaw Control The yawing motions of the aircraft would be principally controlled by the vertical tails. Again, these tails are all moving, allowing desired sideforces to be generated with low deflections. As was the case for the horizontal tails, the rate of change of lift and pitching moment coefficient were the same as for the section since the tails were all moving. The following numbers were supplied by the aerodynamics group. Table A4.2 - Yawing coefficients
Roll Control Spoilers were deployed to provide the primary rolling moments. The spoilers were designed to lower the lift coefficient on one half of the wing and hence create a rolling moment about the centreline. The desired roll rate was 0.25 degrees per second and so the spoiler sizing was adjusted so that the spoilers produced the required rolling moments about the centreline to achieve this. The spoilers were deployed at 90° at x/c = 0.8. They ran from 18.75m away from the centreline, to 25.75m from the centreline in a spanwise direction. The height of the fully deployed spoilers was 0.08m which, crucially, was designed to be significantly taller than the height of the boundary layer at this point. There are small gaps in the spoiler surface to allow for Project Aether
the rib flanges in the wing. An analysis was carried out to make sure that the spoilers did not create excessive yawing moments due to the increased drag on one side of the aircraft. The yawing moments were found to be very small (about 1.5 N) and would not strongly affect the performance of the aircraft while the spoilers were deployed. Table A4.3 - Roll properties Roll Rate with Spoilers Deployed (deg/sec) Deflection (Ëš) Height (m) Boundary Layer Height (m) Inboard edge distance from centerline (m) Outboard edge distance from centerline (m) Porosity (%) 2 Spoiler Area Exposed to Flow (m ) 2 Top Down Spoiler Area (m )
0.25 90 0.08 0.0075 18.75 25.75 4 0.5376 0.0134
Aircraft Performance Stall Speed The stall speed including thrust effects of the aircraft was determined for a wide range of altitudes but Table 5.5.1 shows only values at principal altitudes for brevity. The formula used was a rearrangement of the lift coefficient equation with a known maximum lift coefficient used and thrust accounted for. Rate of Climb The maximum attainable rate of climb and maximum climb gradient were found with an equations involving aircraft weight, propeller efficiency, wing area, maximum motor power, aircraft drag and lift coefficients and relative air density. The inputs had to be in imperial units as the result was in units of feet per minute. The change in motor output power with altitude was supplied by the power systems team and the results were tabulated. The service ceiling of a propeller aircraft is defined as the height at which the rate of climb drops to 100 feet per minute and was found to be at approximately 74000ft. Assuming the climb rate was constant between known stations, with the distance between stations predetermined, an approximate value for minimum climb time from sea level to cruise altitude was found of slightly greater than 7 hours. Maximum Velocity Plotting, for a given altitude, the variation of power required by the propeller and maximum power available from the motor with velocity allowed for the determination of maximum velocity at that altitude. The point where more power was required to go faster than the motor could possibly supply defined the maximum velocity for that altitude. Table A5.1 - Performance at altitude
Altitude (m) 0
Stall Speed -1 (ms ) 5.77
Climb Rate (ft Maximum Velocity -1 per minute) (ms ) 217.2
Maximum Climb Gradient (Âş) 10.5 0
Altitude (m) 5000 10000 15000 20000 24000
Stall Speed Climb Rate (ft Maximum Velocity Maximum Climb -1 -1 (ms ) per minute) (ms ) Gradient (ยบ) 7.44 200.6 51.4 7.4 9.94 181.6 50.0 4.9 14.50 146.7 44.7 2.5 21.53 111.0 39.0 1.2 29.70 90.2 31.9 0.4
Figure A5.1 - Aircraft flight envelope Summary For further information about any topic covered in this section, please refer to the appropriate reports regarding Aircraft Performance.
Control Systems The stability matrices for the aircraft are as follows,
Flexibility Assumption The general formulation of the equations of motion for an elastic plane includes the six translational and rotational degrees of freedom as in the case of the rigid plane as well as the many elastic degrees of freedom. These elastic degrees of freedom make it extremely difficult to accurately model the equations of motion. The closest approximation is to model all the degrees of freedom for a elastic plate in the x-y plane (Bisplinghoff, Ashley, & Halfman, 1955, p. 633). This makes it more sensible to stick with a rigid body assumption than a thin elastic plate assumption.
Flight Simulation Flight simulation is often important in the design process to validate the predicted behaviour of the aircraft by comparing the simulation results with the analytical results. The software to be used for the simulation was X-Plane, a widely used piece of software that was easily accessible and is suitable for the task. The equations used in analysing the dynamic behaviour are for a rigid aircraft, and so a reliable comparison can be made. The simulation will not be true to real life however due to the inability of the software to deal with the probable large flexibility effects in the structure, and so using it as a tool to observe the actual behaviour of the final aircraft will not be useful. (In fact a piece of software that performed this degree of analysis was not found.) So instead, the approach was to use it for the marketing of the project, providing a tool for the business group to use as they wish. This does not mean that the model does not need to be accurate, as it would be used as a validation tool for the rigid aircraft equations used in the analysis. As a note, X-Plane v8.6 was used because X-Plane version 7 does not have the capabilities to model solar powered aircraft and requires specification of fuel storage. Modelling the aircraft in X-Plane Plane Maker required many different parameter values regarding the dimensions, weights, control surfaces and thrust specifications. The aerofoil was designed specifically for the aircraft, and so it needed to be created in X-Plane Airfoil Maker by entering the various aerodynamic performance characteristics regarding lift and drag, camber and thickness. The complete set of values used for the simulation can be found in Appendix 2. Project Aether
Autopilot Design Controller The autopilot system to control the aircraft that is selected is an optimal controller which minimizes the deviation of the control surfaces. The optimal controller selected was a Linear-Quadratic Regulator (LQR). The LQR is used in order to minimise a particular performance index for stabilising a particular linear time-invariant state equation (Williams II & Lawrence, 2007). In this case this performance index will include a high weight for minimising the control surface deflections and energy requirements due to an extremely tight power budget suggested by FATT09 (Power Systems and Integration). The exact weight for each parameter has to be tuned, but such an iteration is necessary as this controller is not only efficient but also inherently robust, reasons for which are discussed later on. Once the controller is decided on, it is worth obtaining the approximate gains for the Stability Augmentation System, which will ensure steady flight or a change in a particular parameter to obtain a desired condition. It is important to note however that this is no way is the actual values for the controller, as such values cannot be obtained till a much later stage in the actual design process involving wind tunnel testing (for correct Aerodynamic parameters). This however can be used as a reference to understand the behaviour of the controller selected and its robustness can be quantitatively analysed. It is however assumed that these results are comparable with the actual controller and proving that if a robust controller is possible to design for this case implies that the robust controller is possible to design for the actual case. In order to obtain these values a simple MATLAB1 code is written which simply gives the controller gains by inputting the â€˜lqrâ€™ function after defining the Q and the R matrix. The Q matrix puts constraints on the state vector outputs, while the R matrix puts constraints on input vector. It must be noted that if too many parameters are prioritised, the controller is seen to be aggressive and thus have large gains, therefore it better to have only the important parameters (like control surface deflection) prioritized in order to obtain a moderate controller. The performance index for a LQR Optimal Controller is given by the following integral (Nelson, 1998)
The Optimal controller maximises this integral parameter for given state and input vector constraints. Longitudinal Stability Augmentation System In the case of the Aether the state parameters that were prioritized were the pitch rate (due to structural reasons) and the pitching angle (due its relation with angle of attack and thus affected by stalling capabilities of the aerofoil). The input vector constraints were placed on the maximum elevator deflection. Thus the performance index matrices Q and R chosen were assumed as follows:
MATLAB 7.63, Mathworks Inc.
Using the above matrices A, B, Q, R the controller gains are found out by using a MATLAB code as
F IGURE 4 - P OLE D IAGRAM WITH C ONTROLLER FOR THE P HUGOID AND SPPO
The eigenvalues obtained by this system are graphed in the figure above Lateral Stability Augmentation System In the case of Lateral motion the state parameters that were prioritized were the roll rate (due to structural reasons) and the yaw rate (due to structural reasons as well). The input vector constraints were placed on the maximum spoiler height and on the rudder deflection. Thus the performance index matrices Q and R chosen were assumed using Brysonâ€™s Theorem (Hespanha, 2007) as follows:
Using the above matrices A, B, C, Q, R the controller gains are found out by using a MATLAB code as
F IGURE 5 –P OLE D IAGRAM WITH CONTROLLER FOR THE D UTCH R OLL (L EFT ), THE S PIRAL AND R OLL S UBSIDENCE (R IGHT )
The eigenvalues obtained by this system for the Dutch Roll, Spiral and Roll Subsidence are graphed in the figure above. Robustness An overall robustness analysis still allows for all three conditions to have large variations to occur at the same time and still maintain stability. In order to test this, the density and the velocity were pushed up to 50% variation and the aerodynamic parameters were allowed to be pushed up to 20% variation. The results (Appendix VIII) show that stability is still maintained, and since these values are so large any further analysis would be redundant and unnecessary, as this clearly established the robustness of the controller. The frequency changes are also observed and insured to be away from the natural frequencies of the aircraft. In order to understand these results better it is important to look at the inherent robustness that an LQR controller is associated with. It is shown by using the Kalman’s inequality that an LQR controller is inherently robust (Hespanha, 2007)and is highlighted by the Nyquist plot for a LQR controller.
F IGURE 6 - N YQUIST P LOT FOR A LQR CONTROLLER
Though this cannot be directly applied to reason out the reasons for the robustness of the controller selected for a multi variable system, but can definitely seen as an indication of how LQR controller have a tendency to be robust. Continuous Flight In order to undertake the complete flight envelope, a mission script has to be prepared which informs the controller the desired state of the aircraft. The controller thus adjusts the current state to match the desired state, thereby enabling the aircraft to attain autonomous control for the complete mission period. The aircraft’s desired motion for each day was calculated by FATT09 (Power Systems and Integration) in order to maximize the power obtained and minimize to power expended. This profile is incorporated in the autopilot module, thus there would be no need to transfer data from the ground station for this purpose. The autopilot computer reads the script every 15 minutes to find the desired position of the aircraft and readjusts the current position to match the desired position. Project Aether
The flight profile also takes into account the latitude at which the aircraft is flying. In order to achieve this it uses its inbuilt GPS system which is assisted by the MAVINS to locate the aircraft position, thereby allowing the autopilot module to select the correct mission profile for the corresponding latitude. Manual Control and Override In the unlikely event that the ground station needs to manually control the plane with the Autopilot system turned off, a manual override option will be included in each controlling station. This will allow the controller to fly the plane using the onboard camera which will transmit live data to the ground control station.
Gust Load Alleviation Systems (GLAS) Atmospheric Turbulence at lower altitudes experienced during takeoff and landing effect the normal operation of this aircraft (due to the high aspect ratio) and is a design constraint. The Aircraft structure is affected by these extra loads, thereby resulting in a heavier airframe. The use of a Gust Load Alleviation System (GLAS) reduces these loads by adjusting the wing control surfaces in accordance to the change in the lift due adverse weather conditions. In order to counter this difficulty a forward looking sensor, a Light Detection and Ranging (LIDAR) sensor is installed on the aircraft, which allows for the easy prediction of the vertical velocity 100m in front of the aircraft. The elevator and the engine thrust are correspondingly adjusted in order to compensate the change in lift produced by the incoming gust, effectively using a feed-forward system. This provides for potentially reducing the gusts loads on the aircraft by 90% but in practice due to gust measurement uncertainty and actuator rate limits it is expected to reduce the load by 50% (David Soreide, Aug. 1996). The minimum requirements for the LIDAR system to provide the desired performance were presented to FATT07 (Systems Selection and Packaging), who chose the appropriate LIDAR for this purpose.
Aircraft Systems Launch & Recovery Method The Launch & Recovery Method designed for the Aether is based on six supporting Launch & Recovery Vehicles (LRV) which effectively provide a state-of-the-art, autonomous, reliable and fully-automated solution which is an essential aspect of the business plan generated. They all integrate a computer, a small electric motor, braking discs and a front-steerable axis. Their general arrangement with respect to the Aether (as determined by FATT 03 Structures) is shown below:
Figure 1 â€“ Diagram showing the location of all six LRV: 4 Front LRV (F) and 2 Rear LRV (R). These distances have been calculated by the department of Structures. All six vehicles are aligned in the above configuration by means of a local-GPS method. There are two types of LRVs: Front LRV Four of these are used and they serve as supporting points along the wing. This type of LRV has a Clamped Release & Recover (CRR) capability. The upper part of this vehicle presents a mechanism which is able to clamp itself onto the Aether (by means of a top-hinged surface) for both the launch & recovery stages, thus providing stable ground operations. In order to minimize structural damage, an airbag system is used for providing sufficient grip. For launching, the Aether is placed on top of the LRV, the top-hinged surface is lowered and the airbags are slowly inflated until providing a firm grip. The LRV is then programmed to automatically navigate towards the runway. This is done with the use of a local-GPS system in which a set of beacons are placed throughout the navigation path of the LRV and which can in turn pinpoint its location with great accuracy. All LRVs share the same software and are therefore synchronized. As they accelerate down the runway, the Aether starts generating lift and eventually it produces enough lift to start the climb-stage. Just before this point is reached, the Aether starts up its propellers and the LRV is released it into the air. Project Aether
Figure 2 – Launching the Aether. The airbag systems keeps it in place until enough lift is generated by the Aether, at which point it is released. Recovering the Aether involves a similar although more complex process. The LRV navigates down the runway in a straight line, at the same approach speed as the Aether (i.e. same frame of reference). The Aether, using an ILS approach, and coordinating with the LRV local-GPS system, enters what is called the “secure-zone” within the LRV after which the top hinged surface descends and the airbag system is inflated (relatively slow to avoid structural damages) At this point, the Aether is firmly secured and the LRVs start to brake thus bringing the Aether to rest.
Figure 3 – Recovering the Aether. The ILS enables a smooth approach during which the Aether enters the “secure zone”. The top surface is lowered and the airbag system is activated. Below are two artistic designs of the Front LRV. At this design stage there are no specific justifiable dimensions since further investigations would be required. We believe however that an approximate sizing would be of the order of about 2.5 m wide and long vehicle, with a height of 2 m.
Figure 4 – Artistic design of Front LRV
Rear LRV The rear LRV is exactly the same as the front LRV only that it doesn’t provide a CRR capability and therefore has no clamping device. It is merely used to support the Aether’s tails due to its fragile nature.
Figure 5 – Artistic design of Rear LRV
Avionics The avionics of Aether are a central part to the aircrafts design, especially as it is an unmanned vehicle. This requires it to have multiple redundant systems and autopilots, in case one of them fails. Our core-avionics component is the RADA MAVINS (Modular Avionics), an all-in-one solution capable of the following functions: flight control surface management mission profile storage - autopilot Project Aether
Inertial navigation system (INS) GPS Air data centre Payload and data-link I/O interface MAVINS Flight Computer Power
110 mm W x 100 mm L x 78 mm H
The MAVINS has two identical sub-units, thus each providing a dual redundant solution. Aether will have two MAVINS thus providing a quad-redundant solution. MAVINS has a built-in autopilot unit, connected to the flight computer. This flight computer controls the actuators on the aircraft as specified by either the autopilot or the remote controller. The INS system in the MAVINS uses a micro-electrical-mechanical-system (MEMS) based on accelerometers and gyroscopes. This system allows you to know the position and attitude of the aircraft. As it is a mechanical system it accumulates errors over time, due to friction and inefficiencies, resulting in erroneous outputs. The INS is therefore coupled with GPS, a much more accurate positioning system, to correct these errors. We cannot use only GPS as it does not tell you the aircrafts attitude, but just its position. This requires a GPS antenna to be connected to the MAVINS. Our GPS antenna only weighs 45 g and is therefore very easy to incorporate into the aircraft. A thermocouple and Pitot tube are connected to each sub-unit allowing the MAVINS to determine velocity, altitude etc, by the Air data centre. Finally the MAVINS has an analog and digital I/O interface allowing various auxiliary components to connect to it. The choice of the payload is up to FATT04, and depends on the function of each Aether as established by FATT08. This allows the payload to communicate with the MAVINS and make use of its functions.
Power Management Power management is dealt with by FATT09, Power Systems. They will give us at least 120 W all year round which is sufficient to power all the essential avionics at once. The power is fed 50/50 into each MAVINS which then power all of the components attached to it, as they all require very little power.
Communication Aether will comprise a communications package based on UHF radio, which is between 300 and 3000 MHz. We chose this as our frequency range because of the very small antenna which it requires. Aether will have two boom-mounted blade antennas, pointing towards the ground to improve its range. Each antenna will be connected to its own transceiver, which is the link between the antenna and the MAVINS. The transceiver is connected to the MAVINSâ€™s mission profile unit where commands from the ground station are sent to. We chose a blade-style antenna to mininise the drag penalty on the aircrafts aerodynamics: Project Aether
UHF Blade Antenna Power
10 W: Receive and Transmit
30 mm W x 450 mm L x 220 mm H
C-TT/505 Transceiver Power
45.5 W: Receive and Transmit 4.9 W: Receive only
171.45 mm W x 128.27 mm L x 92.71 mm H
The communication system will be dual-redundant, so we will have two UHF transceivers each with its own antenna. Each transceiver will connect to one MAVINS, allowing us to save weight on cabling and providing a robust solution.
Camera We have incorporated a camera at the leading edge of the wing whose primary function is highresolution aerial photography, but will also function as a video camera if the aircraft is being remotely controlled. The camera we chose is the Prosilica GE4900, which has the following specifications: Prosilica GE4900 Power
66 mm W x 63 mm L x 66 mm H
It will be connected to the MAVINS by a 10/100 CAT5E Ethernet cable into the ethernet port on the MAVINS. The camera will not be redundant because it is not an essential system for the aircraft given that our autopilot is quad-redundant.
Actuators Each of the six control surfaces on Aether will be controlled by an electric actuator. These are small devices which use an electric motor to drive a lever arm which to control a mechanism. The actuator we chose for our aircraft is the Kearfott K2000, specifically designed for UAVâ€™s our size. Below are its specifications:
Kearfott K2000 Power
68.3 mm W x 111.8 mm L x 22.9 mm H
All of the actuators are dual redundant providing a very robust solution. We are able to achieve this configuration with little weight penalty because of the small and light actuators which we chose. Avionics Summary Average Power (W) Component mass (kg)
Wiring mass (kg) Total mass (kg)
Electrical System Design The diagram below illustrates how all the electric components on Aether will be connected together. The location of the components is determined by FATT04, weights and payloads.
AETHER ELECTRICAL SYSTEM DESIGN
Power & Propulsion Highlights Solar-Powered No refuelling required 1 Year endurance Flight Latitude 45oN â€“ 45oS Flight Altitude 40000ft minimum Configuration flexibility with 60kg maximum payload Continuous power to payload during day and night Semi-Geostationary Redundant Propulsion System High efficiency power and propulsion subsystems Efficiencies 96% Solar Capture (Optical) 15% Photovoltaic 91% Electrical 99% Battery 98% ESC (Motor Controller) 91% Motor 93% Propeller Power Characteristics 44.3kWh/day Minimum Power 116.5kWh/day Peak Power 28.0kWh Battery Storage Payload
Fulfilling the Requirements
Aether faces a number of technical challenges due to the operationrequirements imposed.
The analysis and design of the power and propulsion system goes through a number of stages, consists of three different elements.
Latitude Very low sun altitude Very long period of darkness Batteries Very large battery capacity is required for night operation Additional batteries required for avionics and payload Power Higher power requirement for flying at higher altitude Weight Massive solar panel array area and battery capacity means heavy weight Additional weight put strain to aerodynamic and structure
Optimization Result in an increases of system efficiency, bounded by the limit of physics Discovery Result in a decrease in system efficiency, from finding out unavoidable inefficiencies within the power-propulsion pathway. Assumption Used initially before indepth analysis taken place Used as a last resort, such as assuming a higher component operation value (e.g. efficiency, energy density) than currently available commercially.
214W Minimum Average Power 5000W Maximum Power # Subjected to electrical limitation and time of year
Design for Worst Case Operation Two weeks before Christmas to two weeks after Christmas (Add half a year for Southern Hemisphere) poses a significant challenge to Aether, where the source of power, is in a very short supply. The primary objective is to allow Aether to fly in such worst case scenario. As a result, Aether is heavily optimized for operating at this particular condition. When environmental condition improves, the performance of Aether is severely limited, by flying at off-design flight condition, despite a significant increase in available power.
Changed target requirement The original requirement asks for an operating altitude of 65000ft. The power requirement at 65000ft is 82% higher than 40000ft, which requires a significant increase in battery and PV cell weight. At 40000ft the air mass is stable enough for Aether to fly. In terms of business plan, effect of lowering the altitude is believed to be very small. In summer where the energy intake is considerably higher, Aether is capable of reaching height in excess of 76000ft. However, battery capacity and length of sunlight poses a limit, which Aether is unable to fly higher than 40000ft for a prolonged period. Power to Payload as Technology Improves
Flexible Design Aether has a huge wingspan as well as internal volume, which allow a highly flexible powerpropulsion system arrangement. For example when designed to fly at lower latitude, photovoltaic cells can be replaced by equivalent weight of battery, which provide considerably more power to the payload during the night. Or alternatively a different propeller can be used to optimize for flying faster or at higher altitude. (Subjected to time of year). It can also be changed by using the latest technology available over the lifespan of the airframe.
Power to Payload (W)
900 700 500 300 100 2009
Year Latitude (45')
Figure 1 â€“ Available Power to Payload based on predicted improvement of technology
Power Component Selection Battery Specification Chemistry Lithium Sulphur Gravimetric Density 300 Wh/kg  Volumetric Density 5.0875 Wh/m3  Battery Efficiency 99%
Photovoltaic Cell Specification Chemistry Copper Indium Gallium Selenide Panel Type Polymer thin-film Efficiency 15% (Projected)  Density 0.885 kg/m2 (Assumed) 
Battery Requirement Battery Capacity Battery Weight
Power Requirement Energy Demand Panel Area Panel Weight
Commercially available Lithium-Ion Polymer battery has an energy density of 150 Wh/kg which is impossible to fit in the available weight budget. Lithium-Sulphur battery is chosen specifically by its high energy density. Note that at the last publication of the technology (2004)  Lithium-Sulphur degrades significantly after each battery charge/discharge cycle. The capacity reduces to 57% after 150 charge cycles. This problem is assumed to be resolved by now (2009) and the capacity is projected to increase at a rate of 4%
44.3 88.9 78.7
kWh m2 kg
It is assumed thin-film CIGS cell will weight approximately the same or lower than existing thin-film PV cell. Hence the weight of the lightest thin-film cell available commercially now is used as a reference to calculate the PV cell areal density. The aircraft can fly with lower efficiency PV cell which is available in relatively large quantity, subjected to limit in wing area as well as power requirement of the payload. The efficiency is projected to increase at a rate of 4% a year in relative to current technology.
a year. Power Circuit Design 100
Power Available to Payload (W)
It is calculated that each Watt-Day of power is worth 70~93 grams of PV and Battery, and as such the efficiency of the electrical wiring can significantly impact the power available. And hence the wiring cannot be separated from power calculation. Note that minimum efficiency of each cable section is limited to 95% to provide achieve a set minimum voltage (10.8V / 21.6V). And also PV and Battery cost many times more than cables and it may be more ideal to pursuit a lower cost than absolute highest efficiency.
Variation of Power and Weight
Resistive Loss (W/m) Panel Weight
Figure 2 â€“ Available power against weight of components
Wire material Aluminium Resistive heat loss Wire weight Power to Payload
â‰¤ 8W/m 26.9 kg 214W
Propulsion Battery Photovoltaic Cell Avionics Payload Power Mngr Unit
Aluminium provided the lowest electrical resistance of all known non-reactive material per weight. Operating temperature also affects the electrical resistance. As the wiring is optimized for weight, high resistive loss may cause a high wire temperature. This temperature is not known without performing a detailed wiring analysis.
Each power management unit is isolated from another to reduce the length and mass of wiring required. Power management units are simple arrays of circuits which are highly reliable and the lifespan is considered to be in tens of years. Panel area and battery capacity allocated to each PMU is adjusted to provide just enough power day and night, thus eliminating the need of interconnection and achieve true redundancy. The position of battery packs, which contribute significantly to the cable weight, cannot be moved to optimal position as that will affect the loading of the structure. If time permits for FATT03/04, moving centre battery array closer to PMU can yield extra 100W of power without increasing weight, or alternatively translates to 9kg of weight reduction.
Propulsion Component Selection Propulsion in Aether is provided by 4 propulsion “packages” which can be broken down into different components. The whole package is embedded into the main wing, where the exposed part is covered in thin aluminium plate for improved motor cooling and efficiency. The two motors, turning in opposite direction, are connected to the main drive shaft through a 3.47:1 (Motor:Propeller) reduction gear made in stainless steel, which rotates the propeller. From propulsion point of view, the engine can be placed in any position as long as symmetricity is achieved. Motor Specification Model Number 5320/18  Maximum Rated Power Maximum Rated RPM Motor Type Motor kV Rating RPM/Volt Weight
AXI 1650W 10000 Brushless 370
Figure 3 – Propulsion Package Layout
Electronic Speed Control Model Number HobbyWing Platinum 70A  HV Opto Maximum Current 70A Operation Voltage Range 5V – 50V Fully Programmable Weight 82g
The reason for using dual motor is based on a number of considerations Each Propulsion Package can continue to operate with one failed motor Weight of extra motors (<4kg) are far less than extra wiring weight required for PMU interconnecting (>30kg) Provide additional thrust at ground level for climbing
Figure 4 – AXI 5320/18 HobbyWing ESC 
/ Figure 5 –
Operation Parameter Cruise Thrust 16.6 N (x4) Cruise Power 189 W (40kft) / 344W (65kft) (x8) Cruise RPM 1115 (40kft) / 2124 (65kft) Maximum RPM 2560 Voltage Range 6V ~ 24V (10.8V Cruise) The minimum voltage of the ESC is lower than shown on the left. The lower range is not used as the ESC efficiency is significantly lower below 90% throttle / input voltage.
Figure 7 – ESC Efficiency 
Efficiency at different altitude 92 91
90 89 88 87 86 85 84 83 0
High efficiency at worst month at 40000ft is achieved by using dual voltage rails. Between 40000ft ~ 45000ft, 12V rail is used. Above 45000ft, or high thrust is required, 24V rail is used. Voltage is electronically switched by the PMU. Optimizing Efficiency through Flight Planning
PV Cells are placed on the rear side of the top surface of the wing. This naturally gives an inclination of approximately 6.2o. This angle is significant when compared to the altitude (inclination) of the sun during worst-month. By optimizing the flight path such that the plane is always more-or-less facing the sun, the amount of energy intake is increased. This technique is called Solar-Tracking. However in this case the aircraft tracks the sun not the panels. The decided flight plan is based on previous restriction from FATT08 in terms of flight radius. Although the restriction had been lifted, the flight plan remained as to provide some flexibility in the event of needing it later. The impact of sub-optimal flight path on power intake is minimal (0.85% total power) and is used for reserve. Excess Power against Latitude 48.0
Typical Flight Plan (Winter)
200% 150% 100%
Longitude (Degrees) "Decided Flight Path"
"Optimal Flight Path"
Figure 8 â€“ Flight Planning
Figure 9 â€“ Available power at various Latitude
In Event of Power / Propulsion Failure
Aether cruise at 40000ft, well above the height of clouds. Solar energy intake is constant and reliable. Power system is sufficiently monitored such that any problem associated with battery charge level can be predicted hours before occurring. In the unlikely event of such failures, the operators will be alerted, and power restriction is applied to subsystems based on priority. Switching off power to payload is far more than sufficient to sustain flight until next sun-rise, or plan an emergency landing. PERFORMANCE:
EFFICIENCY VS ALTITUDE: Since density changes with altitude, the performance of the engine does too. The propeller forces and power are directly related to the fluid density. Indeed at higher densities the engine operates much better as less force is required to propel an equivalent mass of air as for 60000feet behind it. Since Aether flies between 40000 ft and 60000 ft, but also needs propellers efficient enough at ground level to allow for take off, the engines were optimized for an altitude of around 40000ft. SHAFT POWER:
Available Shaft Power at different Altitude 7
Maximum Power (kW) (Including Transmission Efficiency)
The propellers are limited in rpm by the motor performance. With the gear ration of 3.47:1, it can achieve about 2560 rpm. This means the shaft power is limited by this rpm. Since it is a function of the incoming flow velocity and the density of the flow, 2560 rpm will translate to varying shaft power at different altitudes implying that the higher aether climbs the less power will be available.
6 5 4 3 2 1 0 0
Altitude (ft) Total Power (Including Efficiency)
Power at Various Altitude 3
Power Requirement (kW)
2.5 2 1.5 1 0.5 0 680
Propeller RPM 00000 ft
N=1 Best Eff
N=1 Max Power
N=2 Best Eff
N=2 Max Power
Likewise the power requirement of the engine will vary with altitude and also with the flight speed. Indeed the faster aether flies the more power it will require, and more so at higher altitudes. This performance analysis is shown in the graph below. Aether only flies at a ground speed of about 7m/s, which equates to 30.6m/s at 40000 ft, and implies we only need around 450W of power at that speed and altitude.
Mach vs. Power Requirement 0.5
30kft 0.25 40kft 0.2
Choosing an airfoil for the blades required going through different airfoils and choosing one on which the L/D is maximized at the operating condition. However since the propellers needs to be able to work well at off design conditions, it was decided that the blades would be set at a slightly lower angle of attack so that it would remain efficient throughout its operating range. The airfoil that was chosen is a MH116 with 9.8%thickness. This is because it is sufficiently thin and flat to avoid supersonic flow at the tips but still sufficiently thick to produce high lift at the root.
Appendix    12   12  12 
"Lithium Sulfur Rechargeable Battery Data Sheet" Sion Power, Inc. Accessed 2009-06-12 http://www.sionpower.com/pdf/sion_product_spec.pdf Page 3, "WinHec 2004 White Paper" Sion Power, Inc. Accessed 2009-06-12 http://sionpower.com/pdf/articles/WinHEC2004WhitePaper.pdf “Best Research-Cell Efficiency” National Renewable Energy Laboratory Accessed 2009-06http://www.nrel.gov/pv/thin_film/docs/kaz_best_research_cells.ppt “PowerFilm Custom / OEM Product” PowerFilm, Inc. http://www.powerfilmsolar.com/products/custom/?voltages=ALL_PRODUCTS "AXI 5320/18" Espirit Model http://www.espritmodel.com/index.asp?PageAction=VIEWPROD&ProdID=7027 "HobbyWing Platinum 70A HV Opto ESC" Espirit Model http://www.espritmodel.com/index.asp?PageAction=VIEWPROD&ProdID=9983 “ESC Efficiency at Various Throttle Point” Matttay @ RCGroups http://www.rcgroups.com/forums/attachment.php?attachmentid=2389256
Accessed 2009-06-12 Accessed 2009-06-
Life Cycle Analysis Our aim for Aether was to try to follow the EU end-of-life vehicle (ELV) directive which states that a vehicle should, by 1 January 2015, have 95% reuse and recyclability (Directive 2000/53/EC). With this in mind we have performed a mixture of cradle-to-cradle and cradle-to-grave assessment to ensure that this target can be met. While the factory of the aircraft will not be built within the EU it was decided that this would be a good benchmark to pursue. Figure a. Global Life Cycle Analysis.
Production Carbon Fibre While it is not possible to fully estimate the production costs and emissions as many of the components of Aether are going to be sourced from other companies. Despite this the internal structure of Aether will be made in-house and as such we have set a target to offset the energy consumption of producing carbon fibre. In order to do this we have developed a strategy to utilize the surface of the roof of the factory by covering the area with solar cells. While this may not totally offset the emissions of the carbon fibre production it can take a step towards reducing the impact. Solar cells CIGS (Carbon Indium Gallium Selenide) cells are a relatively new technology compared to other forms of solar cells such as cadmium telluride or crystalline silicon cells and thus it is hard to estimate the energy intensity required to produce it. Despite this, manufacturers claim that the production of CIGS cells has a 95% material utilization (Nanosolar, 2009). Batteries There is little data available regarding the manufacture of lithium sulphur batteries however other forms of lithium primary batteries require a large energy input to manufacture (Ishihara, Kihira, Terada, & Iwahori, 2002).
Recycling Carbon Fibre Research has been conducted into the recyclability of carbon fibre composites and a company has been sourced (Recycled Carbon Fibre Ltd) that will accept waste carbon fibre reinforced plastics (CFRP) for no charge. They will recover about 90% of the original fibres and these will be recycled into other noncritical uses such as airline seats, medical applications, road surfaces and other applications. Research has also shown that some recycled fibres can improve the performance compared to virginal fibres (McNally, Boyd, McClory, Bien, Moore, & Millar, 2007). It is believed that if this research can be followed up the next iterations of Aether can potentially have a much smaller environmental impact in its production stages; thus potentially having a closed loop in the CFRP life cycle as can be seen figure a.
Figure b. Carbon Fibre Recycling Processes (Pickering, 2006) Solar Cells CIGS solar cells have been used to provide power to the components of Aether. Due to the recent development nature of these cells there is no clear defined method to recycle it although some research has been undertaken and various methods have been introduced that can capture many of the precious materials that are used such as indium (Gaiker, 2006).
Figure c. CIGS solar cell recycling. Batteries In order to reduce the total life cycle impact of batteries a company named TOXCO has been found that will recycle lithium sulphur batteries for a cost of CAD $2.80/lb which roughly equates to GBP ÂŁ3.80/kg. While this will increase the cost of Aether the benefit far outweighs the cost. In addition to this TOXCO is based in British Columbia, Canada which utilizes hydroelectric power thus reducing the carbon footprint of the recycling stage of the battery (Bruce, 2009).
Figure d. Battery Recycling Processes (Bruce, 2009). Avionics, Motors & Cabling Another company has been sourced for recycling of avionics named Blue Star Electronics LLC based in California, USA. The company will attempt to recycle and resell equipment; if neither of these is possible the company will attempt to destroy the equipment in an appropriate manner. Recycling of the motors has been sourced and the company from whom the motors have been sourced will accept motor parts to be recycled. (Vaclavik, 2009) Cables do not pose a major concern for recycling; there are many companies that will recover the materials in cabling.
Satellite Comparison One of the main objectives for Project Aether was to create a design which has the capability to replace satellites. With this in mind, the life cycle of Aether has been compared with that of a satellite with a payload capacity of 60kg, the same as Aetherâ€™s. The major components of both a satellite and Project Aether are solar cells and structure. These solar cell technology used on satellites has been assumed to be CIGS and the structure has been assumed to be a carbon reinforces polymer composite. Due to the confidentiality surrounding solar cell development and the satellite industry as a whole, production energy data for solar cells and the full satellite are very difficult to source. This is why such crude estimates have had to be used. However, using estimates based on a solar energy journal, (Kato, Hibino, Komoto, Ihara, Yamamoto, & Fujihara, 2001) solar cell production for both satellite and Aether can be estimated at 1500 MJ/m 2. Both cases require electrical energy for avionics and payload but Aether requires it for propulsion also, causing the solar cell area for Aether to be much larger. This explains the much higher production energy requirement for Aether in terms of solar panels. The structure of the satellite has been assumed to be very light given the minimal loads experienced by the craft when in use. Aether, however, is a lift generating aircraft requiring a load bearing structure. Thus, the satellite structure and therefore the production energy has been estimated as half that of
Aether. The production energy requirement of Aether has been estimated using data from (Recycled Carbon Fibre Ltd, 2008) a carbon fibre recycling company. Given these assumptions and estimates, the following data can be calculated. Battery production energy has not been included as data on this subject is extremely difficult to source.
Solar cells Structure Major components production energy
Satellite 60000 11750 72
MJ MJ GJ
Aether 180000 23500 204
MJ MJ GJ
In terms of production the satellite is less energy intensive. Despite this being a rough estimate, it can be seen that Aether requires almost three times the energy of a satellite with a similarly sized payload. However, an integral part of a satellite system is the launch vehicle. This, in nearly all cases is a rocket and will require an amount of energy far surpassing that of either Aether or a satellite to produce. As the rocket is a one-time-use vehicle, the production of the rocket should be included in the environmental impact for the satellite system. In this case the Aether aircraft provides a reduced environmental impact. In operations, the satellite requires one huge input of energy as a launch but after that it requires and indeed has no way of receiving external sources of energy. The launch requires the burning of a massive amount of solid rocket fuel. On the other hand, Aether requires a very small amount of energy to launch as it only needs to be accelerated to 20km/h in order to take off. This is achieved by using LRV as described earlier in this report. These have the capability of being electrically powered with the energy being renewably sourced. In the operations case Aether has a much smaller environmental impact compared to a satellite. The disposal method for the satellite and Aether are very different indeed. Companies and methods have been sourced for the recycling or safe disposal of the entire Aether aircraft. On the other hand, the method of disposal for a typical satellite is to change orbit so that re-entry into the earthâ€™s atmosphere occurs. This results in the burning up of the constituent material of the satellite and its dispersal within the atmosphere. Given the toxic and potentially dangerous materials within electronics and solar cells, this is a highly undesirable occurrence. In this part of the life cycle, Aether allows a greatly reduced environmental impact. In conclusion, despite the greater production energy of Aether compared to a satellite, the ancillary equipment required for a satellite and the environmental impact of these makes it have a much greater environmental impact overall. Coupling this with the huge benefits of aircraft disposal compared with satellite disposal, Aether clearly provides an environmentally superior alternative to traditional satellite technology.
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