VIRGINIA POLYTECHNIC INSTITUTE AND STATE UNIVERSITY
Large Crowd Surveillance Unmanned Aerial Vehicle
2007 – 2008 International Design Project
Belle Bredehoft
Robert Briggs
Amanda Chou
Richard Duelley
Alex Kovacic
Jessica McNeilus
Philip Pesce
Dennis Preus
Megan Prince
Anthony Ricciardi
Michael Sherman
Erik Sunday
Tuesday, April 29, 2008
Table of Contents
Large Crowd Surveillance Unmanned Aerial Vehicle ................................................................... 1
2007 – 2008 International Design Project ................................................................................... 1
List of Variables ............................................................................................................................ 11
Introduction ................................................................................................................................... 15
Conceptual Design Requirements ................................................................................................. 16
Comparator UAVs ........................................................................................................................ 16
Comparator UAV Configurations ............................................................................................. 17
Conventional Tractor ............................................................................................................. 17
Flying Wing ........................................................................................................................... 18
Pylon-Mounted Propeller ...................................................................................................... 18
Twin-Tail-boom Pusher ......................................................................................................... 18
Comparator UAV Review ......................................................................................................... 19
Initial Virginia Tech Overall Conceptual Designs........................................................................ 19
Conceptual Analysis ..................................................................................................................... 20
Max Gross Takeoff Weight (MGTOW).................................................................................... 20
Wing Sizing and Placement ...................................................................................................... 20
Power Required and Endurance ................................................................................................ 21
Deployable Landing Gear System............................................................................................. 21
Landing Techniques .................................................................................................................. 21
General Geometry ..................................................................................................................... 22
Conventional Design ............................................................................................................. 22
Twin Tail-Boom .................................................................................................................... 23
Pylon Mount .......................................................................................................................... 23
Tail-Mounted Pusher ............................................................................................................. 24
Concept Comparison ..................................................................................................................... 25
Decision Matrix ............................................................................................................................ 25
Group A Decision Matrix .......................................................................................................... 26
Group B Decision Matrix .......................................................................................................... 28
Reliability...................................................................................................................................... 30
Final Group Concepts ................................................................................................................... 32
Group A Final Concept ............................................................................................................. 32
Constraint Analysis ................................................................................................................... 34
Sizing......................................................................................................................................... 35
2
Wing ...................................................................................................................................... 35
Fuselage ................................................................................................................................. 35
Tail ......................................................................................................................................... 35
Performance Analysis ............................................................................................................... 36
Stall ........................................................................................................................................ 36
Power Required – Straight and Level .................................................................................... 36
Power Required – Climb ....................................................................................................... 36
Engine – Power Available ..................................................................................................... 37
Endurance .............................................................................................................................. 38
Glide Range ........................................................................................................................... 38
Turn Rate ............................................................................................................................... 39
Group B Final Concept ................................................................................................................. 40
Constraint Analysis ................................................................................................................... 41
Airfoil Selection ........................................................................................................................ 42
Tail Sizing ................................................................................................................................. 43
Engine Selection, Power Requirements, and Endurance .......................................................... 44
Starters and Alternators ......................................................................................................... 46
Stability ..................................................................................................................................... 47
Qualitative Reliability Analysis ................................................................................................ 48
Final Values............................................................................................................................... 49
Loughborough University Design Process ................................................................................... 50
Downselection............................................................................................................................... 52
Final Concept ................................................................................................................................ 54
Constraint Analysis ................................................................................................................... 55
Wing Sizing ............................................................................................................................... 56
Tail Sizing ................................................................................................................................. 57
Performance Analysis ............................................................................................................... 57
Preliminary Design Phase ............................................................................................................. 62
Aircraft Overview ......................................................................................................................... 62
Requirements Met ..................................................................................................................... 63
Structural Design .......................................................................................................................... 64
Rectangular Fuselage ................................................................................................................ 64
Keel Design ............................................................................................................................... 64
Materials .................................................................................................................................... 65
3
Modularity of Design ................................................................................................................ 65
The Fuselage Section ............................................................................................................. 65
The Wing Section .................................................................................................................. 66
The Tail Assembly................................................................................................................. 66
Stress Calculations .................................................................................................................... 67
Weights and Balances ................................................................................................................... 67
Moments of Inertia .................................................................................................................... 67
Center of Gravity Location ....................................................................................................... 68
Size Comparison ....................................................................................................................... 69
Aerodynamics ............................................................................................................................... 69
Overall Aerodynamics............................................................................................................... 69
Airfoil Selection ........................................................................................................................ 70
Drag Buildup ............................................................................................................................. 70
Prop Wash Effects ..................................................................................................................... 72
Tail Sizing ..................................................................................................................................... 73
Control Surfaces............................................................................................................................ 73
Ailerons ..................................................................................................................................... 74
Elevator ..................................................................................................................................... 74
Rudder ....................................................................................................................................... 74
Additional Control Surface Points ............................................................................................ 75
Stability Analysis .......................................................................................................................... 76
Stability Derivatives for a Dihedral Wing................................................................................. 76
Stability Derivatives for a Straight Wing .................................................................................. 77
Static Stability ........................................................................................................................... 78
Longitudinal Stability ............................................................................................................ 79
Directional Static Stability..................................................................................................... 79
Lateral Static Stability ........................................................................................................... 79
Dynamic Stability...................................................................................................................... 80
Longitudinal Motion .............................................................................................................. 80
Lateral Motion ....................................................................................................................... 81
Propulsion ..................................................................................................................................... 82
Power Required ......................................................................................................................... 82
The Internal Combustion Engine .............................................................................................. 83
Specific Fuel Consumption ................................................................................................... 83
4
Propeller Sizing ..................................................................................................................... 85
Parts Required for Internal Combustion Engine.................................................................... 86
The Electric Engine ................................................................................................................... 86
Analysis ................................................................................................................................. 87
Propeller Selection ................................................................................................................. 89
Parts Required for an Electric Engine ................................................................................... 89
IC/IC Configuration .................................................................................................................. 89
IC/Electric Configuration .......................................................................................................... 90
Noise Prediction ........................................................................................................................ 90
Noise Prediction Equations ................................................................................................... 91
Prediction Using Xrotor......................................................................................................... 92
Conclusions on Noise Levels ................................................................................................ 93
Advantages and Disadvantages ................................................................................................. 94
Performance .................................................................................................................................. 95
Breguet Equations: Endurance and Range ................................................................................ 95
Glide Range ........................................................................................................................... 96
Landing Distances ..................................................................................................................... 97
Assumptions .......................................................................................................................... 97
Ground Roll Stopping Distances ........................................................................................... 97
Secondary Capture Devices ................................................................................................... 98
Systems Integration ....................................................................................................................... 99
Electromagnetic Shielding ...................................................................................................... 101
Autopilot.................................................................................................................................. 101
Batteries ................................................................................................................................... 102
Servo Sizing and Selection ...................................................................................................... 104
Ground Control ........................................................................................................................... 104
Autopilot Ground Control Station ........................................................................................... 104
Launcher .................................................................................................................................. 106
Landing Gear .............................................................................................................................. 106
Configuration .......................................................................................................................... 106
Sizing and Structure ................................................................................................................ 107
Wheels and Brakes .................................................................................................................. 107
Reliability.................................................................................................................................... 108
Fault Tree Analysis (FTA) ...................................................................................................... 108
5
Failure Modes and Effects Analysis (FMEA) ......................................................................... 108
Mechanical Failure Models ..................................................................................................... 109
Reliability Data ....................................................................................................................... 109
Propulsion Reliability.............................................................................................................. 109
Autopilot Reliability................................................................................................................ 112
Structural Reliability ............................................................................................................... 112
Overall Reliability ................................................................................................................... 114
Costs............................................................................................................................................ 114
Cost Model .............................................................................................................................. 115
Component Cost .................................................................................................................. 115
Operational Costs ................................................................................................................ 116
Maintenance Cost ................................................................................................................ 117
Support Cost ........................................................................................................................ 117
Manufacturing Cost ............................................................................................................. 117
Overall Costs ....................................................................................................................... 118
Conclusions ................................................................................................................................. 118
Further Work ............................................................................................................................... 118
Desired Work .......................................................................................................................... 119
Appendix A – List of Equations ................................................................................................. 120
Appendix B – Initial Concept Drawings ..................................................................................... 123
Concept A1 .......................................................................................................................... 123
Concept A2 .......................................................................................................................... 124
Concept A3 .......................................................................................................................... 125
Concept A4 .......................................................................................................................... 126
Concept A5 .......................................................................................................................... 127
Concept B1 .......................................................................................................................... 128
Concept B2 .......................................................................................................................... 129
Concept B3 .......................................................................................................................... 130
Concept B4 .......................................................................................................................... 131
Concept B5 .......................................................................................................................... 132
Appendix C – Decision Matrix Group A .................................................................................... 133
Appendix D – Decision Matrix, Group B ................................................................................... 136
Appendix E – Constraint Analysis Script ................................................................................... 138
Appendix F – Endurance Script .................................................................................................. 139
6
Appendix G – Power Script ........................................................................................................ 140
Appendix H – Constraint Analysis Script ................................................................................... 141
Appendix I – Final Design CAD 3-View ................................................................................... 144
Appendix J – Stability Derivatives and Crosswind Script .......................................................... 145
Appendix K – Specific Fuel Consumption Script....................................................................... 147
Appendix L – Internal Combustion Engine Propeller ................................................................ 148
Appendix M – Electric Engine Propeller .................................................................................... 149
Appendix N – Electric Propulsion Battery Requirements .......................................................... 150
Appendix O – Noise Footprint Output from Xrotor ................................................................... 151
Appendix P – Landing Matlab Script ......................................................................................... 157
Appendix Q – Mechanical Failure Model Method ..................................................................... 159
7
Table of Figures
Figure 1. Arcturus T-16XL and EMIT Blue Horizon 2 ................................................................ 17
Figure 2. Boeing/Insitu ScanEagle ............................................................................................... 18
Figure 3. The Orca Light Sport Amphibian .................................................................................. 18
Figure 4. AAI Pioneer UAV and AAI RQ-7 Shadow 200 ........................................................... 18
Figure 5. Viking 300 ..................................................................................................................... 19
Figure 6. Average Sources of System Failures For U.S. Military UA Fleet (Based on 194,000
hours)9 ........................................................................................................................................... 30
Figure 7. Constraint Analysis ....................................................................................................... 34
Figure 8. Power Available Plot for Various Altitudes .................................................................. 37
Figure 9. New Conceptual Design ................................................................................................ 39
Figure 10. Exploded View of New Concept ................................................................................. 40
Figure 11. Final Concept – Group B ............................................................................................. 41
Figure 12. Constraint Analysis Curves for Final Concept ............................................................ 42
Figure 13. Cruise Speed and Stall Speed vs. Time ....................................................................... 45
Figure 14. Power Required Curve with 15hp Available ............................................................... 45
Figure 15. AVL Geometry Plot .................................................................................................... 47
Figure 16. LU Conventional Concept ........................................................................................... 50
Figure 17. LU Pylon-Mounted Engine Concept ........................................................................... 51
Figure 18. Final Concept Drawing................................................................................................ 54
Figure 19. Constraint Analysis for New Conceptual Design ........................................................ 56
Figure 20. Power Curves............................................................................................................... 59
Figure 21. Final Concept 3-View.................................................................................................. 61
Figure 22. Pylon Mounting Structure .......................................................................................... 65
Figure 24. Detached Wing Section .............................................................................................. 66
Figure 23. Fuselage with No Skin................................................................................................. 66
Figure 25. Tail Assembly ............................................................................................................. 66
Figure 26. CG and Neutral Point Locations................................................................................. 68
Figure 27. Six-Foot Man with Vulture ......................................................................................... 69
Figure 28. Angles Showing Blanketed Regions (Raymer) .......................................................... 72
Figure 29. Raymer's Angles for Spin Recovery ........................................................................... 75
Figure 30. Thumbprint Criterion for Short Period Mode Handling ............................................. 80
Figure 31. Society of Automotive Engineers (SAE) SFC vs. RPM Curve ................................. 84
Figure 32. Appendix K Script Output for Specific Fuel Consumption Estimates ....................... 84
Figure 33. Propeller Design Plot .................................................................................................. 85
Figure 34. Thrust and Drag vs. Velocity from Analytical Method ............................................... 87
Figure 35. Maximum Allowable Headwind vs. Weight of Battery Packs................................... 88
Figure 36. Diagram of Pylon Packaging for the IC/Electric Configuration ................................ 90
Figure 37. Ground Roll with No Brakes or Flaps ........................................................................ 99
Figure 38. Cloud Cap Piccolo II Autopilot ................................................................................ 102
Figure 39. Ground Control Architecture .................................................................................... 105
Figure 40. Cloud Cap Ground Station ....................................................................................... 105
Figure 41. Landing Gear Sizing11 .............................................................................................. 107
Figure 42. Internal Combustion Engine Fault Tree ................................................................... 110
Figure 43. Electric Engine Fault Tree ........................................................................................ 111
8
Figure 44. Autopilot Fault Tree ................................................................................................. 112
Figure 45. Overall System Fault Tree ........................................................................................ 114
Figure 46. Cost of Aircraft Components..................................................................................... 116
Figure 47. Noise Footprint at Ground of Aircraft at 200 ft AGL at 4500 RPM w/ 2 Blades .... 151
Figure 48. Noise Footprint at Ground for Aircraft at 2000 ft AGL at 4500 RPM w/ 2 Blades . 151
Figure 49. Noise Footprint at Ground for Aircraft at 3000 ft AGL at 4500 RPM w/ 2 Blades . 152
Figure 50. Noise Footprint at Ground for Aircraft at 200 ft AGL at 4500 RPM w/ 3 Blades ... 152
Figure 51. Noise Footprint at Ground for Aircraft at 2000 ft AGL at 4500 RPM w/ 3 Blades . 153
Figure 52. Noise Footprint at Ground for Aircraft at 3000 ft AGL at 4500 RPM w/ 3 Blades . 153
Figure 53. Noise Footprint at Ground for Aircraft at 200 ft AGL at 6000 RPM w/ 2 Blades ... 154
Figure 54. Noise Footprint at Ground of Aircraft at 2000 ft AGL at 6000 RPM w/ 2 Blades ... 154
Figure 55. Noise Footprint at Ground of Aircraft at 3000 ft AGL at 6000 RPM w/ 2 Blades ... 155
Figure 56. Noise Footprint at Ground for Aircraft at 200 ft AGL at 6000 RPM w/ 3 Blades .... 155
Figure 57. Noise Footprint at Ground for Aircraft at 2000 ft AGL at 6000 RPM w/ 3 Blades .. 156
Figure 58. Noise Footprint at Ground for Aircraft at 3000 ft AGL for 6000 RPM w/ 3 Blades 156
Figure 59. Mechanical Failure Method ...................................................................................... 159
Table of Tables
Table 1. Comparator UAV Review .............................................................................................. 19
Table 2. Concept Comparison....................................................................................................... 25
Table 3. Summarized Decision Matrix – Group A ....................................................................... 27
Table 4. Summarized Final Decision Matrix – Group B .............................................................. 29
Table 5. Initial Empennage Sizing Parameters ............................................................................. 35
Table 6. Stall Speed versus Altitude ............................................................................................. 36
Table 7. Endurance at Cruise (3,000 ft, 50 kts) ............................................................................ 38
Table 8. Glide Range and Glide Speed versus Absolute Altitude ................................................ 38
Table 9. Airfoil Comparison ......................................................................................................... 42
Table 10. Results of Airfoil Ratings ............................................................................................. 43
Table 11. Initial Empennage Sizing Parameters ........................................................................... 44
Table 12. Engine Specifications for Engine Comparison ............................................................. 45
Table 13. Engine Selection Decision Matrix ................................................................................ 46
Table 14. Final Values for Group B Concept ............................................................................... 49
Table 15. LU Conventional Concept Sizing ................................................................................. 50
Table 16. LU Pylon-Mounted Concept Sizing ............................................................................. 51
Table 17. LU Final Concepts Advantages and Disadvantages ..................................................... 52
Table 18. Assumptions for Constraint Analysis ........................................................................... 55
Table 19. Wing Sizing Parameters................................................................................................ 56
Table 20. Initial Empennage Sizing Parameters ........................................................................... 57
Table 21. Assumptions for Performance Analysis ....................................................................... 58
Table 22. Stall Speed at Various Altitudes ................................................................................... 60
Table 23. Team Responsibilities .................................................................................................. 62
Table 24. Key Dimensions of the Vulture ................................................................................... 63
Table 25. Mass Breakdown of Vulture (IC/Electric Configuration) ........................................... 63
Table 26. Requirements and Performance Comparison .............................................................. 64
Table 27. Stress Calculations ....................................................................................................... 67
9
Table 28. Moments of Inertia for the IC/IC Configuration ......................................................... 68
Table 29. Moments of Inertia for the IC/Electric Configuration .................................................. 68
Table 30. Wing Dimensions ........................................................................................................ 70
Table 31. Airfoil Selection Characteristics ................................................................................... 70
Table 32. Drag Build-Up with Raymer and Tornenbeek's Approximations ............................... 71
Table 33. Drag Estimations from Mason's Program .................................................................... 71
Table 34. Prop Wash Effects on the Tail ...................................................................................... 72
Table 35. Summary of Tail Dimensions ...................................................................................... 73
Table 36. AVL Output for a Dihedral Wing ................................................................................. 76
Table 37. AVL Output for a Straight Wing ................................................................................. 78
Table 38. Neutral Point and Static Margin Calculation ................................................................ 79
Table 39. Power Required for Different Conditions .................................................................... 83
Table 40. Summary of Noise Prediction Values .......................................................................... 92
Table 41. Noise Footprint Approximations using Xrotor ............................................................ 93
Table 42. Advantages and Disadvantages of IC/IC Configuration.............................................. 94
Table 43. Advantages and Disadvantages of IC/Electric Configuration ..................................... 95
Table 44. Endurance at Constant Angle of Attack and Constant Velocity ................................... 96
Table 45. Powered Range for Different Engine Configurations.................................................. 96
Table 46. No-power Glide Range ................................................................................................. 96
Table 47. Ground Roll Distances for Vulture .............................................................................. 98
Table 48. Parts List .................................................................................................................... 100
Table 49. Batteries Chosen for Vulture ..................................................................................... 102
Table 50. Structural Reliability Calculations .............................................................................. 113
Table 51. Operational Costs of Vulture ...................................................................................... 118
Table 52. Number of Batteries vs. Max Headwind ................................................................... 150
10
List of Variables
Variable
Description
Units
a
acceleration
[ft/s2]
A
Cross-sectional area of propeller airflow
[ft2]
AR
Aspect ratio
[-]
b
Wing span
[ft]
B
Number of propeller blades
[-]
c
Average chord length
[ft]
CD
Coefficient of drag
[-]
CD0
Coefficient of base drag
[-]
C DMD
Coefficient of drag at minimum drag
[-]
C DMP
Coefficient of drag at minimum power
[-]
CF
Skin friction coefficient
[-]
Cht
horizontal tail volume coefficient
[-]
CL
Coefficient of lift
[-]
C Lmax
Maximum lift coefficient = 1.3
[-]
CL/CD
Lift to drag ratio
[-]
Maximum lift to drag ratio
[-]
Lift to drag ratio at minimum power
[-]
Coefficient of lift at zero-angle of attack
[-]
CLMD
Coefficient of lift at minimum drag
[-]
C LM P
Coefficient of lift at minimum power
[-]
Cm
Moment coefficient
[-]
Cvt
Vertical tail volume coefficient
[-]
Δm
Wing sweep at maximum thickness
[°]
d
diameter
[ft]
D
Drag on aircraft
[lbs]
dB(A)
Weighted sound level
[dB]
DI
Noise directivity index
[dB]
dh/dt
Rate of climb at constant speed
[ft/s]
dV/dt
change in velocity with time (ft/s2)
[ft/s2]
CL / CD max
CL / CD MP
CL0
e
emf
Oswald efficiency factor = 0.9
Electromotive force of batteries
[-]
[W]
E
Endurance
F
Form factor for drag build-up
[lbs]
Drag force
[lbs]
Fdrag
[sec], [hr]
11
Ffriction
Friction force
[lbs]
Fmain, friction
Friction force contribution from main gear
[lbs]
Ftail, friction
Friction force contribution from tail skid
[lbs]
FL1
Farfield noise contribution due to power, tip speed
[dB]
FL2
Farfield noise contribution due to blade number, diameter
[dB]
FL3
Farfield noise contribution due to atmospheric absorption
[dB]
g
Acceleration due to gravity
[ft/s2]
γp
Weight specific fuel consumption
h
Altitude
[ft]
I
Current or amperage
[A]
Estimated propeller constant from wind tunnel
[-]
kprop
L
Lift on aircraft
[lbs/hp-hr]
[lbs]
L/D
Lift to drag ratio
[-]
Lht
horizontal tail moment arm
[ft]
Lvt
vertical tail moment arm
[ft]
µmain
Frictional coefficient from main gear
[-]
µtail
Frictional coefficient from tail skid
[-]
m
Mass of subscripted component
M
Mach number
MEP
MGTOW
n
[slug]
[-]
Mean effective pressure of an engine
Maximum gross takeoff weight
[psi/cycle]
[lbs]
Load factor
[-]
ηelec
Electrical efficiency of an engine
[-]
ηmech
Mechanical efficiency of an engine
[-]
Propeller efficiency
[-]
Number of cylinders in an engine
[-]
ηp
Npiston
NC
Noise correction for multi-bladed propellers
[dB]
Overall sound pressure level
[dB]
p
Pitch angle
[in.]
P
Engine power
[ft-lbs/s], [hp]
Pavail
Power available from engine
[ft-lbs/s], [hp]
Preq
Power required from engine
[ft-lbs/s], [hp]
PF
Propeller factor
OSPL
PNL
[-]
Perceived noise pressure level
[dB]
[lbs/ft2]
q
Dynamic pressure
Q
Interference factor for drag build-up
12
[-]
R
Range
[ft], [n.m.]
Rbatt
Battery resistance
[Ω]
Resc
Electronic speed controller resistance
[Ω]
Rglide
Gliding range
Rmotor
Resistance of electric motor
[Ω]
Rtotal
Total resistance of electric motor
[Ω]
Rwires
Resistance of wires
[Ω]
Reynolds number
[-]
Re
RPM
[ft], [n.m.]
Rotations per minute of a propeller
[rotations/min]
ρ
Ambient air density
[sl/ft3]
s
Net displacement distance
[in], [ft]
S
Wing planform area
[ft2]
Sht
Horizontal stabilizer surface area
[ft2]
Sref
Reference area for drag calculations
[ft2]
Svt
Vertical stabilizer surface area
[ft2]
Swet
Wetted area for skin friction calculations
[ft2]
SFC
Specific fuel consumption
SHP
Shaft horsepower
t
time
T
Thrust from engine
T/W
(T/W)SL
[lbs/hp-hr]
[hp]
[s], [hr]
[lbs]
Thrust to weight ratio
[-]
Thrust to weight ratio at straight and level flight
[-]
Volume of engine cylinder/piston
[ft3]
V
Aircraft velocity
[ft/s]
V0
Initial aircraft velocity
[ft/s]
V1
Final aircraft velocity
[ft/s]
Va
Aircraft approach velocity
[ft/s]
Velec
Voltage of electric engine
[V]
VMD
Aircraft velocity at minimum drag
[ft/s]
VMP
Aircraft velocity at minimum power required
[ft/s]
Vstall
Aircraft velocity at stall
[ft/s]
vcyl
𝑤̇ f
Fuel flow rate of an engine
W
Weight of aircraft
[lbs]
W1
Maximum gross takeoff weight
[lbs]
W2
Empty weight
[lbs]
Weight supported by main gear
[lbs]
Wmain
[lbs/s]
13
Wtail
Weight supported by tail skid
W/S
Wing loading
[lbs/ft2]
Wing loading at stall
[lbs/ft2]
(W/S)stall
x
Horizontal position (generic designation)
14
[lbs]
[ft]
Introduction
In the modern world of aerospace engineering, an increasing emphasis on safety, performance, and reliability has driven aircraft designers to develop a variety of autonomous vehicles.
Uninhabited autonomous vehicles (UAVs) have a larger performance envelope, such as increased maneuverability previously limited by the g-forces able to be withstood by humans.
Likewise, the vehicle’s size can be greatly reduced, improving both efficiency and cost. Being
able to predict the vehicle’s timely and consistent response to instructions, as well as the possibility of system or structural failures, the overall reliability of the vehicle can be monitored.
The International Design Team in the Aerospace Engineering departments at Virginia
Tech and Loughborough University has been tasked with the development of a UAV whose purpose is flying surveillance cover over a crowd of interest. The task objectives and requirements
have been developed by Naval Air Systems Command (NAVAIR), the Patuxent River Naval Air
Station in Maryland. Part of NAVAIR’s request is to not only develop an aerial vehicle that can
accomplish the primary objective of surveillance, but also to study the impact of design on aircraft reliability. Specific emphasis is placed upon cost drivers for improved reliability and at
what point does increasing the cost no longer significantly improve the overall integrity of the
vehicle. From this analysis, a better understanding of UAV reliability can be obtained.
To start off the design process, comparator UAVs were researched in order to investigate
what configurations of aircraft were used to fulfill mission requirements. Each university group
involved with the project was then divided up into conceptual design groups. Each Aerospace
Engineering student on the team developed an individual conceptual design which met as many
of the design requirements as possible. The Industrial Systems Engineer on each Virginia Tech
team then created and executed a decision making process as to which design from the five conceptual designs would be used as the team design. The two team designs from Virginia Tech
were presented along with the designs from Loughborough University. A decision was made as
to which design would be used for the remainder of the project, after which, both university
groups will work together to fully develop this design.
NAVAIR will sponsor this design project throughout the next two years and the two corresponding design phases. The first year will focus on the conceptual design of the vehicle,
whereas the second year will focus on the construction and flight testing of the UAV prototype.
The scope of this report is focused on the first year of the program, conceptual design.
15
Conceptual Design Requirements
The objective of this project is to design and conduct prototype development of a remotely piloted UAV that has the following requirements:

Cruise speed of 50 knots (kt)

Top Speed of 70 kt

Range of 15 n.m. (nautical miles)

Minimum endurance of eight hours

Service ceiling of 10,000 ft at half fuel

Normal operational altitude of 3000 ft or 2000 ft above ground level (AGL)

Minimum turn rate of 6 degrees per second

Climb rate of at least 200 ft/min at sea level

Maximum Gross Takeoff Weight (MGTOW) of 300 lbs

Minimum payload of 30 lbs (45 lbs desired)

Payload power source of 10 watts

Noise levels below 50 dB(A) at 200 ft.

All weather operation with a 10 kt crosswind landing capability

Capable of rail catapult pneumatic launch

Landing within a 50 ft x 250 ft parking lot

Less than one flight failure per 100,000 hours of flight
Additionally, the vehicle must be capable of GPS based autonomous operations with dy-
namic re-tasking from ground controllers. In the event of lost communications between the vehicle and the ground station, the vehicle must be capable of autonomous flying to a predetermined location in an attempt to restore communications. Likewise, the vehicle must be capable
of gliding to a predetermined point in the event of an engine-out condition.
Comparator UAVs
In order to start the design process, several comparator UAVs were investigated for their
similarity to mission requirements. Two groups of existing UAVs were found: those that fit the
weight requirement and those that fit the speed or endurance requirement. Once the comparator
16
aircraft were identified, they were split between the two groups and their configuration was noted. The configurations included conventional tractor, flying wing, pylon-mounted engine, and
twin-tail-boom pusher. From these observations, conceptual designs could be modeled after
these aircraft configurations to fit all of the design requirements.
It is apparent that some of the vehicles more closely match some of the design requirements than others. The performance values that are similar to the design requirements are underlined in the table. For example, the T-16 matched the desired airspeeds, whereas its weight and
payload are too low. Likewise, the Shadow’s MGTOW and payload weight are close to desired
while its airspeed range is not. The Viking 300 is the comparator UAV that most resembles the
design requirements in almost every category: speed, weight, endurance, payload.
Comparator UAV Configurations
Each of the comparator UAVs were divided into different configuration categories.
These configurations include: conventional tractors, flying wings, pylon-mounted pushers, and
twin-tail boom pushers.
Conventional Tractor
Figure 1. Arcturus T-16XL1 and EMIT Blue Horizon 22
The Arcturus T-16XL and EMIT Blue Horizon 2 (Figure 1) are two versions of a single
configuration that fit different requirements. The T-16XL matches the maximum and cruise
speeds and exceeds the endurance specified for the project. However, it carries a lighter payload
and has a smaller maximum gross takeoff weight. Specific characteristics of the T-16XL that
stand out are that it is also rail-launched and can be landed conventionally. The Blue Horizon
carries a heavier payload, weighs slightly more than the MGTOW and has greater endurance
than required, but has a much faster cruise and maximum speed.
17
Flying Wing
Figure 2. Boeing/Insitu ScanEagle3
The ScanEagle (Figure 2) is the only flying wing configuration and fits only the speed
and endurance requirements. Like the T-16XL, this aircraft is underweight and carries a much
lighter payload than specified. Interesting characteristics of the ScanEagle, however, include its
capture method, which involves capturing a shock cord on a pole. This alternate landing method
was investigated as a possibility to be used with some of the conceptual designs.
Pylon-Mounted Propeller
Figure 3. The Orca Light Sport Amphibian4
The Orca (Figure 3) is the only pylon-mounted engine configuration and does not fit any
of the requirements listed in Table 1. However, the Orca was used as a comparator aircraft due
to the investigation of noise reduction. For this reason, the propeller is mounted in a duct above
the aircraft.
Twin-Tail-boom Pusher
Figure 4. AAI Pioneer UAV5 and AAI RQ-7 Shadow 2006
18
Most of the comparator UAVs investigated were twin-tail-boom pushers. The Pioneer
and Shadow 200 (Figure 4) both fit the weight and payload requirements but fall a little short for
the endurance and cruise at a much higher speed than is necessary. The Viking300, however, fits
all of the requirements listed in Table 1, showing that it is possible to make an aircraft in this
configuration fit the mission specifications.
Figure 5. Viking 3007
Comparator UAV Review
Table 1 compares the requirements of the comparator UAVs in order to show the differences in performance and structural characteristics. Elements of the existing designs were used
to create five conceptual designs that would fit the requirements as shown.
Table 1. Comparator UAV Review
Maximum Cruise
Airframe
Speed
Speed
(knots)
(knots)
Minimum
Endurance
(hours)
Span
(ft)
Engine
(hp)
MGTOW
(lbs)
Dry
Weight
(lbs)
Payload
(lbs)
Required
70
50
8
---
---
300
---
45
T-16XL
80
50
16
13
2.5
80
40
20
Pioneer
---
64
5
16
26
450
---
---
Shadow
118
90
4
12.8
38
327
---
50
ScanEagle
70
50
20
10
1.5
37.9
---
13.2
Viking 300
70
56
8-10
16.5
22.5
318
210
30
Orca
---
100
3.5
---
---
1430
Blue Horizon 2
120
70
16
21.3
---
397
81.6
Initial Virginia Tech Overall Conceptual Designs
In order to begin the design process, ten concepts were created; one by each aerospace
engineering student on the team. All concepts are shown in Appendix B and a general discussion
of this phase of the design process follows.
19
Conceptual Analysis
A preliminary performance calculation was done on each concept by using a series of
equations found in Introduction to Aerodynamics & Aircraft Performance.8 First, a preliminary
wing area was found using the weight and speed set by the requirements (Equation 1, Appendix
A) where CL and V are for cruise condition. The rest of the wing geometry was found using a desired aspect ratio (anywhere from 5 to 10), structural concerns, and transportability concerns.
With this geometry, a power required for cruise was calculated using Equation 2 and Equation 3,
Appendix A where CD0 and e are estimated from common values. By estimating the amount of
fuel, two endurance numbers were found. Equation 4, Appendix A is for a constant altitude flight
and Equation 5, Appendix A is for a constant velocity flight where ηp and γp are estimated from
common engines and propellers. The glide performance was evaluated using the lift to drag ratio
(Equation 6, Appendix A). After these numbers were found, the designs were then optimized using what was learned from the preliminary calculations. For all designs, the wing area, wingspan,
and overall weight were reduced. This resulted in a lighter and smaller aircraft with much of the
same endurance and performance numbers.
Max Gross Takeoff Weight (MGTOW)
The requirements given by NAVAIR stated that the MGTOW was 300 lbs. After an indepth weight and mission analysis, it was determined that the mission could be completed with a
lighter aircraft. Many of the proposed concepts were sized using a MGTOW up to 100 lbs lighter
than the proposed requirement. This was done to achieve a more transportable aircraft.
Two of the proposed designs did have a MGTOW of 300 lbs. This was to assume a
worst-case scenario. If the final aircraft weighed 300 lbs, the performance of the aircraft still met
the requirements, however, building a lighter aircraft only increases the performance.
Wing Sizing and Placement
Wing sizing for the initial concepts was done by assuming a CLcruise and computing the
required wing area. By using this as a guide the wing span and chord were chosen to optimize
between transportability and a favorable aspect ratio. The concepts vary in wing planform shape
since each design is a balance between these two constraints.
Some wings incorporated taper to achieve a more elliptical lift distribution. The problem
with adding taper is that the complex planform shape makes manufacturing the wing more diffi-
20
cult. Most of the concepts utilize a high wing for the added roll stability and to avoid the need for
dihedral. A low wing might need dihedral to achieve adequate roll stability, and this complicates
manufacturing the wing as well. This design does, however, make loading payloads from the top
of the aircraft more difficult.
Power Required and Endurance
After the wing size was chosen, the power required for normal flight and the required 200
ft/min climb rate was calculated. The power required varied between the concepts due to different wing areas and aspect ratios. Nevertheless, the power for cruise for most concepts fell between 3 and 7 horsepower. Most of the designs utilize a 10-15 hp engine to facilitate a high
climb rate at higher altitudes. Using this information and engines available in this range, a specific fuel consumption was estimated for each design. Endurance calculations showed that most of
the designs have endurance between eight and twelve hours, meeting the minimum requirement
given by NAVAIR.
Deployable Landing Gear System
Several of the proposed concepts utilized a deployable landing gear system. This system
was utilized to provide a possible camera payload with a clear field of view and reduce the overall drag of the aircraft. However, this system does reduce the reliability of the aircraft. To address this, a one-time deployable landing gear system was used. This system would not utilize
any hydraulics or pneumatics, but would be retracted manually on the ground. When the aircraft
is preparing to land, a servo will release a pin and the landing gear will deploy using gravity
and/or a spring mechanism. This system is less reliable than a fixed landing gear, but is more reliable than a actuated retractable system.
Landing Techniques
One requirement given by NAVAIR was that the aircraft have the ability to land in a 250’
x 50’ “parking lot”. The proposed concepts addressed this issue in several different ways. Some
of the concepts utilized a tail hook that would capture a cable stretched across the landing area.
This design is simple, but requires more structure to support the load on the tail hook and requires the ground crew to set up the capture system.
One concept utilized a variable pitch propeller to produce a reverse thrust on landing.
This system seems to prove adequate for stopping the vehicle and it does not require additional
21
structure. In spite of this, the mechanism needed to implement this idea is complicated and possibly unreliable.
Another concept utilizes a constant braking system. The problem with a conventional
braking system is that it requires a mechanism to actuate the brakes. In the constant braking system, the brakes are already closed when the aircraft lands, making the system very simple. This
places a lot of stress on the system, however, and the landing gear and might make the aircraft
hard to control after touchdown.
The last idea proposed to meet the landing requirement was a parachute system. The parachute would be activated over the intended landing area and the aircraft would glide down. The
parachute system would call for a very small landing area, but brings in the concern of packing
the parachute before a flight. Thus, the parachute was deemed a good backup system, but not to
be used for the primary landing method.
General Geometry
Since all the proposed concepts were inspired by comparator UAVs, many of the concepts were similar. The ten proposed concepts could be placed into four groups: conventional
tractors, twin-tail booms, pylon-mounted engines, and tail-mounted pushers.
Conventional Design
Two of the proposed concepts were conventional
designs (Concept A4 and B3 in Appendix B). Both designs incorporated a single piece fuselage, conventional
tail, and single tractor engine. This design has several
benefits. Using a tractor engine reduces the noise generated by the propulsion system and the engine can be
placed near the CG of the aircraft. This allows the payload weight to be changed without the concern of shifting the CG of the aircraft. This also protects the expensive payload from a nose first crash but leaves the less expensive engine to be
damaged in a crash. The simple fuselage structure will make manufacturing the aircraft easier.
The use of a tractor engine does have the problem with possible contamination of the
sensors by exhaust. In addition the safety of the ground crew decreases with this design, because
unlike the two tail-boom designs the propeller is not protected from the crew. Both convention
22
designs also use a tricycle gear for improved ground handling and less structural loads in the tail
section of the aircraft.
One of the conventional design concepts also utilized winglets. These were incorporated
to increase the spanwise efficiency of the wing and also provide some thrust from the wingtip
vortices. However, the incorporation of winglets increases the amount of structure needed in the
wing.
Twin Tail-Boom
Most of the proposed concepts (Concepts A1,
A3, A5, B2, and B5 in Appendix B) have a twin tailboom design with a pusher engine located near the center of the aircraft between the two booms. The twin tailboom design has the benefit of protecting the propeller
from foreign objects. Also, the pusher design should
address the issue with exhaust interfering with the sensor payloads, but could be louder than a
tractor system. One main concern with this design is the more complex tail structure being a
weak point of the aircraft. This design protects the engine for a crash; however, the payload
might need to be placed farther forward to balance the aircraft and thus be susceptible to damage
in a crash.
Three of the concepts of this design incorporate an inverted V-tail. This tail was chosen
due to the lightweight structure and the ease of integration with the twin tail-boom design. The
V-Tail also has a slightly lower interference drag than a conventional tail. The other two concepts use H-Tails: one with the horizontal between the tail-booms, the other with the horizontal
on the top of the verticals. This design allows for multiple rudders increasing reliability. The
structure could be heavier than a conventional or inverted V-tail, especially in the case where the
horizontal is mounted to the vertical tails.
Pylon Mount
One proposed concept is a pylon-mounted engine design (Concept A2 in Appendix B). In
this design, the engine is mounted above the fuselage in a shroud. This design is beneficial because of the noise reduction in having the fuselage and wing between the engine and the ground.
This design also protects the propeller from debris better than any of the proposed concepts.
23
Since the engine is separate of the fuselage, volume is made available that would have otherwise
been used by the engine. This design might have thrust line issues from having the propeller far
away from the vertical CG and the structure for
this design is more complicated, thus harder to
manufacture.
For this design, a T-tail was chosen to
avoid blanketing the horizontal tail by the wing
wake in high angle of attack situations. The T-tail
is, however, a more complicated structure than a
conventional tail.
Tail-Mounted Pusher
Two of the proposed concepts have a conventional fuselage with the engine mounted in
the tail section of the aircraft. This design prevents the sensors from being affected by the engine
exhaust and places the engine far away from any electrical equipment, reducing damage or interference from vibration. This design could be significantly louder than other designs because of
increased aerodynamic interference with the pusher propeller at the tail. The propeller is also
placed in a position that could easily be damaged on landing. Placing the engine in the tail also
increases the structure needed in the tail section of the aircraft and makes balancing the aircraft
more difficult.
Both of the designs of this type utilized Y-tails. The “V” part of the tail was chosen to reduce the weight and drag of the tail. The vertical part of the Y-tail was incorporated because it
protects the propeller from a ground strike and adds the ability to have a redundant rudder for
yaw control.
24
Concept Comparison
Most of the concepts were comparable in size, differing only slightly in wingspan and
weight. As a result, the power required for cruise was roughly the same for each of the aircraft at
an average of about 3.5 hp. Other physical characteristics and performance of each concept are
compared in Table 2.
Table 2. Concept Comparison
Wing
Span (ft)
Wing
Area
(ft^2)
MGTOW
(lbs)
Empty
Weight
(lbs)
Endurance
(hr)
(L/D)max
Concept A1
20
55
300
230
10-14
16.03
Concept A2
20.6
85
250
228
8.55
13.29
Concept A3
19.3
53.6
300
225
9-11
15.7
Concept A4
12.7
20.2
132
67
8.5
11.5
Concept A5
20
65
205
160
8-10
18
Concept B1
20
74
175
140
10
10.63
Concept B2
15
45
175
150
8-9
10.23
Concept B3
16
50
200
150
16.7 - 17.7
10.36
Concept B4
15
45
175
150
8-9
10.23
Concept B5
16
40
200
160
9-10
11.6
Decision Matrix
To narrow down ten conceptual designs between the two groups into two concepts, each
group created a decision matrix consisting of key aspects to take into account that each particular
group deemed important for design. Both teams used a scale of 1 to 5 to score each concept,
with 1 being the worst, and 5 being the best. Each person within the teams then rated each concept, including their own, based on the categories within each decision matrix. The process for
both teams is outlined within the following sections.
25
Group A Decision Matrix
To analyze the five conceptual designs from Group A, a decision matrix was created consisting of key aspects to take into account (see a summarized version in Table 3 below). There
are eight main categories:

wing

performance

tail

payload

propulsion

fuselage

landing gear

overall design
The tail category includes both a division for the vertical stabilizer and a horizontal stabilizer. The overall category is a general category designed to identify areas that are not associated
with a particular aspect of the vehicle, such as storage and portability. Each main category has
been divided into several sub-categories such as: reliability, structural implications, or ease of
manufacture.
In order to determine appropriate “scores” for the individual conceptual designs, several
steps were taken to assure a fair score. All sub-category weights add up to be 100 under each
main category. After each sub-category had been rated, the score was multiplied by the respective weight to get a weighted score. These were then subtotaled for each main category and then
multiplied by the category weight. After each subtotal was multiplied by the category weight,
these scores were totaled for a total concept design score. The design with the highest score was
deemed to be the most appropriate choice to continue with. A table showing a comparison of
basic aircraft performance and parameters for each individual concept (Appendix C) was used to
help make the ratings within the decision matrix. As can be seen, concept design A1 “won” the
decision matrix. However, the group decided to merge some of the best components from each
concept to come up with a new conceptual design to create an optimal aircraft.
26
Wing
Table 3. Summarized Decision Matrix – Group A
Concept Designs
concept
concept
A3 score
A4 score
concept
A1 score
concept
A2 score
285.0
255.0
276.0
291.0
284.0
0.175
49.9
44.6
48.3
50.9
49.7
SUBTOTAL
330.0
289.5
313.0
350.0
304.0
16.5
14.5
15.7
17.5
15.2
336.0
277.0
330.0
347.0
308.0
0.05
16.8
13.9
16.5
17.4
15.4
SUBTOTAL
299.0
234.0
257.0
227.0
256.0
29.9
23.4
25.7
22.7
25.6
255.0
239.5
261.0
280.0
271.0
51.0
47.9
52.2
56.0
54.2
286.0
254.0
258.0
233.0
269.0
28.6
25.4
25.8
23.3
26.9
356.0
309.5
343.0
356.0
338.0
0.175
62.3
54.2
60.0
62.3
59.2
SUBTOTAL
330.0
300.0
380.0
260.0
360.0
0.05
16.5
15.0
19.0
13.0
18.0
SUBTOTAL
326.0
277.5
308.0
356.0
321.5
32.6
27.8
30.8
35.6
32.2
304.1
266.6
294.0
298.7
296.3
SUBTOTAL
CATEGORY
WEIGHT
concept
A5 score
Tail
Vertical Stabilizer
CATEGORY
WEIGHT
0.05
Horizontal Stabilizer
SUBTOTAL
Overall
Payload
Performance
Landing
Gear
Propulsion
Fuselage
CATEGORY
WEIGHT
CATEGORY
WEIGHT
0.1
SUBTOTAL
CATEGORY
WEIGHT
0.2
SUBTOTAL
CATEGORY
WEIGHT
0.1
SUBTOTAL
CATEGORY
WEIGHT
CATEGORY
WEIGHT
CATEGORY
WEIGHT
Total Score
0.1
27
Group B Decision Matrix
A decision matrix (Appendix D) was used to narrow the initial five concepts from Group
B into one final concept. The categories of the decision matrix were as follows:

wing

conformity to requirements

tail

reliability

fuselage

human factors

propulsion

overall aircraft
This allowed for the option of not only one concept design to be selected, but a combination of the concepts depending on the scoring in the categories. These categories were weighted
out of 1.0. Reliability was weighted the heaviest, at 0.3, because it is the main focus of this design project. Conformity to requirements was the next highest weight to make sure that the concepts fit all the requirements of the mission.
Listed in each category are the scored and weighted components. Most of these are typical or self-explanatory concerns for each aircraft component. For instance, the wing and tail both
have aerodynamic efficiency, impact on stability/control, ease of manufacturing (also seen in fuselage) and integration with fuselage. All categories, except reliability and human factor, have
some component that deals with weight. Some of the components specific to reliability require
further explanation. The glide component in the reliability category is a replacement for the reliability of the propulsion system. This was done because none of the concepts have a known propulsion system at this time. The propulsion system, however, is one of the main concerns for
reliability since it is the most likely place for failure. To overcome this hurdle, the glide characteristics were analyzed to see how well the concept could glide to safety should the propulsion
system fail. The control surface (malfunction) component was used to evaluate how the system
would respond to a control actuator failure. These servos are another point of failure. In the
overall aircraft category, the transportability component is used to describe the ease of moving
the proposed concepts on the ground. The components in each section were weighted out of a
total of 1.0 based on their importance.
A summary of the final decision matrix is shown in Table 4. This matrix was examined
by the team to make sure there were no arguments and a final concept was selected.
28
Table 4. Summarized Final Decision Matrix – Group B
Wing
SUBTOTAL
CATEGORY
WEIGHT
0.1
concept
B2 score
concept
B5 score
3.5585
3.7
3.5145
3.6815
3.397
0.35585
0.37
0.35145
0.36815
0.3397
3.2675
3.9215
3.957
3.2745
3.7005
0.32675
0.39215
0.3957
0.32745
0.37005
3.312
3.7715
3.5175
3.4925
3.5445
0.3312
0.37715
0.35175
0.34925
0.35445
2.662
3.6425
4.261
3.3125
3.3745
0.2662
0.36425
0.4261
0.33125
0.33745
3.7625
3.505
3.7685
3.4495
3.4345
0.564375
0.3505
0.37685
0.34495
0.34345
3.899
3.491
3.824
3.59
3.504
1.1697
1.0473
1.1472
1.077
1.0512
3.865
4.515
4.325
3.955
4.805
0.19325
0.22575
0.21625
0.19775
0.24025
3.517
3.825
3.619
3.649
3.6065
0.3517
0.3825
0.3619
0.3649
0.36065
3.559025
3.5096
3.6272
3.3607
3.3972
Tail
SUBTOTAL
Concept Designs
concept
concept
B3 score
B4 score
concept
B1 score
Overall Aircraft Human Factor
Reliability
Conformity to
Requirements
Propulsion
Fuselage
CATEGORY
WEIGHT
0.1
SUBTOTAL
CATEGORY
WEIGHT
0.1
SUBTOTAL
CATEGORY
WEIGHT
0.1
SUBTOTAL
CATEGORY
WEIGHT
0.15
SUBTOTAL
CATEGORY
WEIGHT
0.3
SUBTOTAL
CATEGORY
WEIGHT
0.05
SUBTOTAL
CATEGORY
WEIGHT
TOTAL SCORE
0.1
29
Reliability
According to the Department of Defense UAS Roadmap9, reliability is the “core of …
reducing acquisition system cost and improving mission effectiveness for [UAV’s].” The document goes on to state that reliability underlies the “affordability, availability, and acceptance” of
UAVs. In terms of affordability, an unmanned vehicle should be less expensive to operate and
maintain than a vehicle which is manned. By eliminating the cockpit, the average savings in
terms of weight ranges from 3,000 to 5,000 lbs. However, any further means to reduce the costs
and improve affordability tend to have a negative impact on the reliability of the aircraft. Another aspect to carefully consider is the availability of the aircraft. By including redundant systems
in the vehicle, the reliability tends to increase, as does the cost.
Clearly, reliability plays an essential role in designing an aircraft. By considering dependability issues from the beginning stages of the design process, costs to correct faults can be
reduced. Important aspects to consider include the performance of the vehicle, the payload, and
propulsion methods. Uncontrollable conditions such as weather related problems, for example
icing or high winds, also pose a threat to the overall dependability of the system.
As can be seen in Figure 6, the number one source of failure in military unmanned aircraft is related to the power and propulsion systems. Flight control, communication problems,
and human error are also listed as sources for system failures.
Figure 6. Average Sources of System Failures For U.S. Military UA Fleet (Based on 194,000 hours) 9
During the design process careful attention must be paid to the maintenance procedures
for the aircraft. It would be beneficial to minimize the number of tools required for maintenance
and make sure these functions can be accessible from the ground (versus on a lift.) Also, the material used for construction should be able to withstand corrosion. It is essential to create a design
30
which is “Jack-proof,” or simple enough that the average person would be able to understand and
use it.
With respect to overall aircraft structure, reliability considerations were taken into account in each conceptual design. Aspects of each aircraft were designed based on vehicles already in use. Taking the characteristics and reliability issues of past aircraft in account, aircraft
structure and form were designed to the desired specifications of each student.
The task of achieving the desired reliability of this aircraft is not an easy one. Aside from
specific concerns stated in each conceptual design, the considerations that must be taken into account consist of the following:

engine reliability

navigation system (including autopilot and/or GPS)

servos for flight control

communication between the ground controllers and the aircraft

structural integrity of the aircraft
Once all of these considerations are accounted for, it must be realized that ultimately, the
reliability of the aircraft will be measured by the conformity to the proposed requirements of the
design. Therefore, the performance of the aircraft will be the ultimate determinant of its reliability.
There are many tools and statistical models available for use in this project. For example,
“time until failure” distributions are historically a great choice to use as models, especially for
electrical components. The general form of this distribution function can be seen in Equation 16,
Appendix A. This function analyzes the probability P that the working condition of the component is less-than or equal-to time t.10 This is a very basic, fundamental summary of the distribution function that can be modeled into a specific continuous distribution. From this summarized
function, other functions such as failure rate can be constructed given the same data. To find this
distribution for a component, a sample of data must be collected and analyzed to find the distribution that best fits the data. The main concern with this type of modeling at this point is data
gathering.
One way that data can be gathered is from the production company of each proposed
component of the aircraft. However, this is very unlikely to be obtainable considering the confi31
dentiality and probable lack of knowledge of the manufacturers of the component. A more tangible but tedious task would be to experiment with the actual components enough to obtain a reasonable sample of data to consider it for real results.
However practical it may seem to collect component data, this is an unlikely approach
considering construction of the aircraft will not occur this year. A more feasible approach is to
model a current, similar aircraft, or components based on historical events. This will allow the
team to obtain a “ballpark” measurement of the conceptual design’s components.
Final Group Concepts
Group A Final Concept
After the decision matrix results, the group came up with a design that would combine
the best of all the individual concepts and meet the design objectives, as seen in Figure 9. The
design consists of:

high wing

H-tail

shrouded pusher propeller

twin tail-booms

tricycle landing gear
The shroud around the propeller serves two purposes: increasing the efficiency of the
propeller and significantly reducing the noise level to meet the noise requirement. Placing the
motor at the aft of the fuselage decreases the likelihood of engine damage in a crash situation.
Likewise, exhaust from a rear-mounted engine will be deposited into the airflow downstream of
the payload sensors, thus avoiding any form of contamination on the payload. The engine of
choice for this aircraft is the Desert Aircraft DA-150, outputting 16.5 hp at maximum RPM.
This engine provides more power than necessary to meet the design speeds, cruising altitudes,
and climb rate, thus improving the safety margin for flight operations. The engine fuel consumption allows for an endurance of over 10 hours with only 25 lbs of fuel.
The payload will be located in the middle of the fuselage to protect it in a crash. Keeping
the most expensive parts of the aircraft away from the nose increases the reliability by avoiding
32
unnecessary maintenance and replacement costs. The payload sensors will be mounted on a generic removable cartridge, allowing for easy exchange of payload for a variety of missions.
The high wing configuration was chosen to improve the roll stability of the aircraft, while
increasing the visibility of the payload. As mentioned above, the wing area was calculated to be
66 ft2 from constraint and performance equations to provide optimum performance at slow cruise
speeds. Likewise, an aspect ratio of about 7 was chosen to improve the wing’s ability to glide
without power and decrease its induced drag. This value was chosen such that the wing’s dimensions would support the stresses during a pneumatic launch, obtain the best performances
during flight, and provide enough volume for the fuel needed. The wing was tapered to create
more of an elliptic lift distribution. A Clark Y airfoil was chosen for ease of construction due to
its flat lower surface, which also reduces the complexities of mounting it to the fuselage. The
Clark Y airfoil provides good performance characteristics in addition to its simplistic design.
The dual-boom mounted tail allows the mounting of the horizontal and vertical stabilizers
with an aft-mounted propulsion system. The two booms allow for an H-tail configuration,
providing increased reliability from multiple control surfaces. Additionally, the H-tail provides
adequate propeller clearance during ground operation.
A non-retractable tricycle landing gear system was chosen to decrease the risk of having
a mechanical malfunction. The main gear is placed behind the payload for increased sensor visibility and propeller clearance. They also provide a wide enough base to reduce the risk of wingtip strike during landing and rollout. The main gear is mounted to the fuselage to avoid placing
additional stresses and vibrations on the wing during ground operations. For landing and stopping purposes, the main gear will utilize a constant applied breaking system. At touchdown, the
wheels will apply breaking forces for the aircraft to eliminate a separate break control system.
This increases reliability by reducing the number of systems.
For ease of ground transportation, the aircraft was designed to break apart and fit on the
roof of a truck. The wing breaks apart into three sections and fits together by sliding the wing
spars of adjacent sections into sleeves and then locking them with pins. The assembled wing
then locks into place on top of the fuselage. The dual booms and tail section slide into sleeves on
the aft side of the wing and lock in place with pins. This breakdown allows for any one damaged
component to be replaced without having to replace or repair the entire aircraft.
33
Constraint Analysis
Since maximum gross takeoff weight (MGTOW) is specified as 300 lbs in the requirements, the necessary wing area can be determined (Equation 1, Appendix A). To move forward
with the constraint analysis an Oswald efficiency factor of 0.9 and aspect ratio of 7 were assumed. The values above are constant for all phases of flight except n,
𝑑ℎ
𝑑𝑡
𝑑𝑉
, 𝑎𝑛𝑑 𝑑𝑡 will vary ac-
cording to the stage of the mission. For Figure 7, landing, stall, climb, and turn rate are plotted.
Straight and level flight is omitted from these plots because
𝑇
𝑊
and
𝑊
𝑆
required for straight and
level flight will always be less demanding than those of other conditions.
Figure 7. Constraint Analysis
The optimum design point was chosen as the point that would allow the lowest possible
wing area and thrust required. The wing loading at this design point is 4.475
this demands a wing area of about 66 ft2. The
𝑇
𝑊
33 lbs at MGTOW.
34
𝑙𝑏
𝑓𝑡 2
. At 300 lbs,
is about 0.11, resulting in a thrust required of
Sizing
Wing
The aspect ratio for the design was chosen through careful analysis of the decision matrix
results from the five individual concepts. As the winning concept had an aspect ratio of 7.27, it
was deemed that with a MGTOW of 300 lbs, an aspect ratio of 7 would generate the best performance results. This aspect ratio was verified through an iterative process between aircraft
performance and sizing. This fixed the wing area at 66 ft2, the span at 22 ft, and the mean average chord at 3 ft. To create a more elliptic lift distribution, the wing also had a taper ratio of. 0.8
Fuselage
Using statistical equations for fuselage length developed by Raymer based on MGTOW,
the length can be calculated with Equation 15 from Appendix A. The variables A and C are coefficients and Wo is the takeoff gross weight. Using a sailplane for estimation, A and C are estimated to be 0.71 and 0.48 respectively, giving a fuselage length of 11 ft. To improve the stability, the fuselage was lengthened by 1 ft to give the needed moment arm length.
Tail
To calculate the tail size, the moment arm and tail coefficients (cVT and cHT ), must be estimated. Using typical values quoted by Raymer for a sailplane, cVT and cHT are approximated as
0.50 and 0.03 respectively. The moment arm can be estimated at the conceptual design phase as
a percentage of the fuselage length. With the engine configuration of a pusher prop mounted at
the end of the fuselage, the tail arm is about 60% of the fuselage length, giving a moment arm of
6.6 ft.11 The tail aspect ratio was chosen to be two thirds of the wing’s, giving an aspect ratio of
4.67.12 The results of the initial sizing calculations are shown in Table 5.
Table 5. Initial Empennage Sizing Parameters
Horizontal Tail
Vertical Tail
Area
15.0 ft.²
6.6 ft.²
Span
8.37 ft.
5.55 ft.
Mean Chord
1.79 ft.
1.19 ft.
Volume Coefficient
0.50
0.03
35
Performance Analysis
Stall
The normal operating cruise speed of the vehicle is 50 kts at an altitude of 3000 ft. Compared to other UAVs in the same weight category, this cruise speed is slow. One of the design
goals with this vehicle is to prevent the stall speed from occurring near the normal operating
cruise speed. By placing a large enough gap between the cruise speed and the stall speed, the
safety and reliability of the aircraft can be increased.
From the constraint analysis and sizing of the aircraft in the previous sections, the planform area of the vehicle is 66 ft2 when operating at 300 lbs. Considering altitudes of 0, 3,000,
and 10,000 ft above sea-level (ASL), the plane stalls at the following speeds:
Table 6. Stall Speed versus Altitude
Altitude above Sea-Level [ft]
Stall Speed (True airspeed) [kts]
0
32
3,000
34
10,000
37
At a cruising altitude of 3,000 ft, the buffer velocity is roughly 16 kts. The worst case
scenario for stall speed occurs at 10,000 ft but still provides a 13 kts buffer velocity.
Power Required – Straight and Level
Using Equation 2 from Appendix A, the power required to maintain straight and level flight
is found and can be plotted for different altitudes. In these calculations, it is assumed the aircraft
is operating at MGTOW; the altitudes range from sea-level to 10,000 ft. The following plot depicts the power required curves for the different altitudes.
The plot reveals that a cruise speed of 50 kts would require slightly less than 3 hp to
maintain straight and level flight. The power required at cruise does not change much with altitude; this is because the cruise speed is also the speed at which minimum drag occurs. A cruise
speed of 50 kts is thus optimal for maximum endurance.
Power Required – Climb
The power required to maintain a 200 ft/min climb rate can be calculated using Equation
11 from Appendix A, assuming constant velocity. The power required to climb at 50 kts at any
36
altitude is roughly 4.7 hp, while to climb at 70 kts at sea-level is roughly 7.4 hp (Figure 8). The
power required at climb is the maximum power required to climb condition in the aircraft’s normal operating region.
Engine – Power Available
From the power required data, an engine can be chosen that meets the power requirements. The assumed drag coefficient, 𝐶𝐷0 = 0.02, may be lower than the actual value, so it is safer to choose an engine that outputs more power than calculated. In this case, the maximum power needed comes from the maximum power required for climb at sea-level, 7.4 hp. One ideal
engine for this aircraft is the Desert Aircraft DA-150, outputting a total of 16.5 hp at 8,500 RPM.
At 6,000 RPM, the engine uses 3.3 oz/min of fuel.
Extra power can be used to exceed the performance requirements, such as increasing
maximum speed or ceiling altitude. Plotting the power available (Equation 8, Appendix A), on
top of the power required curves results in the Figure 8. It is important to point out that these
calculations assume the aircraft is operating at MGTOW for all altitude.
increasing
altitude
Power
Available
stall speed at 10000 ft
Climb
S&L
increasing
altitude
Figure 8. Power Available Plot for Various Altitudes
37
Clearly the power available at any point in flight (Figure 8) will be more than needed, offering a little margin of power for safety. This plot shows that that ceiling altitude will be greater
than 10,000 ft at MGTOW.
Endurance
The endurance requirement for the aircraft is at least 8 hr. For the design cruise speed,
the aircraft is operating at minimum drag conditions. Assuming a propeller efficiency of  P =
0.85, a specific fuel consumption of  P = 3.6058e-007 1/ft, and a power output of 13 hp and
constant cruise, 19 lbs of fuel is needed to cruise at 3,000 ft for 8 hr (Equation 13, Appendix A).
Table 7 shows the endurance obtained by increasing the amount of fuel available for a
normal cruise.
Table 7. Endurance at Cruise (3,000 ft, 50 kts)
Fuel [lbs]
Endurance [hr]
19
8.1
25
10.8
36
13.6
This table shows that for each additional gallon of fuel burned, the aircraft can stay aloft
for an additional 2.7 hr.
Glide Range
The glide range of the aircraft is determined using Equation 12 from Appendix A and is a
function of maximum lift-to-drag ratio as well as absolute altitude. Calculating the best glide
speed, as well as the glide range for various altitudes yields the following table:
Table 8. Glide Range and Glide Speed versus Absolute Altitude
Altitude [ft]
Glide Speed [kts]
Range [nm]
3,000
48
7.9
5,000
49
13.2
7,000
50.7
18.5
10,000
53
26.5
38
In the event of an engine failure, this table shows that the best glide speed to maintain is
roughly 50 kts for all altitudes. Calculating the range associated with each altitude shows that
around 6,000 ft above sea level, the aircraft would be able to glide the full 15 n.m. required operational range. If an engine failure were to occur, the vehicle can glide a fair distance and, in
some cases, return to the point of departure. This gliding ability is one of the main reasons the
aircraft was designed with a higher aspect ratio.
Turn Rate
The required turn rate for this vehicle is 6°/sec. Using Equation 14 from Appendix A, the
load factor acting on the aircraft with this rate of turn is 1.04. The bank angle associated with
this rate of turn is 15.3°. With such a small increase in load factor for turning, the stall speed of
the aircraft increases no more than 1 kts.
Figure 9. New Conceptual Design
39
Figure 10. Exploded View of New Concept
Group B Final Concept
Since the final scores of each concept appeared to be so close, the group decided it would
be best to combine the components of each proposed concept. Concept B3 (Appendix B, Concept B3) started out with a slightly higher score, so further evaluation was necessary to figure out
why this was the best. Most of the components were similar for each of the conceptual designs,
except for the placement of the propeller and tail configuration. Concept B3 did not seem to
score much higher than any other conceptual designs, except in the category of acoustics. Concept B3 had a tractor propeller, so it had the potential for much quieter propulsion compared to
the pusher configuration used in all of the other concepts. After the decision to use a tractor propeller, the rest of the concept was created.
In using a pusher propeller, the fuselage and tail of many of the concepts were unconventional. However, since the final concept uses a tractor propeller, a conventional fuselage and
wing configuration was used, much like the one depicted in Figure 11. Instead of using the exact
fuselage of Concept B3, the fuselage was made more streamlined in order to produce more fa-
40
vorable aerodynamic qualities. Winglets were also eliminated from the original conceptual design in order to reduce the structural loading on the wings.
Unlike any of the previous concepts introduced, an H-tail is used on this final concept for
transportability and added reliability. In order to make the vehicle more transportable, multiple
vertical stabilizers may be used in order to help in decreasing the necessary height of the tail.
Another major concern for the conventional tail is that if the rudder fails by getting stuck in one
position or fails to activate, the yaw control of the vehicle would essentially be eliminated. By
adding a second vertical stabilizer, the second rudder should be able to provide a small amount
of yaw control, should this occur. Figure 11 shows an isometric view of the final concept.
Figure 11. Final Concept – Group B
Constraint Analysis
After a concept was chosen a constraint analysis was performed on the aircraft. Using
Equation 17, Appendix A the power loading required for several cases was found and plotted vs.
wing loading. The constraint analysis graph is shown in Figure 12.
41
Figure 12. Constraint Analysis Curves for Final Concept
It was then determined that the proposed wing loading of 4 is close to ideal. The power
loading of 0.075 is more than required and will provide enough power for higher drag flight regimes, if needed.
Airfoil Selection
The goal of the airfoil selection was to choose one that would provide 200 lb. of lift with
the selected area of 50 ft2 and would minimize the total drag. Eight airfoils (Table 9) were compared using Martin Hepperle’s JavaFoil program13 at a calculated Reynolds number of 2,000,000
to determine the needed information for the elimination process.
Table 9. Airfoil Comparison
Airfoil
Clmax
Cl0α
(L/D)max
Cm¼ c max
SD 7062
1.371
0.354
25.327
-0.101
NACA 4412
1.184
0.368
26.362
-0.118
NACA 4415
1.344
0.382
27.187
-0.128
SD7034
1.093
0.295
28.357
-0.083
S 2027
1.045
0.23
25.678
-0.077
NACA 4418
1.515
0.397
26.032
-0.142
Eppler 68
1.096
0.389
28.589
-0.137
NACA 2412
1.008
0.184
28.217
-0.065
42
The first phase of airfoil eliminations was based solely on the maximum lift coefficient of
the airfoil and a Clmax of 1.15 was determined to be the lower limit; thus the SD 7034, S 2027,
Eppler 68, and NACA 2412 were eliminated. The next phase of eliminations consisted of selecting the three airfoils with the highest maximum lift coefficients thus the NACA 4412 was eliminated. With only three airfoils left, the NACA 4415, SD 7062 and NACA 4418, several more
factors were introduced into the airfoil elimination process. The final factors for the airfoil selection criterion were, in no particular order; maximum lift coefficient, lift coefficient at zero angle
of attack, manufacturability, maximum coefficient of moment about the quarter chord and maximum lift over drag ratio.
It is important to look at the manufacturability of the airfoil section because imperfections in manufacturing could reduce the overall reliability. Features of an airfoil that could decrease its manufacturability score are excessive camber, thickness and hard to cut angles. The
maximum lift over drag ratio is a factor that provides a decent overall look at the performance of
the airfoil, a higher maximum lift over drag ratio is desirable. After setting these criteria each
airfoil was given a score and the results compared, see Table 10.
Table 10. Results of Airfoil Ratings
Weighting
NACA 4415
SD 7062
NACA 4418
Clmax
0.35
4
3
5
Cl0α
0.2
4
3
4
(L/D)max
0.1
5
3
4
Manufacturability
0.25
4
4
2
Total Score
1
4
3.35
3.65
The NACA 4415 was chosen as the airfoil to be used for the aircraft concept. This airfoil
will be used for the initial concept tail sizing and other factors to be determined later as constraint analysis progresses.
Tail Sizing
The objective of the preliminary tail sizing was to obtain a rough estimate of the needed tail
size. The tail surfaces were initially sized using Raymer’s equations8 (Equation 18 and Equation
19, Appendix A where Svt is the area of both verticals). With the wingspan and mean chord al43
ready constrained at 16 ft. and 3 ft., respectively, only the volume coefficients, Cht and Cvt, and
the moment arm lengths, Lht and Lvt, needed to be defined. A recommendation in Raymer indicated a horizontal volume coefficient of 0.7 is common; however 0.6 was used to account for the
endplate effect from the H-tail. A vertical tail volume coefficient of 0.04 was chosen based on an
average value for comparator aircraft found in Raymer.
This gave a horizontal tail area of 13.85 ft2. The tail aspect ratio was also selected as two
thirds that of the wing based on Stinton’s12 recommendation, effectively increasing the stall angle
of the horizontal tail above that of the wing. Using a 7° leading edge sweep on the horizontal the
span, root chord and tip chord were found. Using these numbers and the assumption that the vertical tails are mounted to the tips of the horizontal with a chord equal to the tip chord of the horizontal the dimensions of the verticals were found. A NACA 0012 was picked for both the horizontal and vertical tails to allow volume for the required structure. The results of the initial sizing
calculations are shown in Table 11.
Table 11. Initial Empennage Sizing Parameters
Horizontal Tail
Vertical Tail (for one tail)
Area
13.85 ft.²
2.35 ft.²
Span
7 ft.
1.5 ft.
Root Chord
2.22 ft.
1.8 ft.
Mean Chord
2.01 ft.
1.6 ft.
Tip Chord
1.8 ft.
1.4 ft.
Moment Arm
6.5 ft.
6.8 ft.
Taper Angle
7°
15°
Volume Coefficient
0.6
0.04
Engine Selection, Power Requirements, and Endurance
The proposed concept design will cruise at an average speed of 56 knots with a power required of 4 hp at cruise (Figure 13 and Figure 14). The proposed concept can easily achieve the
required 70 knot maximum speed. Figure 13 shows how the cruise speed will change as the mission progresses while Figure 14 shows the power required for the aircraft to stay airborne.
44
Figure 13. Cruise Speed and Stall Speed vs. Time
Figure 14. Power Required Curve with 15hp Available
The engine selection process was much like the airfoil selection process where several engines
were directly compared and a decision matrix was produced in order to make a selection. Three
engines were compared, two from Lightning Aircraft Inc and one from Desert Aircraft14,15,
The relevant engine specifications are provided below in Table 12.
Table 12. Engine Specifications for Engine Comparison
Lightning Aircraft Inc Lightning Aircraft Inc
150D2-B
250D2
Desert Aircraft
DA – 150
Output
15+ Hp
22 Hp
16.5 Hp
Weight
9.25 lbs
12.5 lbs
7.96 lbs
Fuel Consumption
0.8 lb/hp-hr
0.68 lb/hp-hr
0.5625 lb/hp-hr
Other Design Aspects
UAV
UAV
Acrobatics
45
16
.
The main problem with the Desert Aircraft DA-150 is the fact that it was designed specifically for competition aerobatics. This means that this engine is specialized and highly tuned.
The high tuned and high performance nature of this engine put its long term endurance reliability
in question. Furthermore, this engine does not have a rear output shaft.
The Lightning Aircraft 150D2-B and 250D2 are both engines that were designed to be
used on UAVs. Both engines have rear output shafts in order to attach starters, alternators or any
other accessories that may be required.
Table 13. Engine Selection Decision Matrix
Weight
DA-150
150D2-B
250D2
Output
.2
4
4
5
Weight
.3
4
3
2
Fuel Consumption
.25
4
3
2
Output Shaft
.25
1
5
5
Total Score
1
3.25
3.7
3.35
As seen in Table 13, the Lightning Aircraft Inc 150D2-B engine was selected for this
conceptual design. This engine has a good balance of weight, output power and fuel consumption as well as having the ability to attach various engine accessories to the rear output shaft.
Based on this engine selection the endurance of the proposed concept was determined.
This concept aircraft was proposed with an initial fuel capacity of 50 lbs. The endurance of the
aircraft with 50 lbs of fuel is around 16 hours. This is double the requirement which could allow
for an increase in payload weight if less endurance is required.
Starters and Alternators
In order to increase the reliability of the propulsion package on the aircraft it was decided
that a starter is a necessary engine component. Having a starter will allow the engine to be restarted in mid flight, as well as allowing the ground crew to start the engine from a safe distance.
One of the requirements for this concept was to be able to provide 10 watts of power at
12 volts during the duration of the mission. This is difficult and heavy to do with batteries alone.
Thus it was decided that an alternator was needed to charge a smaller set of batteries in order to
run the payload electronics on the aircraft.
46
The alternator recommended for use on this concept is a Sullivan Face Type alternator. 17
This alternator is simple and contains few moving parts and is thus reliable and easy to replace or
repair if a failure does occur. This alternator can be mounted on the rear output shaft or on the
front output shaft behind the propeller. See Reference 16 for an overview of the Sullivan alternator. The starter18 that is recommended for this concept is also from Sullivan. They manufacture several different styles of starters, a rear mount and an under mount style. They also have a
combination alternator/starter available. A starter style will be picked when it is decided how the
motor will be mounted in the aircraft and how much available clearance is around the motor.
The Sullivan alternators and starters can be mounted on any engine; however they may require
custom mounting brackets.
Stability
After a tail size was chosen the stability of the system was evaluated using Athena Vortex
Lattice (AVL).19 A simple AVL model was made and was used to find the neutral point of the
aircraft. Figure 15 shows the geometry plot of the AVL model. For the aircraft to be stable a static margin of 5-15% percent was desired.
Figure 15. AVL Geometry Plot
47
The results from AVL show that the neutral point is 2.06 ft. back from the leading edge
of the wing. This is equal to 0.69c, which is a normal value for stable aircraft in AVL. The value
is high; however the fuselage will cause a destabilizing effect.
Qualitative Reliability Analysis
The final concept was analyzed by the whole team in terms of reliability. One of the major changes to the design was the iteration of the tail. The H-tail form was chosen for this aircraft because it provided two rudders that could allow the plane to function and fly if one were to
fail, most likely a servo failure. Instead of a broad nose, the team also chose a sleek, tapering
nose integrated with the propeller to provide less drag from the fuselage. This will enable a
longer glide distance if the engine were to fail.
Once a final concept is chosen, the engine type, electrical components, and navigation
system will be chosen based on historical data of mean time until failure and failure rates. This
will give the aircraft the highest possible percentage of combined reliability for the entire aircraft
system.
48
Final Values
Table 14. Final Values for Group B Concept
Fuselage
Value
Performance Data
Value
Length (ft)
11.5
CL max
1.344
Width (ft)
1.8
CL cruise
0.4
Height (ft)
1.8
(L/D)max
10.36
Wing
Value
Endurance (Const alt) (hr)
16.23
Airfoil
NACA 4415
Endurance (Const V) (hr)
17.67
Span (ft)
16
Maximum Climb Rate (ft/min)
1500
Area (ft2)
50
Stall Speed (knots)
30
Chord at Root (ft)
3.5
Maximum Speed (knots)
100
Chord at Tip (ft)
2.5
Cruise Speed (knots)
58
Aspect Ratio
5.12
Weight Statement
Value
Horizontal Stabilizer
Value
Payload Range (lb)
35-45
Airfoil
NACA 0012
Dry Weight (MEW) (lb)
150
Span (ft)
7
Gross Takeoff Weight (lb)
200
Chord at Root (ft)
2.22
Engine
Details
Chord at Tip (ft)
1.8
Engine
Lightning Aircraft Inc
150D2-B
Max Power (hp)
15+
SFC (lb/hp-hr)
0.8
Area
(ft2)
13.85
Volume Coefficient
0.6
Vertical Stabilizer
(values for one vertical)
Value
Airfoil
NACA 0012
Height (ft)
1.5
Chord at Root (ft)
1.8
Chord at Tip (ft)
1.4
Area (ft2)
2.35
Volume Coefficient
0.04
49
Loughborough University Design Process
Each student at Loughborough University came up with six conceptual designs. These
designs were then grouped together based on common traits. Five groups in total were needed to
group the initial concepts: conventional, multi-tail, multi-wing, multi-fuselage, and pylonmounted engine. A decision matrix was then used to rule out 3 of the initial concept groups,
multi-fuselage, multi-wing, and multi-tail. With two concepts remaining, a detailed analysis of
the conventional and pylon-mounted concepts was performed and approximate aircraft sizing
was determined (Table 15 and Table 16). Each concept was then sketched in CAD (Figure 16
and Figure 17) and dimensioned in meters.
Figure 16. LU Conventional Concept
Table 15. LU Conventional Concept Sizing
Parameter
Meters
Feet
Total Length
3
9.8
Fuselage Length
1.5
4.92
Fuselage Diameter
0.3
0.98
Wing Span
4.75
15.58
Wing Chord
0.53
1.74
Tail Span
1.08
3.54
Tail Chord
0.53
1.74
Tail Height
0.6
1.97
50
Figure 17. LU Pylon-Mounted Engine Concept
Table 16. LU Pylon-Mounted Concept Sizing
Parameter
Wing
SI
Tail
US
2
27.01 𝑓𝑡
SI
2
0.398 𝑚
US
2
Area
2.511 𝑚
Aspect Ratio
9
9
-
-
Span
4.75 𝑚
15.59 𝑓𝑡
-
-
Mean Chord
0.528 𝑚
1.73 𝑓𝑡
-
-
4.284𝑓𝑡
The advantages and disadvantages for the final two concepts were determined and listed
for future debate (Table 17). These key points were later brought up in the elimination process
with Virginia Tech. It is important to note that both concepts from the Loughborough team contained an internal combustion engine as the main propulsion system and a back up electric motor
in order to get the aircraft back to base safely. This back up engine is to be used in case of main
engine failure or to provide additional power.
51
Table 17. LU Final Concepts Advantages and Disadvantages
Conventional
Advantages
Pylon-Mounted Engine
Disadvantages
Advantages
Tractor propeller reduces risk of propeller strike
Rear pusher adds to
expense and weight
Electric back up can be
mounted near the primary
Modular lightweight tailboom
Chance of tail strike
damaging rear pusher
Avionics, payload and
propulsion at small risk
in a crash
Electric get home/ dash
facility
Airfoils experience wash
from tractor propulsion
system
Easy access for
maintenance
Possible to over-torque
engine on rail launch
Large internal fuselage
volume
Main weight contributions close to center of
gravity
Simplicity in design and
manufacture
Large knowledge base
from existing conventional aircraft
Disadvantages
Large nose down pitching moment requires
counter-action from tailplane
High mounted propulsion system may cause
instability during
taxi/landing/take off
Downselection
A total of four conceptual designs were suggested by Virginia Tech and Loughborough
University. In order to find the ideal design, three of these concepts were eliminated using a variety of methods including the listing of advantages and disadvantages and group discussions.
Key issues that arose from presenting each of the concepts were identified by the combined
group in order to determine points of discussion for each design. Of these key issues, the ones
found to have a greater effect on the overall and final configuration were complexity of design,
thrust line problems, center of gravity placement (lateral and vertical), noise, engine placement,
and propeller protection. These key issues were considered by each group member individually,
and then thoughts on how they applied to each conceptual design were brought up in a group
discussion.
The elimination process included identifying categories that fit each of the conceptual designs to reduce the number of designs to decide between. This involved identifying adaptable
components of each design. For example, wing placement was considered a trivial modification
for each concept, because each design could support a low-, mid-, or high-mounted wing. Other
adjustable features of each design included tail type, aspect ratio, landing gear configuration, and
placement of auxiliary power systems. From this grouping of conceptual design types, the two
52
suggested conventional tractor propeller configurations were combined to reduce the number of
concepts to three.
The second step of the elimination process involved writing a list of advantages and disadvantages for each of the remaining three conceptual designs as the Loughborough team had
done with their concepts (Table 17). These allowed the combined group to analyze the aforementioned key issues and how each conceptual design fared in each of the categories. For example, the twin tail-boom pusher configuration had a clear disadvantage in complexity of design,
lateral center of gravity placement, structural robustness, and noise. Because the twin tail-boom
pusher design had more parts and deviated from the conventional design, it was considered to be
a more complex design. Complex designs were considered to be less reliable, as it would require
a ground crew to have more knowledge of how to assemble the aircraft. The lateral center of
gravity placement was considered to be unstable because the removable cartridge may allow for
the CG to shift depending on how the cartridge was loaded. Furthermore, the overall balance of
the aircraft was seemingly non-existent due to the weight of the engine and the moment-arm of
the tail being compensated by only the payload in the front. From the analysis of key issues for
each concept, the twin tail-boom pusher design was eliminated because its disadvantages far
outweighed the number of advantages it provided.
After the first two steps of the elimination process, only the conventional configuration
and pylon-mounted pusher remained as viable concepts. In order to save time and to try another
method of elimination, the combined Virginia Tech and Loughborough University groups were
split into four groups of six, where the members of each group were a mix of students from each
university. Within these small groups, issues such as whether or not the problems previously
presented could be designed for as well as examples of current UAVs with similar configurations
were discussed. From these small group discussions, each group came away with a consensus
that the pylon-mounted pusher design would be the best choice between the last two concepts as
long as some modifications were made.
The overall reasoning that each group had for selecting the pylon-mounted pusher included the diversity of options that the design provided, the ease of engine cooling, the propeller protection provided by mounting the engine high, the reduction in noise, and the ease of maintenance. The diversity of options in selecting the pylon-mounted engine design included the ability to have all of the same configuration options as the conventional design as well as the ability
53
to have either a tractor or pusher propeller without worrying about the lateral center of gravity
effects. Furthermore, by mounting the engine high, the noise could be reduced by adding some
of the aircraft’s own structure as a physical barrier between the ground and the propeller, the engine could be cooled by the free stream air-flow, easier to access for maintenance would be provided, and the propeller would be well protected from ground strikes.
To make the pylon design the best concept possible, modifications were made by placing
the wing high, using a tractor engine, and using a conventional tail. The higher wing placement
provides better stability, which is necessary because the vertical CG is very high in this configuration. The use of a tractor engine provides less noise because it does not have to cut through the
wake of a pylon and back-up propeller. In this case, the electric engine would be mounted on the
aft portion of the pylon in a pusher configuration where the propeller blades could be folded
back. Finally, the use of a conventional tail as opposed to any other tail was selected because it
provides the most control upon failure of a control surface. Use of a tail boom was also suggested to reduce the amount of structure necessary in the tail.
Final Concept
Figure 18. Final Concept Drawing
From the Virginia Tech and Loughborough University concepts, a new concept was generated (see Figure 18 and Figure 21). The new concept, referred to as a pylon design due to the
placement of the engine, inherited features from the previous designs. The geometry of the concept is conventional with the exception of the engine mounted above the wing. A high wing and
54
a conventional tail design are used along with a tail-dragger landing gear, fixing a skid in the
place of the tail wheel. The skid is intended to provide added friction after landing to help stop
the aircraft. The fuselage is widest in the area under the wing to accommodate the payload, avionics, batteries, and other onboard devices, positioning them close to the CG. The concept utilizes an internal combustion engine mounted at the top of the pylon in a tractor configuration.
Also in the pylon is the auxiliary propulsion unit, an electric motor and folding propeller intended to give the aircraft about a half hour of power in the event of the main internal combustion
engine failing. The electric motor is in a pusher configuration, facing aft. The fuel will be stored
in the pylon, above the wing to raise the vertical position of the CG. For the wing airfoil, the
NACA 4415 was tentatively selected because it gave a good lift to drag ratio, a relatively high
maximum lift coefficient and the shape is not difficult to build.
Constraint Analysis
The relationship between thrust to weight and wing loading is shown in Equation 21 in
Appendix A. The following assumptions in Table 18 were made:
Table 18. Assumptions for Constraint Analysis
Aspect Ratio, AR
6.25
Stall Lift Coefficient, C L , max
1.3
Drag Coefficient,
CD0
Oswald Efficiency Factor, e
0.02
0.9
The constraint analysis in Figure 12 illustrates the thrust to weight ratio and wing loading
for the design specifications including a cruise speed of 50 kts, dash speed of 70 kts, climb rate
of 200 ft/min, and a turn rate of 6 degrees/sec.
55
1.2
T/W Climb
1
T/W Straight
Strait and
Level
T/W
and
Level
0.8
T/W 70 kt Dash
T/W Min Turn
0.6
Landing
T/W
0.4
Stall
Design Point
Design
Point
0.2
0
0
1
2
3
4
W/S [lb/ft^2]
5
6
7
8
Figure 19. Constraint Analysis for New Conceptual Design
𝑙𝑏
The design point selected results in a wing loading of 3.1 𝑓𝑡 2 and a thrust to weight of
0.14. Assuming an aircraft weight of 200 lbs, the required wing area is 64 ft2, and the thrust required is 28 lbs.
Wing Sizing
The first step to wing sizing was determining a CLcruise, which was set to 0.4. This yielded a
wing loading of 3.125 lb/ft2 (150 Pa) using Equation 20, Appendix A. After validating that this
wing loading would work with the constraint analysis, the wing area was found assuming a
MGTOW 200 lb (91 kg). From there the span was varied until it was small enough to fit in the
10 ft. box limit that was imposed and had a high enough aspect ratio. The final wing numbers are
shown in Table 19.
Table 19. Wing Sizing Parameters
Area
64 ft² (5.95 m2)
Span
20 ft. (6.10 m.)
Chord
3.2 ft. (0.98 m.)
Aspect Ratio
6.25
Wing Loading
3.125 lb/ft2 (150 pa.)
56
Tail Sizing
After the wing size was chosen, the tails were sized using Raymer’s equations11
(Equation 18 and Equation 19, Appendix A). The moment arm for both the horizontal and vertical was set to eight feet (2.44 m.) and the volume coefficients, Cht and Cvt, were found from recommendations in Raymer. This gave a horizontal tail area of 16.66 ft2 (1.55 m2). The tail aspect
ratio was also decided to be two thirds that of the wing based on Stinton’s12 recommendation,
effectively increasing the stall angle of the horizontal tail above that of the wing. Using a 10°
leading edge sweep on the horizontal the span, root chord and tip chord were found. The vertical
area was found to be 6.00 ft2 (0.55 m2). The preliminary height, root chord and tip chord were
found by trial and error. Table 20 shows the final tail numbers.
Table 20. Initial Empennage Sizing Parameters
Horizontal Tail
Vertical Tail
Area
16.66 ft² (1.55 m2)
6.00 ft² (0.55 m2)
Span
8.35 ft. (2.55 m.)
3.0 ft. (0.91 m.)
Root Chord
2.36 ft. (0.72 m.)
2.5 ft. (0.76 m.)
Mean Chord
2.00 ft. (0.61 m.)
2.0 ft. (0.61 m.)
Tip Chord
1.63 ft. (0.50 m.)
1.5 ft. (0.46 m.)
Moment Arm
8.00 ft. (2.44 m.)
8.0 ft. (2.44 m.)
Taper Angle
10°
18°
Volume Coefficient
0.65
0.04
Performance Analysis
The performance of this concept can be broken down into a number of categories: power
required for straight and level flight, power required for climb, stall speeds for straight and level
flight and turns, endurance, and glide range. Each of these categories is analyzed using the assumptions in Table 21.
.
57
Table 21. Assumptions for Performance Analysis
Planform Area, S
64 ft2
Wing Span, b
20 ft
Average Chord, c
3.2 ft
Aspect Ratio, AR
6.20
Maximum Weight
200 lbs
Cruise Lift Coefficient, C L , cruise
0.4
Stall Lift Coefficient, C L , max
1.3
Drag Coefficient,
CD0
0.038
Oswald Efficiency Factor, e
0.9
Climb Rate, dh dt
200 ft/min
Specific Fuel Consumption,
P
3.605769 x 10-7
Additionally, the aircraft is assumed to operate at altitudes of sea-level, 3,000 ft, 5,000 ft,
and 10,00ft. For the purpose of analysis, the normal cruising altitude will be interpreted as
5,000 ft mean sea-level. Using the preceding assumptions, the power required for flight can be
estimated.
The power required to maintain straight and level flight can be calculated using Equation
2, Appendix A. Similarly, using Equation 11, Appendix A the power required for climbing
flight can be found. In each case, the power required is found as a function of velocity and plotted in the following graph. For a cruise altitude and speed of 5,000 ft and 50 kts respectively,
the power required for straight and level flight is 3.5 hp. Likewise, the powered required to
climb at 200 ft/min at maximum weight is 4.7 hp. Straight and level flight at 5,000 ft and 70 kts
maximum cruise speed requires 8.0 hp.
58
Increasing altitude
Pavail
Preq Climb
Preq Straight
and Level
Increasing altitude
Figure 20. Power Curves
In Figure 20, estimated power available lines have been drawn to provide a general idea
of the flight performance that is desired. The three horizontal lines on the top half of the plot
represent the power available for flight at sea-level, 5,000 ft, and 10,000 ft, with sea-level being
the uppermost line. Assuming an engine with a maximum power output of 10 hp at sea-level,
the aircraft is able to reach the maximum cruising speed of 70 kts at around 5,000 ft altitude.
The aircraft is easily able to climb to 10,000 ft at cruise speed and at full weight. In fact, the aircraft can exceed the 200 ft/min climb rate at all altitudes up to 10,000 ft and all speeds up to 60
kts.
Using Equation 9 from Appendix A, the stall speed was calculated for all three altitudes
for straight and level flight as well as turning flight at MGTOW (Table 22). For turning flight
defined in the requirements, a load factor n = 1.1 is used in Equation 9, Appendix A.
59
Table 22. Stall Speed at Various Altitudes
Straight and Level
Altitude
Stall Speed n = 1.0
Turning
Stall Speed n = 1.1
0 ft
27 kts
28 kts
5,000 ft
29 kts
30 kts
10,000 ft
31 kts
33 kts
Regardless of altitude, the stall speed of the aircraft is well away from the normal cruise
speed of 50 kts. Keeping a fairly large gap between stall speed and cruise speed improves the
safety of the aircraft in the advent the aircraft encounters gusty winds during flight.
Equation 13 from Appendix A is used to calculate the endurance of the aircraft based on
constant speed flight and a fixed amount of available fuel. Using an iterative method, the
amount fuel required to obtain an endurance of 8 hours at 5,000 ft is roughly 22 lbs. With this
amount of fuel, a range of 485 n.m. is possible at these flight conditions. Since the design calls
for cruise at constant speed and constant altitude, an additional few pounds of fuel can make up
for the difference in flight profile.
In the event that the aircraft losses engine power, the probability for a safe landing is improved dramatically with an increasing glide range. Equation 12 from Appendix A roughly estimates the glide range of an aircraft given a maximum lift-to-drag ratio and an altitude. The
maximum lift to drag ratio for this aircraft is L D max = 10.78. At 5,000 ft mean sea level, the
aircraft can glide about 8.8 n.m. At 10,000 ft the glide range increases to 17.7 n.m. In the latter
case, the aircraft has a good chance of returning to base after an engine failure even while operating at the maximum design range of 15 n.m.
60
Figure 21. Final Concept 3-View
61
Preliminary Design Phase
After selecting a final conceptual design, further analysis was performed in order to determine modifications that needed to be made. Areas in which the design process required further analysis included: Aerodynamics, Structure, Stability and Control, Propulsion, Weights and
Balances, CAD, Performance, Systems Integration, Ground Control, Reliability, and Costs.
Each of these areas was assigned a primary lead on both the American and British teams so that
both sides could work in conjunction. The assignments for both teams are shown in Table 23.
Table 23. Team Responsibilities
Area
Virginia Tech
Loughborough University
Team Lead
Amanda Chou
Dan Marshall
Aerodynamics
Anthony Ricciardi
Ben Hanson
Structure
Richard Duelley
Robert Penn
Stability and Control
Philip Pesce
Andrew Courtneidge
Weights and Balances
Mike Sherman, Alex Kovacic
Alex Humphrey
CAD
Alex Kovacic, Mike Sherman
Alex Humphrey
Propulsion
Dennis Preus
Dan Jones
Performance
Megan Prince
Peter Christie
Systems Integration
Belle Bredehoft
Robert Noble
Ground Control
Robert Briggs
Craig Dillon
Reliability
Erik Sunday
Bal Chand
Costs
Jessica McNeilus
Kris Hanna
In the preliminary design phase, communication was vital between the two teams as well
as between each of the specified disciplines. As a result, an organizational team website with file
uploading capabilities was used in order to keep track of changes in configuration. For communication between smaller groups, instant messaging software such as MSN Messenger and Skype
were used.
Aircraft Overview
The aircraft was designed with the specified requirements in mind, but also with an emphasis on reliability. Once a primary method of performing to those requirements was found, a
secondary redundant system was often considered. The key dimensions of the aircraft are listed
62
below in Table 24. The reasoning behind most of these dimensions is such that the plane can be
broken down into parts and assembled in the field.
Table 24. Key Dimensions of the Vulture
Key Dimensions
Wing Span
21.75 ft
6.6 m
Wing Area
75.6 ft2
7.1 m2
Aspect Ratio
6.25
Aircraft Length
17.3 ft
5.3 m
Overall Height
4.25 ft
1.3 m
Horizontal Tail Area
14 ft2
1.3 m2
Vertical Tail Area
4.5 ft2
0.42 m2
Two configurations of the aircraft were considered: one with an electric back-up propulsion system (IC/Electric Configuration) and one with an internal combustion (IC) back-up propulsion system (IC/IC Configuration).
Both configurations weigh in under the required
MTOGW, therefore, the amount of fuel or payload weight can be increased. The heavier
weighted option is shown below in Table 25 to show the distribution of weight on the aircraft.
Table 25. Mass Breakdown of Vulture (IC/Electric Configuration)
Mass Breakdown
Max Take Off
277 lb
125.6 kg
Operational Empty
191 lb
86.6 kg
Fuel
40 lb
18 kg
Payload
45 lb
20.4 kg
The IC/IC Configuration of Vulture weighs in at 245 lb, which is about 32 lb less than the
IC/Electric configuration. Although both configurations weigh significantly less than the required MTOGW of 300 lb, the aircraft was designed in order to be able to carry such a weight
and cruise at 50 knots. This indicates that a heavier payload or more fuel can be added to the
system in order to meet the MTOGW requirement.
Requirements Met
Both configurations of Vulture were designed with a 300 lb MTOGW, 50 knot cruise
speed, 45 lb payload, and operational altitude of 3000 ft AGL in mind. Therefore, these re63
quirements have been met. A comparison of the specified requirements and performance of the
aircraft are broken down in Table 26.
Table 26. Requirements and Performance Comparison
Required
Predicted for Vulture
Endurance
8 hours
13.9 hours
Range
15 n.m.
793.3 n.m.
Climb Rate
(Sea Level)
200 ft/min
Turn Rate
6°/sec
Service Ceiling
10,000 ft
Landing Distance
250 ft
202.3 ft
Noise Output
< 50 dB(A)
50-60 dB(A)
700 ft/min
(at 70 knots)
8°/sec
(at 50 knots)
900 ft/min climb at 10,000 ft
possible
Structural Design
The design of the structure was intended to make the aircraft as lightweight as possible to
aid in the endurance and range of the vehicle. As a result, many of the elements of the vehicle
are designed to be made with foam core, with major structural elements designed to be made
from aluminum.
Rectangular Fuselage
A rounded square shaped fuselage was chosen to house the internal components of the
aircraft in order to better budget the use of space inside the fuselage. Most of the components
are rectangular in shape and fit more easily inside. The rounded edges of the fuselage are intended to simplify construction and aid with the aerodynamic characteristics of the fuselage.
Keel Design
The keel is made from an extruded aluminum bar and is designed to handle all expected
flight loads with a factor of safety of 1.6. All major elements of the aircraft are attached to or
hung from this major element of the structure. This also means that a flight termination system
may be attached to the keel, as it is the main structural component of the aircraft.
64
Materials
Wing and tail surfaces are made of foam core with a fiberglass skin, which significantly
reduces the weight of the vehicle. The wing foam core structure is supported by an aluminum Ibeam main spar and an aluminum C-section rear spar. The tail foam core structure is supported
by carbon fiber tubes. The fuselage skin will be made up of removable acrylic panels with rubber seals around the edges. This will help ensure that the internal components of the aircraft
cannot be damaged by water.
The tail boom is made of a four-inch carbon fiber tube, which will be slid into frames
mounted on the main keel beam and pinned in place. In order to prevent point loads on the carbon tail boom, a wooden plug will be attached with epoxy into the end of the carbon tube. The
pins that are used to hold the tail boom in place will pass through the carbon tube and the wooden plug, thus distributing the loads from the pins throughout the tail boom.
Modularity of Design
The design is made to be broken down into pieces that can fit within a 6 ft by 6 ft by 12 ft
trailer. Each piece will connect to other portions of the aircraft using pins and plugs. When broken down, the aircraft consists of three major components: the fuselage, the wings, and the tail
assembly.
The Fuselage Section
The main keel beam is a square extruded aluminum beam. The square keel beam simplifies component
integration by providing a surface that is easy to mount
to and carries all expected flight loads.
The pylon
(Figure 22) is integrated into the wing/fuselage joint to
optimize load transfer and is made from the same material as the main keel beam. A simple aerodynamic fair-
Figure 22. Pylon Mounting Structure
ing will wrap around the aluminum box section in order to reduce the drag of the pylon mount.
Machined rear frames connect the tail assembly to the main keel beam using pins.
65
The fuselage (Figure 23) also consists of a replaceable foam nose cone that will dissipate some impact forces if
the aircraft noses over during landing. The skin panels of
the fuselage consist of non-structural acrylic panels that
mount to the fuselage using pins. These panels are easy to
Figure 23. Fuselage with No Skin
remove and will allow for easy access to the entire payload
and avionics systems. They will also have rubber seals around the edges to prevent water from
seeping into the fuselage. A small inboard section of the wing is also integrated with the fuselage. This section of the wing includes portions of the main and rear spar. These box sections
are the mounting points for the outboard wing sections.
The Wing Section
Figure 24. Detached Wing Section
The wing section (Figure 24) is made up of a simple foam core construction. The aluminum I-beam main spar and C-section rear spar slot into the box sections mounted on the fuselage
and pin into place. The wings also consist of replaceable wing tips consisting entirely of foam
with no structural supports within. In the event of damage while landing, these wingtips can be
removed and replaced.
The Tail Assembly
The tail assembly consists of the carbon fiber tail
boom and tail surfaces. In order to minimize any point
loads on the tail boom, the carbon tube is fitted with
wooden plugs at either end. The plug on the fuselage end
is mounted internally and the plug on the tail surface end
is mounted externally to provide mounting points for all
Figure 25. Tail Assembly
tail surfaces, as pictured in Figure 25. The tail surfaces are
of a foam core construction with a fiberglass skin and carbon tubes as reinforcement.
66
Stress Calculations
The major areas of concern include the pylon mount, the wings, the tail, and the main fuselage keel beam. Situations in which these areas would experience extreme loadings include
stresses due to rail launch, due to landing, and due to gusts.
Table 27. Stress Calculations
Main Fuselage
Keel Beam
Fuselage Wing Box
Sections
Wing Main Spar
I-Beam
Max Stress (MPa)
159.5
172.1
159.8
Yield Strength (MPa)
276.0
276.0
276.0
Factor of Safety
1.73
1.6
1.73
As shown in Table 27, the factor of safety in loading on each of the sections of interest never
drops below 1.6 and averages at about 1.7.
Weights and Balances
Weights and balances for both configurations of aircraft were determined by creating a
three-dimensional model of each component and assembling them together. These components
were either assigned a known weight or a known density in the CAD drawing. Unigraphics NX4
and Autodesk Inventor are both capable of calculating the moments of inertia and center of gravity (CG) locations of components drawn in these programs as long as a material, density, or
weight is assigned to them. Instead of solely relying on this program, however, a spreadsheet
summarizing each part drawn using location, density, and weight was composed to double-check
the values. The moments of inertia of each of these pieces and weights were used to find the
overall aircraft moment of inertia and CG location.
Moments of Inertia
For the IC/IC configuration, the moments of inertia in the stability axis frame of reference are as follows:
67
Table 28. Moments of Inertia for the IC/IC Configuration
Ixx (slug-ft2)
56.33066
Iyy (slug-ft2)
85.27103
2
Izz (slug-ft )
135.0672
Ixy (slug-ft2)
4.70×10-4
Ixz (slug-ft2)
-2.41203
2
8.08×10-4
Iyz (slug-ft )
For the IC/Electric configuration, the moments of inertia in the stability axis frame of reference are:
Table 29. Moments of Inertia for the IC/Electric Configuration
Ixx (slug-ft2)
56.50191
Iyy (slug-ft2)
93.30774
Izz (slug-ft2)
142.9307
2
Ixy (slug-ft )
5.27×10-5
Ixz (slug-ft2)
-3.79849
Iyz (slug-ft2)
8.26×10-5
These values are required for dimensionalizing stability derivatives and for understanding
the ease or difficulty in changing the rotational motion of plane about the stability axes.
Center of Gravity Location
Neutral Point
Center of Gravity
Figure 26. CG and Neutral Point Locations
The center of gravity (CG) was determined by analyzing the weight of each of the components on board the aircraft as well as the structure of the aircraft. The point at which the mo-
68
ments of the weights of each of the component summed up to be zero was the point at which the
CG was found to be located.
Vertical (from centerline of keel):
Horizontal (from the first bulkhead):
Neutral Point (from the first bulkhead):
-0.97 in.
55.44 in.
63.89 in.
Size Comparison
In order to understand the size of the aircraft, a six-foot tall man is usually depicted with
the aircraft. In the picture below, a six-foot tall man next to the side view drawing has been provided (Figure 27). A three-view drawing with dimensions is available in Appendix I.
Figure 27. Six-Foot Man with Vulture
As compared to typical aircraft of the same weight, such as the Pioneer or the Shadow, the Vulture is much larger due to the need for such slow cruise speeds.
Aerodynamics
The responsibilities of the Aerodynamics leads on each side were to select an airfoil, perform a drag build-up, find stability derivatives, and to analyze the overall aerodynamic quality of
the aircraft. Programs used to perform these tasks included Fluent, AVL19, JavaProp13, and Tornado.
Overall Aerodynamics
The area of the wing was selected in order to be able to carry the MTOGW at cruise and
dash speeds. Using an aspect ratio of 6.25 to keep the wing within small enough dimensions to
place within a 6 ft by 6 ft by 15 ft trailer, the final dimensions of the wing were:
69
Table 30. Wing Dimensions
Aspect Ratio
6.25
Chord [ft (m)]
3.48 (1.061)
Span [ft (m)]
21.75 (6.63)
Planform Area [ft2 (m2)]
75.69 (7.03)
When a paneled model of the wing was placed in AVL, the Oswald efficiency of the
wing was found to be 0.96 and the 𝐶𝐿𝛼 is 4.8/rad. Other three-dimensional effects of the wing
include that the 𝐶𝐿𝑚𝑎𝑥 is reduced to 1.5 and and the 𝐶𝐿0 is -4.7 degrees with the wing at an incidence angle of 2°.
Airfoil Selection
Desired wing characteristics, such as a high 𝐶𝐿𝑚𝑎𝑥 and high (𝐿/𝐷)𝑚𝑎𝑥 were considered
when selecting an airfoil for the wing. A range of airfoils such as NACA, SD and Eppler airfoils
were considered for these traits, and the comparison of a sample of these airfoils is shown below
in Table 31.
Table 31. Airfoil Selection Characteristics
Airfoil
Clmax
Clα
(L/D)max
Cm¼ c max
SD 7062
1.691
0.094
27.192
-0.106
NACA 4412
1.485
0.089
34.397
-0.122
NACA 4415
1.684
0.096
28.373
-0.133
SD7034
1.369
0.089
36.96
-0.086
S 2027
1.314
0.095
33.666
-0.081
NACA 4418
1.895
0.084
24.536
-0.147
Eppler 68
1.324
0.092
28.589
-0.137
NACA 2412
1.294
0.094
36.242
-0.069
Other surfaces requiring an airfoil, such as the tail or pylon mount, used symmetric airfoils which were chosen based on structural or stability needs.
Drag Buildup
The values for drag build-up (Table 32) are based on the method suggested by Raymer
with additional material form Tornenbeek20. The method is based on the skin friction of a flat
70
plate. To account for surface irregularities and surface roughness, the boundary later is assumed
to be fully turbulent. In this case, the skin friction coefficient C F, is given by the PrandtlSchlichting formula corrected for Mach number and associated equations (Equation 22, Appendix A). The 𝐶𝐷0 of each component are summed to arrive at the parasite drag of the aircraft:
Table 32. Drag Build-Up with Raymer and Tornenbeek's Approximations
Horizontal
Vertical
Engine
Wing
Fuselage
Stabilizer
Stabilizer
Nacelle
Span or
Diameter
6.96
2.64
1.3
0.52
0.22
(m)
Length
1.11
0.425
0.425
5.2
0.85
Scale (m)
Reynolds
1.87×106
7.15×105
7.15×105
8.74×106 1.43×106
Number
Landing Gear
𝐶𝐷𝑤ℎ𝑒𝑒𝑙
𝐶𝐷𝑤ℎ𝑒𝑒𝑙
0.25
𝑆𝑤𝑒𝑡
𝑆𝑟𝑒𝑓
6.45×104
t/c
0.15
0.15
0.15
Pylon
x/c
0.2
0.2
0.2
𝐶𝐷𝑝𝑦𝑙𝑜𝑛
1.40
𝑆𝑤𝑒𝑡
𝐶𝐷𝑝𝑦𝑙𝑜𝑛
𝑆𝑟𝑒𝑓
2.71×10-3
Cf
0.00399
0.00476
0.00476
0.00307
0.00418
Q
1.1
1.1
1.1
1
1.3
F
1.303
1.303
1.303
1.085
2.050
Swet
14.5
1.88
1.11
8.71
0.625
Cd
1.07×10-2
1.65×10-3
9.72×10-4
3.74×10-3
8.99×10-4
Total Cd
Profile
7.35×10-3
2.53×10-2
Another method used to find the drag build up is a program written by W.H. Mason from
Virginia Tech,21 which models the skin friction and form drag. This method validated the previous method and provided a way of rapidly achieving results for the various mission altitudes as
well as the cruise and dash speeds. The results for these flight conditions are listed in Table 33.
Table 33. Drag Estimations from Mason's Program
Mach Number
Altitude (ft)
CD,FRICTION
CD,FORM
CD,FRICTION+CD,FORM
0.07
0
0.01085
0.01594
0.02679
0.1
0
0.01017
0.01492
0.02508
0.07
3000
0.01102
0.01618
0.0272
0.1
3000
0.01032
0.01515
0.02546
0.07
10000
0.01144
0.01681
0.02825
0.1
10000
0.0107
0.01571
0.02642
71
Prop Wash Effects
Because the thrust line of the vehicle is above the CG, the control surfaces will have to
work to trim the aircraft every time the throttle setting changes. JavaProp was used to determine
whether or not placing the horizontal stabilizer directly in line with the thrust line will solve this
problem.
JavaProp’s flow field capability was used to estimate the projection of the propeller’s
wake onto the tail surface. A change in thrust would increase the wake velocity causing an increased flow rate over part of the horizontal tail. The flow increase would effectively produce
more “pitch-up” moment from the tail, which would counteract the “pitch-down” moment
caused by an increase in thrust. The resulting output from JavaProp shows placing the horizontal
stabilizer in line with the thrust line will reduce the moment caused by changing throttle.
Table 34. Prop Wash Effects on the Tail
Moment From
Thrust
Thrust
(lbs)
MThrust
(ft-lbs)
Moment
From Tail
MTail
(ft-lbs)
𝒅𝑴𝑻𝒉𝒓𝒖𝒔𝒕
𝒅𝑻
𝒅𝑴𝑻𝒂𝒊𝒍
𝒅𝑻
189.9
282.3
292.6
1.487
0.8485
30.8
45.79
157.6
1.487
0.9924
2.27
3.375
129.3
1.487
1.016
Placing the horizontal stabilizer in line with the thrust line, however, causes other problems, as
shown in Figure 28. According to Raymer, these angles illustrate the blanketing effect of the
wing on the tail. If the horizontal stabilizer is placed in this zone, it has almost no effect.
Figure 28. Angles Showing Blanketed Regions (Raymer)
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Tail Sizing
Using Raymer’s equations for initial tail surface sizing (Equation 18 and Equation 19,
Appendix A), the horizontal and vertical stabilizer sizes were found to be 16 ft2 and 6 ft2 respectively. When drawn in the CAD model, the proportions of these surfaces appeared slightly larger
than desirable, so another method of finding the tail surfaces was used. Because a moment balance method had already been used to find the CG, the same method was used to balance the
aerodynamic forces of the aircraft, assuming stick-fixed control. From the required tail moment,
a planform area for the horizontal stabilizer was determined to be 14 ft2. The aspect ratio of the
horizontal stabilizer, as suggested by Stinton, was dimensioned to be two-thirds that of the wing.
However, this balancing method was not suitable for determining the size of the vertical tail, and
therefore, another method was needed.
The vertical stabilizer plays an important role in directional stability, so the ability for the
aircraft to handle a 10 knot crosswind became the deciding factor for the stabilizer size. A simple script was used to analyze control surface deflections needed for two crosswind landing cases: for a given crosswind component and for maximum allowable crosswind components. Using
this method of sizing the vertical stabilizer, the area was reduced to 4.48 ft2. By reducing the
size of the stabilizer, less of a control surface deflection was needed to achieve the 10 knot
crosswind landing requirement. A summary of the key dimensions of both the horizontal and
vertical stabilizers is given in Table 35.
Table 35. Summary of Tail Dimensions
Surface
Average Chord
Span
Area
Taper Ratio
Horizontal Stabilizer
1.85 ft
7.72 ft
14.3 ft2
0.7
Vertical Stabilizer
1.82 ft
2.46 ft
4.48 ft2
0.7
Control Surfaces
Control surfaces were sized in order to control the aircraft in basic maneuvers such as a
2-g turn, a 2-g pull-up, and a 10-knot crosswind landing. Further considerations were made into
allowing for the aircraft to maintain control authority after the failure of a surface with any remaining functioning surfaces. Upon the failure of an elevator, aileron, or rudder, the aircraft will
have sufficient command authority to return for a safe landing.
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Ailerons
Ailerons were sized using Raymer’s suggestions of a percent span and percent chord.
Specifically, the ailerons should take up to roughly 40% of the wingspan and be positioned as far
outboard of the wing as possible. Historic data then showed that for ailerons spanning 40% of
the wing, the average aileron chord should be roughly 22% that of the average wing chord. With
this, the ailerons were designed with a rectangular planform area for simplicity.
To account for servo or control surface failure, each aileron on each side of the wing was
split into two equally sized surfaces for four surfaces in total. A single servo properly sized for
the aerodynamic forces controls each surface. By implementing four servos individually attached to each of the four aileron surfaces, the hard-over failure of any one control surface can be
counteracted by the remaining three surfaces.
The deflection range of each of the ailerons is taken to be ±15°. In the event of a 10-knot
crosswind landing, assuming the landing speed is 35 knots, the ailerons need only to deflect 2°
when coupled with a 12° rudder deflection.
Elevator
The elevators were sized using Raymer’s suggestion of using a percent horizontal stabilizer chord. Specifically, the elevator chord should be roughly 45% that of the horizontal stabilizer chord. For simplicity, the elevator surface is rectangular and its average chord is 45% the
average chord of the tapered horizontal stabilizer. This allows for the surface hinge to remain
perpendicular to the longitudinal plane. Similar to the ailerons, the elevator is split into two independently driven surfaces. The failure of one servo or surface can be compensated for by the
remaining functioning surface.
The elevators were sized in order to be able to perform a 2-g turn and a 2-g pull-up. For
a 2-g turn, the elevators must deflect 21°. For the 2-g pull-up, they must only deflect 18°. The
deflection range of the elevators was assumed to be ±25° based upon experience. However, for
the given turn rate requirement of 6°/sec, the elevator only needs to deflect a fraction of this
range.
Rudder
The rudder was sized using Raymer’s suggestion of a percent chord, similar to that of the
elevator. Specifically, the average rudder chord should be roughly 40% that of the average stabi-
74
lizer chord. Once again, the surface has a rectangular planform area and the hinge remains perpendicular to the longitudinal axis.
The rudder remains as a single servo-driven surface. The basis for this decision is that
the aircraft can fly in a constant sideslip in the event a servo fails in the deflected position. The
remaining surfaces can be used to control the aircraft safely to landing. Having either of the other surfaces fail in a deflected position is far more detrimental to flight than that of the rudder.
The use of a single rudder surface also reduces the amount of weight created by the additional
servo and added structure.
Another consideration that was made included the ability of the rudder to maintain authority in a spin. Using Raymer’s blanketing angles, as shown in Figure 29, the position of the
vertical stabilizer was originally too far aft for at least a third of the rudder to be unblanketed by
the horizontal stabilizer. Therefore, the vertical tail shown below was moved forward in order to
aid in spin recovery and control.
Figure 29. Raymer's Angles for Spin Recovery
Additional Control Surface Points
The control surface configurations are not capable of recovery in the event of multiple
surface failures, but splitting the surfaces reduces the possibility of loss of control authority. In
the event that an elevator fails in its maximum deflected position, the remaining surface will
have to compensate for this undesired pitching moment by deflecting in the complete opposite
direction. Maxing out the elevator surfaces cancels out undesired pitching moments from the
stuck elevator, yet it also prevents the aircraft from being able to maneuver in the longitudinal
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axis. Such a situation is less likely to occur with the ailerons as they are split into twice as many
surfaces as the elevator.
Depending on the limitations of the installed autopilot system and the ability to reprogram control laws within that system, surfaces could be used to control the airplane about axes
other than their primary intention. In the event of total aileron failure or lack of aileron control
authority, for example, the elevators can deflect in opposite directions to create a rolling moment. Ailerons could also be deflected symmetrically to control pitch and adjacent ailerons
could be deflected asymmetrically to create yawing moments. The throttle can also be used to
pitch the aircraft at the expense of altitude and speed, due to the fact that the engine is mounted
high above the center of gravity. The difficulty in this method of control is closely linked to the
autopilot’s ability to identify failed surfaces, but is possible with future development and alternative hardware.
Stability Analysis
Stability derivatives are used to determine the static and dynamic stability of the aircraft.
In order to determine whether or not dihedral is needed in the wing, for example, stability coefficients concerning roll and yawing motions are necessary. Using Athena Vortex Lattice (AVL),
the following stability derivatives were found for both a dihedral wing and a straight wing with
no taper.
Stability Derivatives for a Dihedral Wing
Due to the high CG of the aircraft, a 13° dihedral to the outboard 2 feet of the wings was
added in order to make the aircraft more stable in the lateral direction. The stability derivatives
for this case from AVL are shown in the following tables:
Table 36. AVL Output for a Dihedral Wing
alpha (α)
beta (β)
Z’ force - 𝑪𝑳
4.798
0
Y force - 𝑪𝒀
0
-0.1905
X’ moment - 𝑪𝒍
0
-0.0680
Y moment - 𝑪𝒎
-0.6835
0
Z’ moment - 𝑪𝒏
0
0.3087
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Roll Rate (p)
Pitch Rate (q)
Yaw Rate (r)
Z’ force - 𝑪𝑳
0
7.638
0
Y force - 𝑪𝒀
0.2260
0
0.1283
X’ moment - 𝑪𝒍
-0.4571
0
0.1438
Y moment - 𝑪𝒎
0
-12.72
0
Z’ moment - 𝑪𝒏
-0.0365
0
-0.0658
Aileron (δa)
Elevator (δe)
Rudder (δr)
Z’ force - 𝑪𝑳
0
0.005486
0
Y force - 𝑪𝒀
0.000251
0.000278
0.001901
X’ moment - 𝑪𝒍
0.006386
-0.000291
0.000089
Y moment - 𝑪𝒎
0
-0.015115
0
Z’ moment - 𝑪𝒏
0.000296
-0.000138
-0.000894
Neutral point xnp: 1.783 ft from the leading edge of the wing
𝐶𝑙𝛽 𝐶𝑛𝑟
𝐶𝑙𝑟 𝐶𝑛𝛽
= 1.009 > 1
As can be determined from a quick glance at the signs of these stability derivatives, the
aircraft with dihedral is extremely stable. In fact, the stability of the aircraft may be so stable
that it is not easily handled because the responses to control input may be too slow. To avoid
complicating structure unnecessarily, the stability derivatives for a straight wing were also analyzed.
Stability Derivatives for a Straight Wing
The stability derivatives for a straight wing were found with AVL in order to determine
the static and dynamic responses of the aircraft due to small perturbations. If the stability of the
straight-winged configuration of the aircraft was sufficient, the final wing configuration of the
aircraft would also be straight. This was done in order to keep the structure of the aircraft simple
and to reduce weight that may be added at the joints of the dihedral sections.
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Table 37. AVL Output for a Straight Wing
alpha (α)
beta (β)
Z’ force - 𝑪𝑳
4.799
0
Y force - 𝑪𝒀
0
-0.1873
X’ moment - 𝑪𝒍
0
-0.0469
Y moment - 𝑪𝒎
-0.6617
0
Z’ moment - 𝑪𝒏
0
0.0338
Roll Rate (p)
Pitch Rate (q)
Yaw Rate (r)
Z’ force - 𝑪𝑳
0
7.691
0
Y force - 𝑪𝒀
0.1023
0
0.1066
X’ moment - 𝑪𝒍
-0.4550
0
0.1419
Y moment - 𝑪𝒎
0
-12.74
0
Z’ moment - 𝑪𝒏
-0.0354
0
-0.0655
Aileron (δa)
Elevator (δe)
Rudder (δr)
Z’ force - 𝑪𝑳
0
0.005487
0
Y force - 𝑪𝒀
-0.000247
0.000279
0.001902
X’ moment - 𝑪𝒍
0.006396
-0.000291
0.000089
Y moment - 𝑪𝒎
0
-0.015113
0
Z’ moment - 𝑪𝒏
0.000286
-0.000138
-0.000894
Neutral point xnp: 1.767 ft from the leading edge of the wing
𝐶𝑙𝛽 𝐶𝑛𝑟
= 0.6397 > 1
𝐶𝑙𝑟 𝐶𝑛𝛽
From a quick analysis of the signs of the stability derivatives, it is clear that the aircraft is
statically stable. Further analysis of this configuration is detailed in the following sections.
Static Stability
The static stability of an aircraft indicates whether or not the aircraft will make a restoring motion due to a small perturbation. This does not indicate any long-term effects of pilot input to the aircraft, but indicates whether or not the aircraft is initially stable.
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Longitudinal Stability
The longitudinal stability of an aircraft is heavily dependent on the static margin of the
aircraft. To determine the static margin of the aircraft, the position of the neutral point and center of gravity must be determined. The values for critical numbers to calculate static margin are
given in Table 38.
Table 38. Neutral Point and Static Margin Calculation
Position from First
Bulkhead (in.)
c/4
53.09
Center of Gravity
55.44
Leading Edge
42.65
Neutral Point
63.89
Static Margin
20.22%
A statically stable aircraft will have a positive static margin, but the range of acceptable
values varies according to the type of aircraft. One website comments that a 5% static margin
correlates to “twitchy” controls, 20% static margin correlates to “mushy” controls, and more than
that runs the risk of stalling the elevator at takeoff and landing22. Modern UAVs, however, tend
to have larger static margins around 28% to 38% because they require a computer response.
Also, it should be noted that because the stability derivative 𝐶𝑚𝛼 is less than zero, the aircraft is statically stable with respect to angle of attack perturbations. With respect to pitch rate,
the aircraft is also stable because 𝐶𝑚𝑞 is less than zero.
Directional Static Stability
The aircraft is directionally statically stable with respect to sideslip angle perturbations
because 𝐶𝑛𝛽 is greater than zero and with respect to yaw rate because 𝐶𝑛𝑟 is less than zero.
Lateral Static Stability
Without dihedral, the aircraft is laterally statically stable with respect to roll perturbations
because 𝐶𝑙𝑝 is less than zero. Another indicator of this is the fact that the CG is located below
the wing surface, which causes the aircraft to experience a pendulum effect, as the weight of the
aircraft itself creates a restoring moment. The dihedral effect, characterized by 𝐶𝑙𝛽 is also less
than zero.
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Dynamic Stability
The dynamic stability characteristics of the aircraft indicate how the vehicle responds to
pilot or autopilot input. Long-term responses indicate whether or not the vehicle will maneuver
in such a manner that causes it to either stall or structurally fail. It is therefore important to analyze the dynamic motions of the aircraft.
Longitudinal Motion
The longitudinal dynamic motion will indicate whether or not the aircraft will stall given
a small perturbation from pilot input. Depending on the frequency and amplitude of oscillation,
the computer may be able to recover from undesired motions. An analysis of the two modes of
longitudinal motion is described below.
Short Period Mode
The following characteristics of the dynamic response of the short period mode to longitudinal perturbations are as follows:
𝜔𝑛 = 6.369 𝑟𝑎𝑑/𝑠𝑒𝑐
𝜁 = 0.9542
𝑡ℎ𝑎𝑙𝑓 = 0.114 𝑠
Figure 30. Thumbprint Criterion for Short Period Mode Handling 23
It should be noted that the short period mode is classified as having “poor” handling qualities as defined by the “thumb print criterion” provided by pilot opinions (Figure 30). This is due
to the fact that the static margin of the aircraft is larger than that of a typical manned vehicle.
80
The natural frequency of the short period mode can be reduced by decreasing the size of the static margin. However, most UAVs on the market have higher static margins closer to 25% for increased stability in the longitudinal axis. This is, of course, at the expense of the short period
mode.
Phugoid Mode
Characteristics of the Phugoid mode are as follows:
𝜔𝑛 = 0.338
𝜁 = 0.1384
𝑡ℎ𝑎𝑙𝑓 = 14.8136 𝑠
𝜔𝑛,𝑝
= 0.053
𝜔𝑛,𝑠
Criteria for the Phugoid mode include a minimum damping ratio of 𝜁 = 0.04. It is also recommended that the ratio of the Phugoid and short period mode natural frequencies be well separated
such that
𝜔𝑛,𝑝
𝜔𝑛,𝑠
≤ 0.1. From these definitions, it is clear that this aircraft mode is stable.
Lateral Motion
Lateral motion is important to consider due to the implications that if the amplitude of
such oscillations will cause the aircraft to receive loadings higher than the vehicle can candle.
The spiral mode and Dutch roll mode are considered for these reasons. In the case that either of
these modes are unstable, however, it is important to find the time to double amplitude, as it will
indicate the amount of time that the autopilot has to respond to unstable motion.
Spiral Mode
If the value of
𝐶𝑙𝛽 𝐶𝑛𝑟
𝐶𝑙𝑟 𝐶𝑛𝛽
is greater than 1, the vehicle is spirally stable. The IC/IC configura-
tion does not have a value greater than 1, therefore, it is not spirally stable. However, because
the spiral mode has a longer period, it is not a significant concern, so long as the time to double
amplitude is greater than 10 seconds. This allows for the autopilot to have time to respond to
any unstable motion. In comparison, for level 1 quality (mission phase) and flight phase categories C and B (requiring gradual maneuvering and less precise to precise tracking), the time to
double amplitude should be between 17.3 seconds and 28.9 seconds. The Vulture has the following spiral mode characteristics:
81
𝜆𝑠𝑝𝑖𝑟𝑎𝑙 = 0.0343
𝑡𝑑𝑜𝑢𝑏𝑙𝑒 = 20.2291 𝑠
Therefore, despite the fact that the spiral mode is inherently unstable, there is enough time for a
human pilot or autopilot to correct this motion.
Dutch Roll Mode
Characteristics of the Dutch roll mode of the aircraft are as follows:
𝜔𝑛 = 2.0271
𝜁 = 0.2781
𝑡ℎ𝑎𝑙𝑓 = 1.2295 𝑠
For a variety of flight phases and quality levels, these values show that the aircraft is stable in
this mode.
Propulsion
Due to the eight-hour endurance requirement, the primary propulsion system was required to be an internal combustion piston engine. Most available aircraft engines are two-stroke
gas engines, which may or may not be manufactured with the intent of being the most reliable or
the most suitable for flight. Therefore, the option of a backup propulsion system was also added,
as propulsion failures account for 38% of all UAV failures. Two options for this backup engine
were considered. First, a back-up electric engine (IC/Electric Configuration) with only enough
endurance to return to base was devised. Second, a completely component redundant system
with two internal combustion engines (IC/IC Configuration) was proposed.
Power Required
The power required for straight and level flight, a 200 ft/min climb rate at sea level, 60°
banked turn at 50 knots and 100 ft/min climb at the service ceiling were calculated (Table 39).
The power required for each of these conditions was used to size the necessary amount of horsepower for the engine.
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Table 39. Power Required for Different Conditions
Power Required
At 50 knots
(hp)
At 70 knots
(hp)
Straight and Level Flight at 3000 ft AGL
4.0
8.1
200 ft/min Rate of Climb
6.1
10.8
60° Banked Turn at 50 knots
8.5
11.2
100 ft/min Climb at Ceiling
4.8
8.1
The Internal Combustion Engine
An emphasis on engines capable of using heavy fuels such as JP5 or diesel was made, but
commercial off-the-shelf (COTS) engines such as these could not be found for the horsepower
required. An option to convert a gasoline engine was also considered, but this reduces the overall reliability of the engine and causes to the engine to run less efficiently. The ZDZ 160 B2RV
Champion engine was selected for use on this aircraft as it is one of the first two-stroke engines
manufactured for use on aircraft.
Specific Fuel Consumption
The specific fuel consumption (SFC) for the IC engine operating in the flight range listed
above was estimated to be 0.65
𝑙𝑏
ℎ𝑝∙ℎ𝑟
𝑙𝑏
and 0.67 ℎ𝑝∙ℎ𝑟 for the IC/Electric and IC/IC Configurations
respectively. Estimates were made by first taking the manufacturer-listed fuel consumption rate
of 3.3
𝑜𝑧
𝑚𝑖𝑛
for 6000 rpm and converting to
𝑙𝑏
ℎ𝑝∙ℎ𝑟
. This process indicates that at 6000 rpm, or the
maximum power output range of the engine, the SFC is 0.77
𝑙𝑏
ℎ𝑝∙ℎ𝑟
. To find the SFC for the en-
gine operating at reduced power settings for cruise, an estimate of 85% of the internal combustion engine’s maximum power SFC was taken to be the SFC for both configurations.
A second method used to compare the estimates with existing data includes using the
SFC curve for an existing aviation engine and scaling it to match the selected power plant. To
do this, a base SFC was first estimated using (Equation 2326 and Equation 2426, Appendix A).
Engine displacement, fuel use and RPM were used to find the operating fuel-to-air ratio along
with volumetric flow rates of air and fuel according to RPM. An RPM-dependent fuel flow rate
in lb/hr was then found assuming a density of 6 lb/gal, or the density of aviation gas. This meth𝑙𝑏
od results in a constant SFC of 0.83 ℎ𝑝∙ℎ𝑟 that does not vary with RPM or power due to the nature
83
of assumptions made. After examining the results of this analysis, there was a large discrepancy
found in the initially assumed/calculated value for SFC as compared to the scaled result.
Figure 31. Society of Automotive Engineers (SAE) SFC vs. RPM Curve 24
Due to the discrepancies in SFC found in the second method, a third method had to be
used to obtain a SFC vs. RPM curve. Using a SFC vs. RPM chart supplied by Society of Automotive Engineers (SAE) (Figure 31), a fourth-order polynomial fit was applied to the data and
then combined with an approximated power curve (Appendix K) to create plots of both SFC vs.
RPM and power vs. RPM. A scaling factor was found to make the maximum value from the approximated power curve match the power output of the selected engine at 6000 rpm. This scaling factor was then applied to the power and SFC curves to obtain plots of estimates of the ZDZ
160 engine’s fuel consumption and power output.
Figure 32. Appendix K Script Output for Specific Fuel Consumption Estimates
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While the scaled power curve resembles what would be expected of an engine in this size
range, the scaled and constant calculated values for SFC found using the second method appeared much higher than expected (Figure 32). The scaled SAE fuel consumption curve data for
the lowest compression ratio available closely matched the initial assumed values of 0.65
𝑙𝑏
ℎ𝑝∙ℎ𝑟
𝑙𝑏
and 0.67 ℎ𝑝∙ℎ𝑟, so these values were used for all calculations of endurance and range. Due to the
unique nature of the engine choice for the IC/IC Configuration of this aircraft (two full-powered
engines), and lack of available data for engines in this displacement and power class, it was decided that the original estimates for SFC would be best to use for the remainder of the design
process.
Propeller Sizing
XRotor was used to design a propeller to fit the operational performance requirements
given for the UAV. The pitch distributions (Figure 33) for commercially available propellers
were compared with XRotor results to examine the feasibility of using a COTS unit in place of
the custom-designed propeller.
Figure 33. Propeller Design Plot
In order to produce enough thrust for cruise at 50 knots and a 70-knot dash, a threebladed 28-inch diameter prop with 12-inch pitch was selected. The designed blade from XRotor
is shown in Appendix L.
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Parts Required for Internal Combustion Engine
Because the fuel tank is situated below the engine, a fuel pumping system must be used.
To ensure reliable fuel flow to the engine, the fuel tank will also be pressurized with free-stream
air captured from an intake in the pylon. This will provide fuel to the engines with minimal
complexity as long as there is airflow across the intake either from forward velocity or “prop
wash.”
The exhaust system required for the IC engine is an exhaust manifold and muffler or
tuned pipe. Mufflers are significantly smaller and half the weight and price of a tuned pipe. A
tuned pipe, while longer and more difficult to place on the pylon, would provide increased engine performance but could also affect noise levels produced by the engine. Either option can be
used for either configuration.
An alternator increases the weight of the system by approximately two pounds. As a custom item, there is a higher cost associated by adding this part. This cost is offset, however, by
the ability to always have the onboard batteries in a charged state. The alternator is a brushless,
simple, small, and reliable system and can keep the onboard batteries charge using a charging
circuit and a voltage regulator.
All internal combustion engines will be ground-started, including the backup engine in
the IC/IC Configuration. Therefore, there is no need for an onboard starter for internal combustion engines.
The Electric Engine
As one of the newer technologies on the electric motor market, brushless outrunner motors have many advantages over their brushed DC counterparts. The efficiency, reliability, noise
output, lifetime and reduced electromagnetic interference of the brushless outrunner motor sported all the characteristics desired in a secondary propulsion system. The reliability of the electric
engine is often touted by amateur and professional aircraft makers and pilots as being anywhere
from five to ten times that of an internal combustion engine. This added component creates a
great debate as to whether or not the cost of such a component at $11000 – more than twice the
cost of a single internal combustion engine – is worth the extra reliability.
A Plettenberg Motors Predator 30 11kW brushless outrunner motor was selected for
analysis of the IC/Electric configuration. Since the motor is brushless there is less friction than a
brushed motor because there are no brushes contacting the coil. This will increase the life span of
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the motor. An outrunner motor is where the outside casing rotates around stationary coils and
was selected due to their good power-to-weight ratio. Also, an outrunner does not require a gearbox, further reducing the amount of complex parts.
At a weight of 3.4 pounds, the electric engine requires 14 packs of lithium polymer
(LiPoly) batteries in order to return to base. A majority of the cost of this electric system comes
from the batteries.
Analysis
Electrical system analysis is relatively new, so the methods used were validated using
wind tunnel test data gathered previously. The first step to an analytical solution is to find power
available 𝑃𝑎𝑣𝑎𝑖𝑙 , and then the amps of the system I and the total resistance Rtotal using Equation
25, Equation 26, and Equation 27 from Appendix A where PF is 3.2 for an APC propeller. From
here the thrust T can be found using Equation 28 and Equation 29 from Appendix A.
Figure 34. Thrust and Drag vs. Velocity from Analytical Method
A graph showing the thrust and drag, and thus, the envelope in which the aircraft can operate with the electric motor is shown above.
Endurance Calculation
Using the above methods it was decided the aircraft would fly close to 43 knots while
cruising with the electric motor – a slightly conservative estimate. Assuming the worst case sce-
87
nario, where the aircraft loses the primary engine at 15 n.m., then it can be found how many batteries are needed. The max weight of the aircraft dictated that 14 battery packs were the max allowed (2 in series, 7 in parallel). This would allow a return from 22.2 n.m. in ideal conditions.
One limitation on this would be a headwind on return. A headwind will lower the ground
speed of the aircraft, increasing the return time. With the chosen system, the maximum headwind
allowable would be 16 knots. Another limitation is the inability to climb. Since the system is
running at full throttle and must fly at 43 knots to return home, the aircraft will not be able to
climb if necessary.
Maximum Headwind
One of the major considerations in using electric engines are that the limited power supply may not have enough range to allow for the aircraft to return to base in the event of a main
engine failure. The maximum allowable headwind increases with the addition of battery packs,
but so do the weight and cost of the vehicle (Figure 35).
Max Allowable Headwind (knots)
60
50
40
30
20
10
0
0
50
100
150
200
Weight of Battery Packs (lbs)
Figure 35. Maximum Allowable Headwind vs. Weight of Battery Packs
In the case that the winds exceed that allowable for the number of battery packs onboard,
the aircraft will not be able to return if that wind creates a headwind on the vehicle. The probability that the aircraft will not be able to return in the case of a main engine failure is about half
the probability that the winds will be above 16 knots. This is due to the fact that half of the time,
the wind should act as a tail wind, which will aid in the return of the aircraft by increasing its
groundspeed.
88
Propeller Selection
The propeller chosen for the electric backup system is a COTS two-bladed 24-inch diameter reward-folding unit with a 12-inch pitch. It was sized to fit the electric motor’s power output and to provide enough thrust to fly at 43 knots. In addition to the COTS item, a custom designed unit was created with XRotor and shown in Appendix M.
Parts Required for an Electric Engine
Almost any electric engine will have fewer types of components to implement than an internal combustion engine. This means that the system is much simpler, and therefore has fewer
minimal cut sets and less of a chance of failure.
Lithium Polymer Batteries
The source of major cost in terms of dollar amounts and weight lies in the power supply
to the electric motor. In order to return home safely in a 16 knot headwind, 14 packs of LiPoly
batteries must be used. This brings the total cost of the batteries to $6150 and the total weight of
the batteries to 31.5 pounds.
Another concern regarding the use of LiPoly batteries is their tendency to explode when
charged incorrectly. The use of LiPoly batteries on this vehicle means the ground crew must
have sufficient training in the proper method to charge and discharge the batteries between or
before missions. Furthermore, LiPoly batteries must be exchanged every three years for safety,
because they tend to wear down and become unbalanced.
Electronic Speed Controller (ESC)
The Jeti SPIN 200 Opto Electronic Speed Controller weighs 0.6 pounds and costs $548.
The Jeti controller was selected since it is very programmable and logs all relevant data, which
should help with preliminary testing of the system. Its purpose is to receive the signals from the
autopilot or manual controller and change the voltage to the motor accordingly. In doing so,
large amounts of heat can be generated. As a result, component placement must be carefully
considered.
IC/IC Configuration
Because the price and weight of this engine is only slightly more, both of the engines in
this configuration are the full 16 horsepower ZDZ engines. This means that in the case that one
89
engine were to fail, the other engine could not only return to base, but could essentially complete
the mission. At a price of only $4000 more and a weight cost of only 6 pounds more, the added
reliability of a second internal combustion engine is well worth its cost.
In the IC/IC Configuration, both engines will be running at the same time in order to cut
down on the risk that the backup engine will not start. This implies that the specific fuel consumption will be slightly higher for this configuration than for a configuration where only one
internal combustion engine is running at a time.
IC/Electric Configuration
The placement of the internal combustion and electric engines within the pylon are illustrated below.
Figure 36. Diagram of Pylon Packaging for the IC/Electric Configuration
As can be seen, this configuration has a reduced number of components, and therefore has less
weight within the pylon. The IC/IC Configuration would have the left half of the pylon packaging as depicted in Figure 36 reflected across to the right half as well. The reduced number of
components lessens the structural loading of the pylon, and therefore, the aircraft can be made
lighter, should this configuration be used.
Noise Prediction
Key sources of aircraft noise come from airframe turbulence, sonic booms, propellers,
“jet” or turbine engines, fans and compressors, thrust reversers, combustion, and cabin noise.
Due to the configuration and purposes of the proposed aircraft, the major consideration for noise
90
will regard the propeller system. On most aircraft, the propulsive noises will overpower any other noise factor of a subsonic aircraft.25
Airframe turbulence noises mostly affect in-cabin noise, which is not a concern for an
unmanned vehicle. The far-field noise of such turbulence will be broadband and can be relatively small, especially when dissipated by the atmosphere. “Clean” airframe noises add almost imperceptible amounts of noise, but high-lift devices and landing gear increase the noise level on
landing by about 10 dB. This 10 dB occurs in higher frequencies which may be imperceptible to
human ears or masked by the propulsive noise. However, the ground crew will be very close to
the aircraft upon landing, so the need to wear earplugs at that point will be likely.
Noise Prediction Equations
Using Roskam and Lan’s book26, a far-field noise prediction can be used to determine the
noise levels a single propeller can output at a given distance away. For the proposed aircraft,
power input at 4500 RPM is assumed to be 6 hp and at 6000 RPM is assumed to be 10 hp. To
extrapolate the far field partial noise level FL1 for such a low horsepower, Equation 30 in Appendix A was used.
Because there is no requirement indicating the direction of the loudest noise, the highest
decibel level, which is ~1 dB, is used to calculate the Directivity Index. To find the NC for multi-bladed propellers, the OSPL, the PNL, and dB(A) respectively, Equation 31, Equation 32,
Equation 33, and Equation 34 from Appendix A were used. Values for FL2, FL3, and ΔPNL
were approximated from charts in Roskam and Lan’s book26. The changes in perceived noise
level ΔPNL are taken to be 3 dB and 4 dB for a tip Mach number of 0.5 and 0.6 respectively.
91
Table 40. Summary of Noise Prediction Values
Distance
FL1
FL2
RPM Blades
Away
(dB) (dB)
(ft)
4500
2
200
45.85
19
3
6000
2
3
FL3
(dB)
DI
(dB)
NC
(dB)
OSPL
(dB)
ΔPNL
(dB)
PNL
(dB)
dB(A)
8
1
3
76.85
3
79.85
65.85
2000
45.85
19
-12
1
3
56.85
3
59.85
45.85
3000
45.85
19
-17
1
3
51.85
3
54.85
40.85
200
45.85
17
8
1
4.75
76.6
3
79.6
65.6
2000
45.85
17
-12
1
4.75
56.6
3
59.6
45.6
3000
45.85
17
-17
1
4.75
51.6
3
54.6
40.6
200
53.55
19
8
1
3
84.55
4
88.55
74.55
2000
53.55
19
-12
1
3
64.55
4
68.55
54.55
3000
53.55
19
-17
1
3
59.55
4
63.55
49.55
200
53.55
17
8
1
4.75
84.3
4
88.3
74.3
2000
53.55
17
-12
1
4.75
64.3
4
68.3
54.3
3000
53.55
17
-17
1
4.75
59.3
4
63.3
49.3
Because the rotational tip Mach numbers were assumed to be the same for each RPM,
Table 40 does not differentiate between the different density altitudes at which the aircraft is flying.
Prediction Using Xrotor
Using the program Xrotor, noise calculations were also performed for the aircraft flying
at 50 and 70 knots at a density altitude of 5000 ft. Because the preliminary noise predictions
were done for 4500 and 6000 RPM, the input of Xrotor was changed to accept an RPM input.
Noise Footprint from Xrotor
Both two- and three-bladed propellers were run through Xrotor and noise footprints of the aircraft at 200 ft, 2000 ft, and 3000 ft away from the aircraft were determined. For a two-bladed
propeller, the following images were produced:
92
Table 41. Noise Footprint Approximations using Xrotor
Distance
RPM Blades
Noise Footprint Image
Away (ft)
4500
2
200
Appendix O, Figure 47
3
6000
2
3
Loudest
dB(A)
59
2000
Appendix O, Figure 48
41
3000
Appendix O, Figure 49
38
200
Appendix O, Figure 50
58
2000
Appendix O, Figure 51
38
3000
Appendix O, Figure 52
35
200
Appendix O, Figure 53
73
2000
Appendix O, Figure 54
54
3000
Appendix O, Figure 55
49
200
Appendix O, Figure 56
72
2000
Appendix O, Figure 57
52
3000
Appendix O, Figure 58
47
For an aircraft at a density altitude of 5000 ft, the noise footprints at 3000 ft, 2000 ft, and
200 ft AGL are about the same as predicted through Roskam and Lan’s noise prediction equations. However, Xrotor shows a much larger difference in the “loudest” noise level of the twobladed propeller versus the three-bladed propeller. The noise level difference at 2000 ft and
3000 ft away are also slightly larger in difference than predicted, but this is due to the way
Xrotor finds noise levels as opposed to the way Roskam and Lan suggest finding noise levels.
Roskam and Lan ask for an approximation using graphs, so because the rotational tip Mach
number were rounded in order to find the right curve to read, the coarseness of the approximations masked the subtle differences of the 2 and three bladed props as well as the 2000 ft and
3000 ft distance.
Conclusions on Noise Levels
The conclusions that we can draw from these values are that the aircraft will be almost
unheard during cruise conditions. At a level of 50-60 dB(A), the noise output of just the propeller is slightly higher than desired, but should not affect the airworthiness of the aircraft. A noise
level of 60 dB is roughly the same noise level as a ventilation fan or hair dryer, 50 dB is the same
noise level as a window air conditioner or a quiet street, and 40 dB is equivalent to a refrigerator
or bird singing. Therefore, if the proposed aircraft were to survey an outdoor dinner party with
93
fine wine and quiet classical music at an altitude of 3000 ft above ground level, it would likely
be heard as a quiet drone.
At higher speeds and closer to the ground, however, the aircraft will approach the same
noise level as that of a noisy office (around 75 to 80 dB), and should then only be used to survey
larger, noisier events, such as a Virginia Tech football game, if the desired effect is noiseless
stealth. At a distance of only 200 ft away, the noise levels of the propeller will not damage the
hearing of the ground crew, and therefore, should not be a major concern. However, as a precaution, it may be desirable to wear earplugs or other hearing protection devices when the aircraft is
at ground level.
One major consideration not made is in the case is that both a primary and a backup internal combustion engine are running at the same time in one of these configurations. This situation would cause as much as a 27 dB noise increase, depending on the spacing of the propellers
and the speed at which they are rotating27. To adjust for this, a 15 dB noise increase should be
added to all of the above numbers run for 4500 RPM. Another consideration, however, is that
the propellers are mounted on a pylon above the fuselage, so any directional noise directed at the
ground may be blocked by the physical structure of the entire aircraft.
Advantages and Disadvantages
A listing of advantages and disadvantages of both configurations to help determine the
best choice possible is shown in Table 42 and Table 43 below in bullet format.
Table 42. Advantages and Disadvantages of IC/IC Configuration
Advantages


No failure detection necessary, autopilot will
automatically adjust the throttle setting of the


Noise will be up to 30 dB(A) louder than
IC/Electric Configuration
remaining engine

Fuel consumption rate will be much higher
Only one set of spare parts or knowledge of

Component reliability for each engine is any-
maintenance required

Disadvantages
where from 1/5 to 1/10 that of an electric option

More fuel or payload may be carried due to
Fuel contamination may end up affecting both
weight savings
engines (unlikely that fuel from two different
Cheaper alternative (half the price!)
sources would be used to fill each tank)

94
Accuracy of syncing the two engines
Table 43. Advantages and Disadvantages of IC/Electric Configuration
Advantages


Component reliability of electric engine is 5 to
10 times that of an IC engine

Starting of electric engine is simple

Fewer parts to maintain

Less noise output from having only one en-
Disadvantages
Failure detection and deployment of second engine is subject to false alarms

How will folding propeller of electric engine be
opened? – extra system

Possible stall in time it takes to respond to
switch in engines
gine run at a time

Possibility of not able to return in a headwind

LiPoly batteries prone to catch fire if punctured

More expensive

Heavier option

Propellers for 70 knot dash requirement (one
engine running at a time) require large pitch or
multiple blades

Forces ground crew to maintain 2 systems
The final conclusion made from the above lists was that the price of an electric back-up
system was far too great for the number of advantages it provided. As a result, the IC/IC Configuration is strongly recommended and used for calculations in the remainder of the paper.
Performance
In order to determine the degree to which the Vulture accomplished its goals, the performance of the aircraft was calculated using equations from Marchman8 and Raymer11.
Breguet Equations: Endurance and Range
Assuming a propeller efficiency of 0.7 for the IC/IC configuration, Breguet range equations were used to find the endurance and range of the aircraft. At constant angle of attack and
cruise velocity, the endurance of the same aircraft is given in the table below for a single internal
combustion (IC) engine and for two IC engines. It is important to note that for each aircraft configuration, the endurance was calculated for the 40 lb of fuel and for the minimum amount of
fuel (26 lb) required for an eight-hour endurance with 45 minutes of emergency fuel.
95
Table 44. Endurance at Constant Angle of Attack and Constant Velocity
Endurance
Single IC Case
(SFC 0.65 lbs/hp-hr)
40 lb fuel
14.32 hr
26 lb fuel
8.74 hr
Double IC Case
(SFC 0.67 lbs/hp-hr)
40 lb fuel
13.89 hr
26 lb fuel
8.73 hr
As shown in Table 44, the single IC case for the IC/Electric configuration has 26 minutes longer
endurance than the IC/IC Configuration. This difference can be attributed to the higher specific
fuel consumption for the IC/IC Configuration running at reduced power.
Using the Breguet range equation (Equation 35, Appendix A), the vehicle’s range for different configurations was calculated. The ranges displayed in Table 45 show that a 15 n.m.
range is not a limiting design condition for this aircraft, as it can be achieved quite easily with
the required high-endurance capabilities.
Table 45. Powered Range for Different Engine Configurations
Range
Single IC Case
(SFC 0.65 lbs/hp-hr)
40 lb fuel
793.3 n.m.
26 lb fuel
516.8 n.m.
Double IC Case
(SFC 0.67 lbs/hp-hr)
40 lb fuel
769.6 n.m.
26 lb fuel
495.0 n.m.
Glide Range
As an additional reliability consideration, the vehicle should have a fairly high no-power
glide range capability. Using a (𝐿/𝐷)𝑚𝑎𝑥 of 12.82 (assuming a CD0 of 0.0257) and Equation 12
from Appendix A, the glide range at different altitudes for the vehicle are given in the table below.
Table 46. No-power Glide Range
Starting Altitude
(ft)
10,000
Glide Range
(n.m.)
21.1
5,000
10.55
3,000
6.33
96
Due to the smaller aspect ratio of the aircraft, the aircraft does not have the ability to return to
base unless it is above 5000 ft, assuming it is 15 n.m. away from base. However, at 6.33 n.m.
closer to base, there may be a higher possibility of recovering the vehicle safely.
Landing Distances
Requirements for the proposed aircraft state that it must be able to land within a 250 ft by
50 ft parking lot. A goal of a 200 ft landing distance was set to allow for the pilots to touchdown
in the first 50 ft. In order to find the ground roll distance for the vehicle, the deceleration was
found through the longitudinal forces on the aircraft as described in Equation 36 from Appendix
A. Using Matlab’s ode45 function and Equation 37 from Appendix A, the deceleration was used
to calculate the stopping distance.
Assumptions
In order to calculate the landing distances, assumptions about the aircraft had to be made.
These assumptions are listed below:
Weight
Wing Planform Area (S)
Tail Planform Area (Stail)
Aspect Ratio (AR)
Oswald Efficiency Factor (ε)
Parasitic Drag Coefficient (CD0)
Maximum Lift Coefficient (CLmax)
Zero-Angle Lift Coefficient (CL0)
245 lb
76 ft2
14 ft2
6.25
0.861
0.0257
1.48
0.3
Density at Altitude (Sea Level)
Acceleration Due to Gravity
0.002378 slug/ft3
32.2 ft/s2
Coefficient of Friction on Main Gear (μmain)
Coefficient of Friction on Tail (μtail)
Percent of Weight on Main Wheels (Wmain/Wtotal)
Horizontal Touchdown Speed (VTD)
0.4 (with brakes)
0.8 (rubber pad on concrete)
85%
1.15 Vstall
Ground Roll Stopping Distances
The final stopping distance given by the MatLab program in Appendix P for four different cases
is detailed in the table below.
97
Table 47. Ground Roll Distances for Vulture
𝑪𝑳𝒎𝒂𝒙
MTOGW
Half Fuela
Half Fuelb
(245 lbs)
(225.2 lbs)
(217.9 lbs)
Case 1: No Brakes/No Flaps
1.48
310.5
285.5
276.2
Case 2: Brakes/ No Flaps
1.48
220.1
202.3
195.7
Case 3: No Brakes/ Flaps
1.5
306.3
281.6
272.4
1.6
286.4
263.2
254.7
1.7
268.2
246.5
238.5
1.8
251.6
231.2
223.8
1.9
236.4
217.2
210.2
1.5
217.1
199.5
193.1
1.6
202.6
186.2
180.1
1.7
189.1
173.9
168.2
1.8
176.7
162.4
157.1
Case 4: Brakes/Flaps
From this information, it is possible to deduce that a mere application of brakes is capable
of stopping the vehicle in 200 ft when landing with half fuel. While the ideal case for reliability
would be to land without brakes or flaps within the desired distance, brakes will allow the vehicle to avoid added structural complications and marginal performance from flaps and the use of
nets or secondary capture devices.
Secondary Capture Devices
Secondary capture devices, such as nets and wires were considered for use in the case a
paved runway was not available. A net capture is possible for use in the case that the runway is
shorter than expected, or that the aircraft needs to make an emergency landing in the early stages
of flight. Assuming the net is placed near the end of the runway, for a ground run of 200 ft, the
landing distance and velocity versus time is given in Figure 37.
a
b
40 lbs fuel initially
26 lbs fuel initially
98
Ground Roll - No Brakes/ No Flaps
Distance X, ft
300
200
100
0
0
2
4
6
Time t, seconds
8
10
12
8
10
12
Velocity V, ft/s
60
X: 5.31
Y: 27.75
40
20
0
0
2
4
6
Time t, seconds
Figure 37. Ground Roll with No Brakes or Flaps
The aircraft hits the net approximately five seconds after landing at 16.4 knots. The distance the
net would need to displace to stop the aircraft was calculated using
where V1 is zero to stop the aircraft, V0 is the initial landing speed of 16.4 knots, the deceleration
a is assumed to be -3 times the acceleration due to gravity, and s is the displacement. This gives
a net displacement requirement of four feet.
Systems Integration
A list of all system parts not falling under any other categories is listed in Table 48. The
dimensions, weights and small notes on placement are listed as well. Larger systems that do not
fall under other major systems such as the autopilot, the batteries, and the servos are further explained in the following sections.
99
Table 48. Parts List
Weight
(lb)
Dimensions
Location notes
FCS
Required
Piccolo II
0.52
5.6"×3.0"×2.4"
Servos
(S3306 1/5 scale HiTorque/Speed)
0.282
2.6"×1.18"×2.25"
GPS Antenna
UHF Antenna
(900MHz or 2.4 GHz)
Piccolo Battery (8V-20V)
(1 pack, need 2)
(TP400032 3cell LiPoly
4000mAh)
Servo Battery (4.8V- 6.0V)
(1 pack, need 2)
(Elite 400mAh 5cell)
very small
~4"-6" long
90degrees to the flight axis within 6in
of the CG
Only 2 per ailerons, elevator then 1
per rudder, and throttles making 8-9
total. Mounted on two mounting
blocks for each servo
the tail
very small
~4"-6" long
away from any other antennae
0.57
7.5"×1.875"×0.6875"
where ever they need to be for CG
balance
0.7
4.36"×1.8"×0.9"
where ever they need to be for CG
balance
Control Linkages
very small
Magnetometer
(Honeywell HMR2300)
0.216
4.20"×1.5"×0.876"
Manual Override(Rx)
Manual Override(RxMuX)
0.0625
0.0101
1.1”×2.2"×0.8"
2.25"×2.25"×0.162"
Total Pressure Tapping
0
0.252" diameter
Static Pressure Tapping
0
0.252" diameter
PC-104 (failure detection)
2.3
3.6"×3.8"×0.6"
2nd Piccolo II
0.52
5.6"×3.0"×2.4"
the control lines for the servos and
stuff
as far back within the fuselage the
connector of the magnetometer goes
toward the direction of flight with the
label up
can put where ever you have room
can put where ever you have room
on the wing or fuselage away from
the prop wake
on the wing or fuselage away from
the prop wake and fuselage effects
Optional
DGPS
Payload
45
Payload internal
Payload Batt (12V & 10W)
(pack, need 2)
(TP45004 4cell LiPoly
4500mAh)
--
15"×11"×7"
0.99
6.25"×1.875"×1.1875"
--
5" diameter
0.27
15.5"×11.5"×7.5"
very small
~4"-6" long
Camera gimbal
Faraday 200 copper wire
mesh cage
Antenna (ground plane)
100
don't really need it unless we have
backup piccolo
may need one need to study cost
analysis
Upgrade from GPS
Towards the front of the fuselage
where ever they need to be for CG
balance
Add to the middle of the internal payload for now
around the internal payload
away fom any other antennae
IC Engine
IC Engine
6.5
11.89"×12.6"×3.74"
IC Propellor
IC Spinner
IC Ignition Module
Ignition Batt (7.4V)
(1 pack, need 2)
(TP400022 2cell LiPoly
4000mAh)
IC Switch
1.32
0.19
0.44
27.56" diameter
4.5" dia w/5.51" length
2.36"×2.36"×1.57"
Nacelle
IC Propshaft
Nacelle
0.4
5.25"×1.875"×0.6875"
where ever they need to be for CG
balance
0.15
1.57"×1.57"×1.57"
Nacelle/Fuselage
5"dia w/1.1" length
IC Propshaft
Nacelle/Fuselage
Nacelle
Nacelle
Alternator
Voltage Regulator
Carburettor
Exhaust Pipe
Venturi (inplace of fuel
pump)
Small
Fuel Tank
39.68
Outlet (fuel pipe) dia
0.0945"
Inlet (freestream) dia
0.185", w/ 0.169"
length
19.69"×7.87"×8.66"
Electric Engine
3.42
Electric Propellor
1.32
Electric Spinner
0.22
Speed Controller
Elec Engine Batteries
(14 off)
[2 series sets of 7
in parallel]
0.6
Electric Engine
4.06" dia w/ 2.93"
length
24.02" diameter
4.5" dia w/ 4.88"
length
4.69"×1.06"×2.52"
22.51
11.57"×1.75"×1.25"
Nacelle
place the inlet in the freestream but
in an area where it can still reach the
fuel tank
Fuselage
Nacelle
Nacelle
Electrice Propshaft
Fuselage
place where needed for CG balance
Electromagnetic Shielding
Electromagnetic shielding must be taken into consideration for onboard systems operating on similar wavelengths or power output. Some systems that may interfere, for example, include the GPS and the magnetometer. To protect from electromagnetic interference, a Faraday
cage of 200 mesh copper was designed to fit over the internal payload. Aluminum shielding
wrapped about the wires and antennae will protect them from electromagnetic inference as well.
Autopilot
The autopilot chosen for the initial design and testing of the UAV was the Cloud Cap
Piccolo II Autopilot. The Piccolo II (Figure 38) is a COTS solution and comes equipped with an
101
inertial measurement unit, full differential GPS, pitot and static air measurement sensors and 900
MHz communication. This autopilot was chosen primarily due to the availability of both the
hardware and the documentation. The Piccolo is not International Trade and Arms Regulation
(ITAR) restricted, so the documentation would be available to the British members of the team.
The Piccolo II also has a fairly robust modeling program to find the autopilot gains, which will
simplify the integration process. Lastly, Cloud Cap support is very helpful in the integration and
testing of the autopilot and will test models submitted to them.
Figure 38. Cloud Cap Piccolo II Autopilot28
Batteries
To increase the reliability of the electrical system, an alternator was placed on the primary engine. The purpose of this alternator is to provide power to the payload, the autopilot, and to
each of the servos. Even though an alternator was placed on board to service the autopilot batteries, the servo batteries and the ignition batteries were sized to complete the mission with no alternator. The payload, however, is not flight critical, so the batteries can be downsized. A summary of the chosen batteries, weights and costs are given below in Table 49.
Table 49. Batteries Chosen for Vulture
Autopilot Batteries
Redundant TP400032 3 Cell LiPoly 4000 mAh
Weight
1.14 lb
Cost
$299.98
Size
7.5 x 3.75 x 1.375"
Servo Batteries
Redundant Elite 4000 mAh NiMH
Weight
1.4 lb
Cost
$67.00
Size
4.36 x 3.6 x 1.8"
102
Ignition Batteries
IC/IC Configuration
Four TP400022 2 Cell LiPoly 4000 mAh
Weight
1.6 lb
Cost
$399.96
Size
2 packs of 5.25 x 3.75 x 1.375"
IC/Electric Configuration
Two TP400022 2 Cell LiPoly 4000 mAh
Weight
Cost
Size
0.8 lb
$199.98
5.25 x 3.75 x 1.375"
Payload Batteries
TP21004 4 Cell 2100 mAh "Pro Lite" 14.8V LiPoly
Weight
0.3875 lb
Cost
$99.99
Size
4 x 1.25 x 1"
The chosen alternator for the aircraft is a Sullivan Face Type Alternator. This alternator is
simple, contains few moving parts, and is thus reliable and easy to replace or repair if a failure
occurs. This alternator can be mounted on the rear output shaft or on the front output shaft behind the propeller. The output of the alternator is 10 watts and should be able to run the payload
under its own power.
The autopilot batteries need to be between 8V and 20V and able to run the autopilot at
around 0.35 amps for 8 hours. This current draw was found by testing what the Piccolo II autopilot needs running at the lowest voltage it will see. The chosen battery pack for the autopilot was
the Thunder Power TP400032. Redundant battery packs have been added onboard for reliability.
The packs would simply be put in parallel so if one fails, the others can complete the mission.
The servo batteries were sized off of the estimate that the servos will draw a maximum of
0.2 amps and to allow the servos to draw maximum current for three hours. The servo packs
were also made of a 5 cell nickel-metal hydride (NiMH) battery, so the voltage would not be
high enough to deteriorate the life of the servos. Redundant packs are also added to the servos in
the same method used for the autopilot to increase reliability.
The ignition batteries were sized from the requirements of the chosen engine. The engine
needs a 7.4 V LiPo with 8000 mAh. For IC/IC Configuration, this will need to be doubled, but
no added weight will come from these batteries, as they will not be onboard. The chosen system
103
for the ignition system includes two Thunder Power TP400022 LiPo packs for each engine. The
total weight is for two internal combustion engines.
Since the payload runs primarily off of power supplied by the alternator, the battery can
be smaller, since it will only be used to maintain a constant voltage to the payload. The chosen
battery was a Thunder Power TP21004.
Servo Sizing and Selection
The servos were a primary concern for the reliability of the aircraft. After the control surface size was chosen, the servos were sized by finding the moment required to deflect the control
surface at cruise speed as well as dash speed. It was determined that a 15º deflection was needed
for the ailerons and a 25º deflection was needed for both the elevator and rudder at cruise speed.
The calculated torque required was 260 oz-in for the ailerons and elevator and 310 oz-in for the
rudder. The servo selected was a Futaba S3306 rated for 333 oz-in of torque at 6 volts. This servo gave the ability to deflect all the surfaces the amount needed at cruise and the ability to deflect each 10º at dash speed.
To increase the reliability of the aircraft the ailerons were split into halves. If one servo
failed the other can compensate and get the aircraft home. The selected servo remained the same.
Firstly, the S3306 would give more than the amount of torque needed for the application and
would increase the factor of safety. Secondly, these are the lightest commercially-available hobby servos in the torque range required.
Ground Control
Autopilot Ground Control Station
The ground control station provided with the Piccolo II autopilot will be set up as illustrated in the diagram below (Figure 39).
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Figure 39. Ground Control Architecture
The autopilot is controlled by a ground station made by Cloud Cap, pictured in Figure 40,
using a radio frequency of either 900 MHz or 2.4 GHz. The ground station connects to a computer running the Piccolo Command Center. From the command center, the user can quickly
view all of the Piccolo stats including GPS health, COMM health, RPM, airspeed, altitude, etc.
The command center is also the interface to create and execute flight plans using point-and-click
navigation software.
Figure 40. Cloud Cap Ground Station
To increase reliability, another computer can be networked to the first. By default, all the stats
sent to the ground station are sent to the local network. A second computer could analyze this
data and test for certain conditions, such as too high a g-loading, overheating Piccolo, etc. If
these conditions were exceeded the computer could send the command for the autopilot to return
105
to home. This solution allows critical flight statistics to be monitored autonomously; however
extensive software design would be needed to avoid false positives and/or negatives.
Another completely separate system for manual override will be able to communicate
with the aircraft at a 72 MHz frequency. Should the ground station, operator interface, and/or
pilot console fail, then the manual override controller can be used. This device will provide the
ability to fly the aircraft manually on a completely separate radio frequency. In the event that the
primary radio frequency used for contact with the autopilot gets jammed, this controller will be
useful if the aircraft is expected to complete part of a mission instead of returning to base.
The last computer set up will be a system dedicated solely to downloading sensory information provided by the payload. This computer can interact with the aircraft through a completely separate radio frequency from both the manual override system and the Piccolo II ground
station.
Launcher
According to the specifications given by NAVAIR, the aircraft must be capable of a
pneumatic launch. Research yielded several launchers that were in the correct weight and speed
range. The common length was found to be up to 45 feet when extended, so the g-loading on the
aircraft was found to be 4.2 for launch. To accommodate for a smaller launcher and a factor of
safety, the structure was designed to take a 6-g longitudinal loading.
Landing Gear
Configuration
The final landing gear configuration chosen was a tail-dragger setup. This was chosen
because it offered a lightweight and simple solution over a tricycle setup. One of the biggest disadvantages of a tail-dragger is handling a crosswind on takeoff, however, since the proposed
UAV is pneumatically launched, this is not a problem. Instead of a tail wheel, a skid is used. Due
to the launched takeoff, it was determined that the ability to turn the vehicle on the ground is not
very important. A skid also helps to stop the vehicle in a shorter distance and will actually help
to straighten out the vehicle upon landing.
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Sizing and Structure
The gear was initially sized using angle recommendations given by Raymer11 as well as
the constraints of the aircraft. Figure 41 shows the aircraft angle recommendations. A tail scrape
angle of 10º was chosen to put the wing close to stall when the aircraft is resting on all three
gear. The angle between the main gear and the CG of the aircraft was set at 20º. This was chosen
so a slight movement of the CG moving does not force the angle outside of the range given by
Raymer. The lateral spacing of the gear was determined from the 25º restriction.
Figure 41. Landing Gear Sizing11
Once the gear height and lateral spacing was known, the breadth and thickness of the
landing strut was calculated. Aircraft grade aluminum was chosen to simplify the manufacturing
of the strut. A basic bending calculation was done on each side of the strut, assuming a 4-g, 2point landing. It was then determined that to keep the deflection under the recommended 37%, a
hollow box section with a breath of 2.25”, a thickness of 0.75” and a wall thickness of 0.125”
was needed. The tail skid would also be made of a curved piece of aluminum with an attached
replaceable rubber pad.
Wheels and Brakes
The chosen wheels are Glennis 6” Remotely Piloted Vehicle (RPV) wheels. These
wheels are specifically designed for larger UAV operations and can handle the weight demanded
107
by this aircraft. The wheels come in varying tread styles and can be customized for the specific
mission. Glennis also produces a pneumatic brake that can be used with this wheel. The brake
unit attaches to the axle and fits inside of the wheel hub. When activated, the air pressure will
push the brake pad into the hub, helping to stop the vehicle. This system will require a small air
tank about the size of a hairspray can and a single servo to actuate the valve.
Reliability
Aircraft reliability was the main focus of the project’s design. Several methods to analyze and improve reliability were used in this design, including:

Fault Tree Analysis (FTA)

Failure Modes and Effects Analysis (FMEA)

Mechanical Failure Models
Each of these methods was used in one of the following areas, which were determined to be the
most critical to reliability:

Engine/propulsion

Flight Control/Communications

Structures

Human/Ground Station
Fault Tree Analysis (FTA)
FTA is a graphical representation of system hierarchy and its dependency of components,
usually derived from FMEA data. The fault tree shows the minimal cut sets of component failures and their relative probabilities. The collection of all of the probabilities constitutes the
probability of failure of the entire system.
Failure Modes and Effects Analysis (FMEA)
FMEA is a procedure to calculate all of the possible ways for a system to fail, all the way
down to the minimal cut sets. Each of these has a certain probability associated to the possibility/proportion of time for the event of a failure to occur for that component. The analysis is done
to determine whether or not the component failure by itself constitutes a total failure for the system analyzed. Once every possible failure mode is accounted for, the cumulative probability for
108
the system to fail is then determined. FMEA is usually represented by a fault tree to depict the
hierarchical structure of the system, and to show subsystem dependence on components.
Mechanical Failure Models
Mechanical failure models provide a way to analyze mean operating stresses on a component and compare them to the material strength. This representation of mechanical component
reliability is known as the “stress-strength interference model”29. This model contains a joint
probability distribution function in which one random variable (Y) represents the operating
stresses of the device and another represents material strength. The dispersions of these two random variables are represented by hY(y) and gx(x) respectively. The component reliability, F, is
when the strength exceeds the operating stresses as given by Equation 39 from Appendix A. The
failure probability is just the complement of the reliability. Each of these dispersions of the respective variables can be represented by either normal, lognormal, or Weibull distributions.
Reliability Data
It can be concluded that, for the majority of components, reliability data is very difficult
to obtain. Two methods to get this data include historical databases and testing and simulation.
While the former method can be more inaccurate to relate across designs/components, the latter
tends to be impractical. Historical databases, to begin, are very scarce with regard to aircraft
components. Depending on the component or system, there might be no data available. For this
project, massive amounts of time were spent on searching the internet, literature, and companies
to find data. Data that was acquired included failure modes and probabilities of the autopilot
system, failure modes of IC engines (no probabilities), failure modes and probabilities for electric motors, material coupon data for metal and composite materials with mean and standard deviation of physical properties.
This data was vital to serve the purpose of analyzing and improving reliability of the design. However, the data that was required of the propulsion system was not enough to analyze.
For the IC engine, there were no probabilities to calculate the proportion of time that the minimal
cut sets will fail.
Propulsion Reliability
It was decided that, following the method of redundancy, the aircraft designed was to
have two engines. The two types of engine that were deemed to be appropriate were the IC and
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electric engines. All combinations of the two were reviewed and conclusions were made on two
designs:

IC/IC Configuration: two of the same model IC engines operating at partial power simultaneously during normal flight. A single engine of this model would be sufficient
enough to operate by itself; so, throttle setting of one engine would increase if the other
were to fail.

IC/Electric Configuration: one IC engine would be used as the main tractor engine.
This engine would be operating by itself at the sufficient power. If an event occurs that
causes the engine to lose power, a secondary electric motor will activate that will be sufficient enough to return the aircraft to base.
The fault tree for the internal combustion and electric engines are provided
IC Engine Failure
0.001
Lack of
Compression
Bad Fuel Mix
Clogged Air
Intake
Out of Gas
Impurities
Incorrect
Amount of
Fuel Supply
Hole in
Cylinder
Worn Piston
Rings
Lack of Spark
Intake/
Exhaust
Valves not
Sealing
Properly
Damaged
Wire
Figure 42. Internal Combustion Engine Fault Tree
110
Worn Spark
Plug
Ignition
Timing
Off
Electric Motor Fails
2.425E-4
Motor Fails
4 Batteries
Fail
1E-9
Shaft Failure
6.6E-6
Bearing
Failure
0.000166
Rotor Failure Stator Failure
1.66E-5
5.33E-5
Figure 43. Electric Engine Fault Tree
The fault tree for the electric motor contains probabilities for the minimal cut sets while
the IC engine does not. This was because no data could be obtained for the IC engine. The
overall system probabilities for failure for the IC and electric engines are 0.001 and 2.425×10-4
respectively.
The probabilities of failure for the two proposed combinations are the multiplication of
the failure probabilities of each engine used. This is because a propulsion system total failure
would be the result of both engines failing in the same mission. Assuming eight hours per mission, and that each engine failure is independent of the other, the probabilities of failure for the
IC/Electric and the IC/IC Configurations are 6.4×10-5 and 1.55×10-5 respectively.
The IC/Electric Configuration is five times more reliable, but for the requirements of this
aircraft, both are equally acceptable. Assuming eight hour per mission and up to five missions
per week, the mean time between failure for the IC/Electric and the IC/IC Configurations are approximately 60 years and 247 years respectively. An operational requirement of 20 years of service makes both configurations acceptable.
111
Autopilot Reliability
Just as in the propulsion system, the FTA method was used to determine the reliability for
the Piccolo II autopilot (Figure 44). Data for the autopilot system was much easier to obtain than
other components due to communication via email with the company. Historical failure mode
data based on customer reports and returns was obtained and reported through this communication. The percentages and probabilities are not based around units of time for operation, but rather on the percentages of total sales for the system. The data represents the percentage of time
that the component is operational. It contains probabilities for both availability and reliability,
depending if it causes an operational failure or mission abort.
Figure 44. Autopilot Fault Tree
From the FMEA and fault tree, it can be noticed that the Piccolo II system is comprised
of both the autopilot and ground station. Both process and firmware or software related failures
are included in this fault tree. Compared to the other components of the aircraft, the autopilot is
has a much higher failure rate at 2.7%. There were no other autopilots for comparison available
as this system was chosen based upon availability before the analysis was performed.
Structural Reliability
To determine the reliability of the aircraft’s structure, the mechanical failure model was
used. The portions of the aircraft that were deemed critical to construct a structural analysis in-
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cluded the main and rear wing spars, the tail boom, and the fuselage. The analysis used load data, maximum bending moment, and operating stresses that the component would experience during the course of normal flight operation (Table 27).
The standard deviations of the structural loadings were then converted to coincide with
the means of the operating stresses. The operating stresses with their respective standard deviations were then compared to material physical property data. The stress-strength interference
was modeled and probabilities based on bending moment failure were calculated:
Table 50. Structural Reliability Calculations
Operating Stresses
Component
(psi)
Mean
St. Dev.
Main Wing Spar
26893
475.5
Rear Wing Spar
11219
Tail Boom
Fuselage
Material
Material Strength
(psi)
Probability
of Failure
Mean
St. Dev.
Al 6061-T6
41670
3512
1.40×10-4
463.0
Al 6061-T6
41670
3512
1.40×10-17
26259
577.1
Carbon Fiber
87023
3512
3.01×10-65
23138
508.5
Al 6061-T6
41670
3512
4.17×10-7
Total
1.404×10-4
From Table 50, it can be concluded that the main wing spar is the largest contributing factor to
the probability of structural failure.
A more in-depth model can be performed with more components to analyze. Ideally, a
structural analysis will include all materials of as many components as possible. An entire analysis, for example, would be constructed for the wing. Not only would the spars be taken into
account, but the wing skin as well. Environmental factors, operational and manufacturing defects, gust loading, and aircraft age would be taken into account as well. These factors would
provide either a translational or scale factor shift to reduce the material strength distribution,
providing a more realistic modeling schema.
Monte Carlo simulation trials constructed for each structural component would produce
random scale factors and probabilities. This algorithm is known as the Northrop Grumman
Commercial Aircraft Division (NGCAD) Probabilistic Design Model. The flowchart is displayed in Appendix Q. This approach, however, was constructed for existing aircraft to provide
means and targets for reliability improvement, which is not the objective of this project. For this
113
reason, and time constraints as well, this approach was not taken but simplified to the aforementioned process.
Overall Reliability
The reliability of the entire aircraft “system” is derived from the cumulative analyses that
were performed for each component. A summarized fault tree is shown below with all of the
components and their respective total probabilities of failure:
Aircraft System Failure
IC/IC Configuration: 0.02720
IC/Electric Configuration: 0.02716
Propulsion Failure
IC/IC Configuration: 6.4×10-5
IC/Electric Configuration: 1.55×10
Structural Failure
Autopilot Failure
1.40417×10-4
0.027
-5
Figure 45. Overall System Fault Tree
The probability for total aircraft system failure can be derived from either of the three independent subsystems or any combination of them. Using the IC/IC Configuration as an example, the complement of the probability of the aircraft system failure of 0.0272, the reliability of
the UAV is 0.9728. The mishap rate that would occur would be 2720 per 100,000 hours. This
obviously exceeds the goal of 1 uncontrolled crash per 100,000 hours. The large majority of the
unreliable contribution is from the Piccolo II autopilot. If a custom system was used instead of a
COTS product, the system would yield a much higher reliability closer to the goal.
Costs
Currently, the rising expense of UAVs is a major topic of debate in the US department of
defense. This is largely associated with unreliability and unavailability of a UAV to perform a
114
given mission. Subsequently, the expense of manufacturing, maintaining, and operating the UAV
is a major aspect of the design phase.
While additional costs are incurred in an unmanned vehicle due to the autopilot technology and operating costs, it is still remains beneficial to use an unmanned system. The elimination
of the possibility of injury and human error paired with the elimination of the pilot and necessary
cockpit systems is estimated to save $1,500 per pound.9
Early in the design phase, when selecting components, it was considered appropriate to
use COTS products. By utilizing such items as opposed to custom-made ones, additional manufacturing costs could be avoided. In the future, if the aircraft were mass-produced, manufacturing
costs would theoretically decrease due to large quantities required.
Cost Model
In order to account for appropriate costs, a cost model was developed. There were five
main categories: operation costs, maintenance costs, component costs, manufacturing costs, and
support costs. The operation costs account for consumables such as batteries and fuel. Maintenance costs estimate the time required to keep the aircraft functioning including replacement of
basic components. Manufacturing costs mainly consist of the labor required to produce the vehicle, as the material costs used are included in component costs. Support costs encompass the
preparation work at launch and recovery required by aircraft operators each mission.
Support
Costs
Maintenance
Costs
Operational
Costs
Manufacturing
Costs
Cost
Model
Elements
Component
Breakdown
Component Cost
The cost of components required to build an autonomous vehicle contribute to the majority of the overall cost. There are three major categories that a component can contribute to:
115

Structure – The cost of the structure of the vehicle was roughly $4,000 USD. This
accounted for the structural components in the tail, wing, fuselage, pylon, and landing
gear. It also includes the servos used on the control surfaces. As mentioned previously, it was desirable to use COTS parts in an effort to avoid additional costs customization. The major contributors in this category were: aluminum frames for the fuselage
($1038.00 USD) and the carbon fiber tail boom ($550.00 USD).

Propulsion– The cost of the propulsion systems were calculated for two separate cases: two IC engines and then one IC engine with one electric motor. When the UAV is
operating in the IC/IC Configuration, the total cost is calculated to be $5,900 USD.
For the IC/Electric Configuration, the cost is $5,700 USD. The most expensive component again is the IC engine. It is important to note this case does not include the
batteries required to operate the motor as they are accounted for in the Operational
Costs section.

Avionics– The avionics sector of the component costs accounted for 66% of the total
cost. This is primarily due to the fact that the Piccolo Autopilot costs $28,000.
Figure 46 shows the breakdown of the costs of the components.
$30,000
$28,000
$25,000
$20,000
$15,000
$10,000
$5,900
$5,000
$4,000
$0
Avionics
Propulsion
Structure
Figure 46. Cost of Aircraft Components
Operational Costs
The operational costs for the Vulture UAV encompass the main components for the
“Use” phase of the aircraft. This includes the fuel required to operate the vehicle for the specified
length of the mission as well as the appropriate amount of batteries required to operate the electronic components.
116
The amount of fuel required for the entire flight is approximately 40 lbs of aviation gas.
This translates to nearly 6.5 gallons of fuel. The average cost of aviation gas over the past year
has been $4.50 per gallon. This translates to a total cost of fuel to be $29.25.
The electric motor requires 14 Lithium Polymer Pack batteries. These are capable of
22.2V and 6600 mAh. Each unit is $439.00 bringing to cost of operation for the electric motor to
$6,146 USD. In addition to the batteries used for the electric motor, two additional batteries are
required for the Autopilot. These batteries are rated for 11.1 volts and 4350 mAh and are priced
at $159.90 each.
Maintenance Cost
Maintenance is of utmost importance for an autonomous vehicle in order to keep it running smoothly, safely, and efficiently. The cost of maintenance should not be overlooked, as it
requires labor, money, and new parts to adequately maintain any type of system. The main elements of the cost contributed by routine maintenance for this vehicle allowed for:

Replacement of main engine once per annum

Replacement of all nine (9) servos
Support Cost
The cost of support labor for the vehicle was taken from an estimate of the average wage
of a U.S. soldier. Based on data gathered from the U.S. army, the average lower-ranking soldier
earns $14.00 per hour. It is estimated the soldier will be involved in routine maintenance for up
to two hours per flight. This allows for such activities as changing the fuel, loading the batteries,
installing the payload, and an overall system check. In addition to routine maintenance performed during this time, the soldier will also be provided support for launch and recovery activities. The UAV needs to be mounted on the rail-launcher as well as gathered up post-mission. It
was assumed that there would be two individuals on hand to perform this work per flight.
Manufacturing Cost
It is difficult to estimate the amount of manufacturing and costs it will require to build the
UAV. Costs will vary among manufacturers. The costs will also vary depending on the degree of
customization and/or special equipment required to produce the part.
117
Overall Costs
The question often arises with regards to the benefit/costs analysis for two engine setups.
In order to fully examine this it is important to consider the costs incurred from the use or operational phase. These can be seen in Table 51.
Table 51. Operational Costs of Vulture
Configuration Engine
IC/IC
$5,900
Electric
$5,700
Operational Phase
$29.25
$6,146
Total Cost
$5,929
$11,846
The cost analysis is important to understanding the full expense of the project and UAV
over its lifetime. By considering as many inputs to the cost model as possible, it reduces the
chance for hidden expenses along the way.
Conclusions
All goals of this design project were met with the exception of the noise requirement and
the reliability. This is due, however, to the fact that the position of the fuselage and wing were
not taken into consideration for the noise calculations and the reliability of the autopilot was on
the order of thousands of times larger than all other calculated component reliabilities. The noise
and reliability disciplines, however, are among the hardest to predict, as it often requires actual
testing or esoteric equations that may not apply to the vehicle. Reliability data is especially hard
to obtain for numerical calculations. In contrast, disciplines with more tools helped to create an
aircraft that offers much more than required. Performance of the aircraft more than exceeds what
was expected, and should provide well for large crowd surveillance missions.
Further Work
During the summer of 2008, a small-scale model weighing a sixth of the designed aircraft
will be built and tested. The creation of a larger scale wind tunnel model for testing in the Virginia Tech Stability Wind Tunnel should first be made and tested for aerodynamic forces and
balances. This will help validate the theoretical calculations made. Plans for a full-scale model
will be made after the small scale model is accomplished.
118
Desired Work
A much more detailed reliability analysis should be conducted to grasp a better estimate
of MTBF and availability of the aircraft. Due to the time constraints, resources available, and
innovation of the design, the work that was done was accurate enough to get feasible results.
Another implication of the time constraint was the design method entirely. A preferred approach
would be to analyze and compare alternatives for each component. This would allow the design
team to weigh the reliability of each alternative against one another along with other constraints
such as cost, weight, and integration with the rest of the aircraft subsystems. The limited time
and resources available in the initial phases of the design prevented the team from taking this approach.
119
Appendix A – List of Equations
Eqn. #
Equation 1
Equation 2
Equation 3
Equation
2𝑊
𝐶𝐿 𝜌𝑉 2
1 2
2𝐾𝑊 2
𝑃 = 𝐶𝐷0 𝜌𝑉 𝑆 +
2
𝜌𝑉𝑆
𝑆=
𝐾=
1
𝜋𝐴𝑅𝑒
𝟑
Equation 4
Equation 5
Equation 6
𝜼𝒑
𝑪𝑳 𝟐
𝟏
𝟏
𝑬=
(
−
)
√𝟐𝝆𝑺
𝜸𝒑
𝑪𝑫 √𝑾𝟐 √𝑾𝟏
𝐸=
𝜂𝑝 1 𝐶𝐿 𝑊1
ln
𝛾𝑝 𝑉 𝐶𝐷 𝑊2
𝑳
𝟏
( )
=
𝑫 𝒎𝒂𝒙 𝟐√𝑪𝑫𝟎 𝑲
D  CD

1
2
V 2 S 
2

CL 


  CD0 


AR
e


Equation 7
 CD0

1
2
2 V S  
Equation 8
 
Pavail  TV  PSL  alt 
  SL 
Equation 9
Vstall 
Equation 10
CLMP 

1
2
V 2 S 
W2
 AR e 12 V 2 S
nW
C Lmax
1
2
S
3CD0
k
 3CLMD
Equation 12
dh V  dV  Pavail  Preq 
 

dt g  dt 
W
Rglide  hL D max
Equation 13
E
Equation 11
 P 1 CL  W1 
ln  
 P V CD  W2 
120


Equation 14
Equation 15
Equation 16
Equation 178
Equation 18
d g 2
SCL

n 1  g
dt V
2W
C
Length  AWO
F (t )  P(T  t )   f (u )du for t > 0
𝑃
𝑞𝐶𝐷0
𝑘𝑛2 𝑊
1 𝑑ℎ 1 𝑑𝑉
=[
+(
)+
+
]𝑉
𝑊
𝑊
𝑞 𝑆
𝑉 𝑑𝑡 𝑔 𝑑𝑡
𝑆
C ht c wing S wing
Horizontal tail area: S ht 
Lht
S vt 
Equation 19
Vertical tail area:
Equation 20
𝑊 1 2
= 𝜌𝑉 𝐶𝐿𝑐𝑟𝑢𝑖𝑠𝑒
𝑆
2
Equation 21
n2  1
n
 
Cvt bwing S wing
Lvt
𝑇
𝑊 𝑘𝑛2 𝑊
1 𝑑ℎ 1 𝑑𝑉
= 𝑞𝐶𝐷0 +
+
+
𝑊
𝑆
𝑞 𝑆
𝑉 𝑑𝑡
𝑔 𝑑𝑡
1
where: 𝑞 = 1⁄2 𝜌𝑉 2 𝑘 = 𝜋𝐴𝑅𝑒
𝐶𝐹 =
Equation 22
0.455
(𝑙𝑜𝑔10 (𝑅𝑒))2.58 (1 + 0.144 𝑀2 )0.65
𝑆
where: 𝐶𝐷0= 𝐶𝐹 𝐹𝑄 𝑆𝑤𝑒𝑡 and
𝑟𝑒𝑓
Equation 2326
0.6
4
𝐹 = [1 + 𝑥 (𝑡⁄𝑐 ) + 100(𝑡⁄𝑐 ) ] [1.34𝑀0.18 (𝑐𝑜𝑠(∆𝑚 )0.28 )]
( ⁄𝑐)
𝜂𝑚𝑒𝑐ℎ 𝑀𝐸𝑃 ∗ 𝑣𝑐𝑦𝑙 ∗ 𝑅𝑃𝑀 ∗ 𝑁𝑝𝑖𝑠𝑡𝑜𝑛
𝑆𝐻𝑃 =
792,000
𝑆𝐹𝐶 =
Equation 2426
𝑤̇𝑓
𝑆𝐻𝑃
Equation 25
𝑃𝑎𝑣𝑎𝑖𝑙 = (𝑒𝑚𝑓 − 𝐼 ∙ 𝑅𝑡𝑜𝑡𝑎𝑙 )(𝐼 − 𝐼0 )
Equation 26
(𝑒𝑚𝑓 − 𝐼 ∙ 𝑅𝑡𝑜𝑡𝑎𝑙 )(𝐼 − 𝐼0 ) = 𝑘𝑝𝑟𝑜𝑝 ∙ 𝑅𝑃𝑀𝑃𝐹 ∙ 𝑑 4 ∙ 𝑝
Equation 27
𝑅𝑡𝑜𝑡𝑎𝑙 = 𝑅𝑏𝑎𝑡𝑡 + 𝑅𝑒𝑠𝑐 + 𝑅𝑚𝑜𝑡𝑜𝑟 + 𝑅𝑤𝑖𝑟𝑒𝑠
0.9𝑉𝑒𝑙𝑒𝑐
𝜂𝑒𝑙𝑒𝑐 =
Equation 28
𝑉𝑒𝑙𝑒𝑐 +
Equation 29
1 𝜂 𝑃𝑎𝑣𝑎𝑖𝑙
2
−𝑉𝑒𝑙𝑒𝑐 ± √𝑉𝑒𝑙𝑒𝑐
+ 2 𝜌𝐴 𝑒𝑙𝑒𝑐
𝑉
𝑒𝑙𝑒𝑐
2
𝑇=
121
𝜂𝑒𝑙𝑒𝑐 𝑃𝑎𝑣𝑎𝑖𝑙
𝑉𝑒𝑙𝑒𝑐
Equation 30
𝐹𝐿1 = 16.037 log 𝑆𝐻𝑃 + 35.377
𝐹𝐿1 = 15.873 log 𝑆𝐻𝑃 + 39.661
𝑁𝐶 = 4.328𝑒 𝐵
𝑂𝑆𝑃𝐿 = 𝐹𝐿1 + 𝐹𝐿2 + 𝐹𝐿3 + 𝐷𝐼 + 𝑁𝐶
𝑃𝑁𝐿 = 𝑂𝑆𝑃𝐿 + 𝛥𝑃𝑁𝐿
𝑑𝐵(𝐴) = 𝑃𝑁𝐿 − 14 𝑑𝐵
𝑉(𝐿/𝐷) 𝑊1
𝑅=
ln
𝑔 ∙ 𝑆𝐹𝐶 𝑊2
𝐹𝑙𝑜𝑛𝑔𝑖𝑡𝑢𝑑𝑖𝑛𝑎𝑙 = −𝐹𝑑𝑟𝑎𝑔 − 𝐹𝑓𝑟𝑖𝑐𝑡𝑖𝑜𝑛
For Tip Mach of 0.5:
For Tip Mach of 0.6:
Equation 31
Equation 32
Equation 33
Equation 34
Equation 35
where:
Equation 36
Equation 37
Equation 38
1 2
𝜌𝑉 𝑆𝐶
2 𝑎 𝐷
𝐹𝑓𝑟𝑖𝑐𝑡𝑖𝑜𝑛 = 𝐹𝑚𝑎𝑖𝑛,𝑓𝑟𝑖𝑐𝑡𝑖𝑜𝑛 + 𝐹𝑡𝑎𝑖𝑙,𝑓𝑟𝑖𝑐𝑡𝑖𝑜𝑛
1
𝐹𝑚𝑎𝑖𝑛,𝑓𝑟𝑖𝑐𝑡𝑖𝑜𝑛 = 𝜇𝑚𝑎𝑖𝑛 (𝑊𝑚𝑎𝑖𝑛 − 𝜌𝑉𝑎2 𝑆𝐶𝐿 )
2
1 2
𝐹𝑡𝑎𝑖𝑙,𝑓𝑟𝑖𝑐𝑡𝑖𝑜𝑛 = 𝜇𝑡𝑎𝑖𝑙 (𝑊𝑡𝑎𝑖𝑙 − 𝜌𝑉𝑎 𝑆(𝐶𝐿 − 𝐶𝐿0 ))
2
2
𝑑 𝑥 𝐹𝑑𝑟𝑎𝑔
𝑎= 2=
𝑑𝑡
𝑚
2
2
𝑉1 = 𝑉0 + 2𝑎𝑠
𝐹𝑑𝑟𝑎𝑔 =
∞
Equation 39
𝐹 = 𝑃𝑟[𝑋 > 𝑌] = ∬ ℎ𝑌 (𝑦)𝑔𝑋 (𝑥)𝑑𝑥𝑑𝑦
𝑦
1 [hp] = 550 [lbs-ft/sec];
1 [ft/sec] = 0.5924838 [kt];
122
1 [n.m.] = 6076.131 [ft]
Appendix B – Initial Concept Drawings
Concept A1
123
Concept A2
124
Concept A3
125
Concept A4
126
Concept A5
127
Concept B1
128
Concept B2
129
Concept B3
130
Concept B4
131
Concept B5
132
Appendix C – Decision Matrix Group A
Design Matrix
Wing
Wt.
Concept A1
Concept A2
Concept A3
Concept A4
Wt.
Score
Wt.
Score
Wt.
Score
Wt.
Score
Score
Score
Score
Score
Concept A5
Score
Wt.
Score
Wing Loading
15
3.2
48.0
3.2
48.0
3.8
57.0
3.1
46.5
3.4
51.0
Structural Implications
15
3.4
51.0
3.4
51.0
3.0
45.0
2.8
42.0
3.2
48.0
Aspect Ratio
15
3.8
57.0
2.4
36.0
3.8
57.0
3.6
54.0
3.4
51.0
Wing Position
10
3.4
34.0
3.0
30.0
3.1
31.0
4.0
40.0
3.6
36.0
Weight
10
2.8
28.0
2.7
27.0
2.8
28.0
3.8
38.0
3.2
32.0
Reliability
10
3.2
32.0
3.2
32.0
3.0
30.0
3.2
32.0
3.2
32.0
Maintenance
10
3.4
34.0
3.0
30.0
3.2
32.0
3.4
34.0
3.4
34.0
Integration
10
3.4
34.0
3.2
32.0
3.6
36.0
3.2
32.0
3.4
34.0
5
3.0
15.0
3.4
17.0
3.4
17.0
3.8
19.0
3.4
17.0
284.
0
Ease of Manufacture
SUBTOTAL
CATEGORY WEIGHT
0.175
285.0
49.9
255.0
44.6
276.0
48.3
291.0
50.9
49.7
Vertical Stabilizer
Impact on Stability/Control
25
3.6
90.0
2.9
72.5
3.2
80.0
3.8
95.0
3.2
80.0
Aerodynamic Interference
15
3.4
51.0
2.8
42.0
3.8
57.0
3.4
51.0
4.0
60.0
Reliability
15
3.2
48.0
2.6
39.0
3.2
48.0
3.4
51.0
2.7
40.5
Structural Implications
15
3.4
51.0
3.0
45.0
3.0
45.0
3.4
51.0
2.5
37.5
Ease of Manufacture
10
3.2
32.0
3.2
32.0
2.8
28.0
3.6
36.0
2.5
25.0
Integration
10
2.8
28.0
3.0
30.0
2.6
26.0
3.4
34.0
3.0
30.0
Maintenance
5
3.0
15.0
3.0
15.0
2.8
14.0
3.2
16.0
2.8
14.0
Weight
5
3.0
15.0
2.8
14.0
3.0
15.0
3.2
16.0
3.4
17.0
304.
0
SUBTOTAL
0.05
16.5
289.5
14.5
313.0
15.7
350.0
17.5
15.2
Tail
CATEGORY WEIGHT
330.0
Horizontal Stabilizer
Impact on Stability/Control
25
3.8
95.0
2.6
65.0
3.6
90.0
3.8
95.0
3.4
85.0
Aerodynamic Interference
15
3.2
48.0
2.6
39.0
3.8
57.0
3.2
48.0
4.0
60.0
Reliability
15
3.4
51.0
2.8
42.0
3.4
51.0
3.4
51.0
2.7
40.5
Structural Implications
15
3.2
48.0
2.6
39.0
3.2
48.0
3.4
51.0
2.5
37.5
Ease of Manufacture
10
3.4
34.0
3.2
32.0
3.0
30.0
3.6
36.0
2.5
25.0
Integration
10
3.0
30.0
3.0
30.0
2.4
24.0
3.4
34.0
3.0
30.0
Maintenance
5
3.0
15.0
3.0
15.0
3.0
15.0
3.2
16.0
2.8
14.0
Weight
5
3.0
15.0
3.0
15.0
3.0
15.0
3.2
16.0
3.2
16.0
308.
0
SUBTOTAL
CATEGORY WEIGHT
0.05
336.0
16.8
277.0
13.9
133
330.0
16.5
347.0
17.4
15.4
Fuselage
Weight
20
3.0
60.0
2.6
52.0
3.4
68.0
3.2
64.0
3.0
60.0
Integration
15
3.2
48.0
3.2
48.0
3.2
48.0
3.0
45.0
3.4
51.0
Structural Implications
15
2.6
39.0
3.0
45.0
3.2
48.0
2.6
39.0
3.2
48.0
Ease of Manufacture
10
3.4
34.0
2.8
28.0
3.4
34.0
2.8
28.0
3.6
36.0
Maintenance
10
3.0
30.0
3.0
30.0
3.2
32.0
2.8
28.0
3.2
32.0
Reliability
10
2.6
26.0
3.0
30.0
2.8
28.0
2.6
26.0
3.2
32.0
Size
10
3.2
32.0
2.5
25.0
3.4
34.0
2.6
26.0
2.8
28.0
Aesthetic
5
2.8
14.0
2.6
13.0
3.2
16.0
3.0
15.0
3.0
15.0
Drag
5
3.2
16.0
3.0
15.0
3.4
17.0
4.0
20.0
2.8
14.0
256.
0
SUBTOTAL
Propulsion
CATEGORY WEIGHT
0.1
299.0
29.9
23.4
257.0
25.7
227.0
22.7
25.6
Reliability
20
3.2
64.0
2.7
54.0
3.6
72.0
3.0
60.0
3.0
60.0
Maintenance
15
3.2
48.0
3.0
45.0
2.8
42.0
3.0
45.0
3.0
45.0
Noise
15
3.0
45.0
3.1
46.5
3.0
45.0
3.0
45.0
3.4
51.0
Fuel Consumption
10
3.4
34.0
3.2
32.0
3.8
38.0
3.8
38.0
3.4
34.0
Integration
10
3.2
32.0
2.4
24.0
3.2
32.0
3.4
34.0
3.4
34.0
Position
10
3.2
32.0
2.6
26.0
3.2
32.0
3.2
32.0
3.2
32.0
Power Output
10
3.6
36.0
3.0
30.0
4.0
40.0
4.4
44.0
3.8
38.0
Weight
10
2.8
28.0
3.6
36.0
3.2
32.0
4.2
42.0
3.7
37.0
271.
0
SUBTOTAL
CATEGORY WEIGHT
Landing Gear
234.0
0.2
255.0
51.0
239.5
47.9
261.0
52.2
280.0
56.0
54.2
Braking
15
3.4
51.0
3.0
45.0
3.4
51.0
3.8
57.0
3.2
48.0
Ground Handling
15
4.2
63.0
3.0
45.0
3.2
48.0
2.2
33.0
4.0
60.0
Structural Implications
15
3.4
51.0
3.0
45.0
3.0
45.0
2.6
39.0
3.0
45.0
Ease of Manufacture
10
2.9
29.0
3.0
30.0
3.2
32.0
2.8
28.0
2.8
28.0
Ground Clearance
10
3.6
36.0
2.9
29.0
2.8
28.0
3.4
34.0
3.0
30.0
Integration
10
3.2
32.0
3.0
30.0
3.0
30.0
2.4
24.0
3.0
30.0
Maintenance
10
3.0
30.0
3.0
30.0
3.0
30.0
3.0
30.0
3.0
30.0
Reliability
10
3.0
30.0
3.0
30.0
3.0
30.0
3.0
30.0
3.1
31.0
5
3.0
15.0
3.0
15.0
3.0
15.0
3.0
15.0
3.0
15.0
269.
0
Weight
SUBTOTAL
CATEGORY WEIGHT
0.1
286.0
28.6
254.0
25.4
134
258.0
25.8
233.0
23.3
26.9
Performance
Endurance
20
4.0
80.0
3.4
68.0
3.8
76.0
4.0
80.0
3.8
76.0
Fuel Weight
15
3.4
51.0
2.4
36.0
3.6
54.0
4.0
60.0
3.6
54.0
Glide Range
15
3.8
57.0
2.9
43.5
3.4
51.0
3.2
48.0
3.2
48.0
Ceiling
10
3.6
36.0
3.2
32.0
3.4
34.0
3.4
34.0
3.2
32.0
Landing
10
3.2
32.0
2.8
28.0
3.2
32.0
2.8
28.0
3.2
32.0
Maximum Velocity
10
3.2
32.0
3.2
32.0
3.6
36.0
3.6
36.0
3.2
32.0
Range
10
3.8
38.0
4.0
40.0
3.2
32.0
3.6
36.0
3.4
34.0
Velocity at Stall
10
3.0
30.0
3.0
30.0
2.8
28.0
3.4
34.0
3.0
30.0
338.
0
SUBTOTAL
Payload
CATEGORY WEIGHT
62.3
309.5
54.2
343.0
60.0
356.0
62.3
59.2
Location
50
3.4
170
3.0
150
3.8
190
2.8
140
3.6
180
Visibility
50
3.2
160
3.0
150
3.8
190
2.4
120
3.6
180
SUBTOTAL
CATEGORY WEIGHT
Overall
0.175
356.0
0.05
330.0
16.5
300.0
15.0
380
19.0
260
13.0
360
18.0
CG/Weight Distribution
25
3.4
85.0
2.5
62.5
3.8
95.0
3.6
90.0
3.3
82.5
Feasibility
20
3.2
64.0
2.9
58.0
2.8
56.0
3.8
76.0
3.5
70.0
Storage/Portability
20
3.4
68.0
2.6
52.0
2.9
58.0
4.0
80.0
3.2
64.0
Reliability
15
3.4
51.0
3.0
45.0
3.0
45.0
3.2
48.0
3.2
48.0
Aesthetics
10
3.2
32.0
3.0
30.0
2.8
28.0
3.2
32.0
2.7
27.0
Crashworthiness
10
2.6
26.0
3.0
30.0
2.6
26.0
3.0
30.0
3.0
30.0
SUBTOTAL
CATEGORY WEIGHT
Total Score
0.1
326.0
277.5
308.0
356.0
321.5
32.6
27.8
30.8
35.6
32.2
304.1
266.6
294.0
298.7
296.3
135
Wing
Weight
Appendix D – Decision Matrix, Group B
Tail
concept 5
weighted
score
weighted
score
weighted
score
weighted
score
weighted
3.92
1.176
3.25
0.975
3.75
1.125
3.67
1.101
3.17
0.951
Weight
0.20
3.25
0.65
4.25
0.85
3.17
0.634
3.42
0.684
4
0.8
Impact on Stability/Control
0.20
3.33
0.666
4
0.8
3.92
0.784
4.17
0.834
3.17
0.634
Structural Implications
0.20
3.5
0.7
3.5
0.7
3.42
0.684
3.5
0.7
3.33
0.666
Integration with Fuselage
0.05
3.58
0.179
3.75
0.1875
3.33
0.1665
3.58
0.179
3.42
0.171
Ease of Manufacture
0.05
3.75
0.1875
3.75
0.1875
2.42
0.121
3.67
0.1835
3.5
0.175
SUBTOTAL
3.5585
3.7
3.5145
3.6815
3.397
0.1
0.35585
0.37
0.35145
0.36815
0.3397
Weight
0.35
3.42
1.197
4.5
1.575
3.83
1.3405
3.33
1.1655
4
1.4
Impact on Stability/Control
0.20
3.08
0.616
3.83
0.766
4
0.8
3.42
0.684
3.67
0.734
Structural Implications
0.20
3.25
0.65
2.92
0.584
4.58
0.916
3.17
0.634
3.08
0.616
Aerodynamic Efficiency
0.15
3.17
0.4755
4.17
0.6255
3.42
0.513
3.33
0.4995
3.92
0.588
Integration with Fuselage
0.05
3.33
0.1665
3.5
0.175
3.92
0.196
2.75
0.1375
3.5
0.175
Ease of Manufacture
0.05
3.25
0.1625
3.92
0.196
3.83
0.1915
3.08
0.154
3.75
0.1875
SUBTOTAL
3.2675
3.9215
3.957
3.2745
3.7005
0.1
0.32675
0.39215
0.3957
0.32745
0.37005
Weight
0.35
2.75
0.9625
4.42
1.547
3.25
1.1375
3.5
1.225
4.08
1.428
Payload Integration
0.20
3.5
0.7
3
0.6
3.5
0.7
3.17
0.634
3.33
0.666
Structural Implications
0.20
3.75
0.75
3.58
0.716
3.92
0.784
3.58
0.716
3.42
0.684
Aerodynamic Efficiency
0.15
3.83
0.5745
3.67
0.5505
3.42
0.513
4.17
0.6255
2.83
0.4245
Ease of Manufacture
0.10
3.25
0.325
3.58
0.358
3.83
0.383
2.92
0.292
3.42
0.342
SUBTOTAL
CATEGORY WEIGHT 0.1
Propulsion
concept 4
score
0.30
CATEGORY WEIGHT
3.312
3.7715
3.5175
3.4925
3.5445
0.3312
0.37715
0.35175
0.34925
0.35445
Structural Implications
0.35
2
0.7
3.83
1.3405
4.5
1.575
3.17
1.1095
3.17
1.1095
Integration with Aircraft
0.25
2.7
0.675
3.8
0.95
4.4
1.1
3.5
0.875
3.6
0.9
Acoustics
0.20
3.1
0.62
3.3
0.66
4.2
0.84
3.2
0.64
3.2
0.64
Fuel Consumption
0.10
3.17
0.317
2.67
0.267
3.33
0.333
3
0.3
3
0.3
Weight
0.10
3.5
0.35
4.25
0.425
4.13
0.413
3.88
0.388
4.25
0.425
SUBTOTAL
2.662
3.6425
4.261
3.3125
3.3745
0.1
0.2662
0.36425
0.4261
0.33125
0.33745
CATEGORY WEIGHT
Conformity to Requirements
Concept Designs
concept 3
concept 2
Aerodynamic Efficiency
CATEGORY WEIGHT
Fuselage
concept 1
Endurance
0.35
4.25
1.4875
3.67
1.2845
4.33
1.5155
3.5
1.225
3.67
1.2845
Landing Distance
0.25
3
0.75
3.2
0.8
2.7
0.675
3.2
0.8
2.8
0.7
Weight of Payload
0.20
3.8
0.76
3.9
0.78
3.9
0.78
3.8
0.76
3.8
0.76
Speed
0.15
4
0.6
3.17
0.4755
3.92
0.588
3.33
0.4995
3.5
0.525
Acoustics
0.05
3.3
0.165
3.3
0.165
4.2
0.21
3.3
0.165
3.3
0.165
SUBTOTAL
3.7625
3.505
3.7685
3.4495
3.4345
0.15
0.564375
0.3505
0.37685
0.34495
0.34345
CATEGORY WEIGHT
136
Reliability
Glide Characteristics
0.30
4.33
1.299
3.67
1.101
3.58
1.074
4
1.2
3.08
0.924
Structural Integrity
0.40
3.9
1.56
3.5
1.4
4.2
1.68
3.5
1.4
3.9
1.56
Landing Gear
Control Surfaces (malfunction)
0.10
3.4
0.34
3.5
0.35
3.3
0.33
3.3
0.33
3.4
0.34
0.20
3.5
0.7
3.2
0.64
3.7
0.74
3.3
0.66
3.4
0.68
SUBTOTAL
3.899
3.491
3.824
3.59
3.504
0.3
1.1697
1.0473
1.1472
1.077
1.0512
Overall Aircraft
Human Factor
CATEGORY WEIGHT
Ease of Construction
0.50
3.25
1.625
3.75
Maintenance
0.50
3.08
1.54
4
1.875
3.5
1.75
3.42
1.71
4.08
2.04
2
3.67
1.835
3.17
1.585
4.17
2.085
SUBTOTAL
3.865
4.515
4.325
3.955
4.805
0.05
0.19325
0.22575
0.21625
0.19775
0.24025
CATEGORY WEIGHT
Drag
0.30
4.17
1.251
3.67
1.101
3.08
0.924
3.83
1.149
3.25
0.975
Feasibility
0.30
3.5
1.05
4
1.2
3.83
1.149
3.75
1.125
4.17
1.251
CG/Weight Distribution
0.20
3
0.6
4.33
0.866
4
0.8
3.5
0.7
3.67
0.734
Control Surfaces
0.15
3.33
0.4995
3.08
0.462
3.75
0.5625
3.5
0.525
2.92
0.438
Transportability
0.05
2.33
0.1165
3.92
0.196
3.67
0.1835
3
0.15
4.17
0.2085
SUBTOTAL
CATEGORY WEIGHT
TOTAL SCORE
0.1
3.517
3.825
0.3517
3.559025
0.3825
3.5096
137
3.619
0.3619
3.6272
3.649
0.3649
3.3607
3.6065
0.36065
3.3972
Appendix E – Constraint Analysis Script
W=300; %weight (pounds)
Z=0; %altitude
(Feet)
Vknots=50;
%velocity (knots)
V=Vknots*1.68780986;
%velocity (feet per second)
[T,R,P,A,MU,TS,RR,PP,RM,QM,KK] = stdatmf(Z,1);
dh_dtminuet=200;
%rate of climb (ft/min) at sea level
dh_dt=dh_dtminuet/60;
%rate of climb (ft/sec) at sea level
dV_dt=0;
%acceleration
e=.9;
%oswalds effciency factor
AR=7;
%aspect ratio
Cdo=.005;
%skin friction drag coefficient (guess)
Clmax=1.6; %mas lift coefficient (guess)
Vstallknots= 40 %knots
Vstall=Vstallknots*1.68780986 %fps
L=W;
k=1/(pi*AR*e);
q=.5*R*V^2;
%Cl = L/(q*S);
%lift (pounds)
%induced drag coefficient
%dynamic pressure (psf)
%strait and level flight
W_S=[.05:.002:10];
T_Wsl= (q.*Cdo)./W_S+ (k./q).*W_S;
%stall
W_Sstall=.5.*R.*Vstall.^2.*Clmax
%climb
n=3;
%load factor
T_Wclimb=(q.*Cdo)./(W_S)+(k*n^2/q).*W_S+(1/V)*dh_dt;
%Landing
W_Slanding=.5.*R.*(.8.*Vstall).^2.*Clmax;
plot(W_Sstall,T_Wsl,'--.b',W_S,T_Wclimb,W_S,T_Wsl,W_Slanding,T_Wsl,'--.g');
title('Constraint Analysis ')
xlabel('W/S');
ylabel('T/W');
legend('Crusie','Climb','Stall','Landing')
138
Appendix F – Endurance Script
%Max Endurance for a prop occurs at minimum POWER
W=300; %wieght (pounds)
Z=3000; %altitude
(Feet)
Vknots=50;
%velocity (knots)
V=Vknots*1.68780986;
%velocity (feet per second)
dh_dtminuet=200;
%rate of climb (ft/min) at sea level
dh_dt=dh_dtminuet/60;
%rate of climb (ft/sec) at sea level
dV_dt=0;
%acceleration
e=.9;
%oswalds effciency factor
AR=7;
%aspect ratio
Cdo=.005;
%skin friction drag coefficient (guess)
Clmax=1.6; %mas lift coefficient (guess)
S=53.5714;
%planform area (ft^2) (guess)
[T,R,P,A,MU,TS,RR,PP,RM,QM,KK] = stdatmf(Z,1);
L=W;
%lift (pounds)
k=1/(pi*AR*e); %induced drag coefficient
q=.5*R*V^2;
%dynamic pressure (psf)
Cl = L/(q*S);
TR=Cdo.*q.*S+2.*k.*W^2./(q.*S); %thrust required (pounds)
D=TR;
Cd=D/(q*S);
W1=300; %pounds
W2=275; %pounds
np=.8; %propulsive efficiency
gpHP=.55; %specific fuel consuption (pounds/HP.hr)
gp=gpHP*(1/550); %specific fuel consuption (pounds/(ft.lb/sec).hr)
gpsec=gp*(1/3600); %specific fuel consuption (pounds/(ft.lb/sec).sec)
E=(np/gp)*(1/V)*(Cl/Cd)*log(W1/W2)
E2=(np/gp)*sqrt(2*R*S)*Cl^(3/2)*(1/Cd)*(1/sqrt(W2)-1/sqrt(W1))
Rfeet=(np/gpsec)*(Cl/Cd)*log(W1/W2);% range feet
R=Rfeet/5280
%range miles
Rknot=R*0.868976242 %range knotical miles
139
Appendix G – Power Script
W=300; %wieght (pounds)
Z=10000; %altitude
(Feet)
Vknots=[30:.5:70];
%velocity (knots)
V=Vknots*1.68780986;
%velocity (feet per secound)
dh_dtminuet=200;
%rate of climb (ft/min) at sea level
dh_dt=dh_dtminuet/60;
%rate of climb (ft/sec) at sea level
dV_dt=0;
%acceleration
e=.9;
%oswalds effciency factor
AR=7;
%aspect ratio
Cdo=.005;
%skin friction drag coefficient (guess)
Clmax=1.6; %mas lift coefficient (guess)
S=53.5714;
%planform area (ft^2) (guess)
[T,R,P,A,MU,TS,RR,PP,RM,QM,KK] = stdatmf(Z,1)
L=W;
%lift (pounds)
k=1/(pi*AR*e); %induced drag coefficient
q=.5.*R.*V.^2;
%dynamic pressure (psf)
%Cl = L/(q*S);
TR=Cdo.*q.*S+2.*k.*W^2./(q.*S);
PR=TR.*V;
PRhp=PR./550; %horesepower
figure(1)
plot(V,PR)
title('Power Required')
xlabel('Velocity (fps) ')
ylabel('PR (ft*lb/sec)')
figure(2)
plot(Vknots,PRhp)
title('HP Required')
xlabel('Velocity (knots) ')
ylabel('PR (HP)')
140
Appendix H – Constraint Analysis Script
Author: Mike Sherman
clear
clc
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
%%%%% Design Constraints %%%%%%
UNITS
%%%%
Vcruise=50;
%%%
knotts
%%%%
Vmax=70;
%%%
knotts
%%%%
MTOGW= 200;
%%%
lbs
%%%%
Range= 15;
%%%
natuical miles %%%%
Endurance= 8;
%%%
hours
%%%%
Ceiling= 10000;
%%%
ft
%%%%
Op_Altitude=3000;
%%%
ft
%%%%
Altitude or 2000 ft AGL)
Turn_Rate= 6;
%%%
degrees/second %%%%
Climb_Rate= 200;
%%%
ft/min
%%%%
Payload= 30;
%%%
lbs
%%%%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
(@half fuel)
(Operational
(@ Sea Level)
%%%%%%%%%%%%%% Converts Units to English %%%%%%%%%%%%%%%%%%
Vcruise=Vcruise*1.6878098;
%%%%
ft/s %%%%
Vmax=Vmax*1.6878098;
%%%%
ft/s %%%%
Range=Range*6076.11548556;
%%%%
ft
%%%%
Climb_Rate=Climb_Rate/60;
%%%%
ft/s %%%%
Endurance=Endurance*3600;
%%%%
s
%%%%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%%%%% Planform Esitmation %%%%%%%%%%%%%%%%%%%
%%%%%%%%%% Assumptions:
Cl Cruise = 0.3
%%%%%%%%
%%%%%%%%%% rho @ sea level = 0.002378
slugs/ft %%%%%%%%
%%%%%%%%%% rho @ 10,000 ft = 0.001755
slugs/ft %%%%%%%%
%%%%%%%%%% Lift = Weight
%%%%%%%%
%%%%%%%%%% Lift = Cl*1/2*rho^2*S
%%%%%%%%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
Cl_cruise = 0.3;
%
Rho_sl = 0.002378;
%
Rho_10k = 0.001755;
%
S = MTOGW/(Cl_cruise*1/2*Rho_sl*Vcruise^2);
%
S_10k = MTOGW/(Cl_cruise*1/2*Rho_10k*Vcruise^2);
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%%%%%%% Drag Calculation %%%%%%%%%%%%%%%%%%%
% Drag = (CDo)(1/2*rho*V^2*S)+2*k*W^2/(rho*V^2*S)
%
% Assumptions: e = .9
CDo = 0.02
AR = 10
%
%
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
e = .9;
%
CDo = 0.02;
%
AR = 10;
%
k=1/(pi*AR*e);
%
D = (CDo)*(1/2*Rho_sl*Vcruise^2*S)+2*k*MTOGW^2/(Rho_sl*Vcruise^2*S);
141
D_10k =
(CDo)*(1/2*Rho_10k*Vcruise^2*S)+2*k*MTOGW^2/(Rho_10k*Vcruise^2*S_10k);
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%%%%%%% Power Required %%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%% Straight and Level Flight %%%%%%%%%%%%%%%
%
%
% Power Required = (CDo)(1/2*rho*V^3*S)+2*k*W^2/(rho*V*S) %
% Assumptions: e = .9
CDo = 0.06
AR = 10
%
%
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
P = (CDo)*(1/2*Rho_sl*Vcruise^3*S)+2*k*MTOGW^2/(Rho_sl*Vcruise*S);
P_10k = (CDo)*(1/2*Rho_10k*Vcruise^3*S)+2*k*MTOGW^2/(Rho_10k*Vcruise*S_10k);
P_hp=P/550;
%%%% Converts to Horse Power%
P_10k_hp=P_10k/550;
%%%% Converts to Horse Power%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%%%%%%% Power Required %%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%%%%%% Climb 200 ft/min
%%%%%%%%%%%%%%%%%%
%
%
% Rate of Climb = (Power Available - Power Required)/W
%
%
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
Pclimb=Climb_Rate*MTOGW+P;
%
Pclimb_hp=Pclimb/550;
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%%%% Max Endurance %%%%%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%% Constant Speed
%%%%%%%%%%%%%%%%%%%%%%%%
%
%
% Range = (Prop_eff/Fuel_consump)(L/D)ln(W1/W2)
%
%
%
%Assumptions:108D2 enginge Fuel=1.75 gal/hour (10+ HP)
%
% Prop_eff = .85;
Avgas= 6.02 lb/gallon
%
% L/D at min drag conditions
%
% Aircraft composed of 50% fuel
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
Prop_eff=.85;
%
Fuel_consump=1.75*6.02/3600/12/550;
%
L_D_minDrag=1/(2*(CDo*k)^(1/2));
%
Wfuel=MTOGWMTOGW/exp((Endurance*Vcruise*Fuel_consump/Prop_eff/(L_D_minDrag)));
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%%%% Constraint Analysis %%%%%%%%%%%%%%%%%%
%
%
%
%
%%%%%%%%%%%%%%%%%
Climb
%%%%%%%%%%%%%%%%%%%%%%%%
% Assumptions: load factor, n = 4
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
W_S=[.05:.002:5];
%
142
n=4;
%
q_10k=(1/2*Rho_10k*Vcruise^2);
%
T_W=(q_10k*CDo)./W_S+(k*n^2/q_10k).*W_S+(1/Vcruise)*Climb_Rate;
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%%
Stall
%%%%%%%%%%%%%%%%%%%%%%%%
% Assumptions: Vstall =
35 knotts
Cl max = 1.3
%
%
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
Vstall=35;
%
Cl_max=1.3;
%
%%%%%%%%%%%%%% Converts Units to English %%%%%%%%%%%%%%%%%
Vstall=Vstall*1.6878098;
%%%%
ft/s %%%
%
W_S_stall=0.5*Rho_10k*Vstall^2*Cl_max;
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%%
Landing
%%%%%%%%%%%%%%%%%%%%%%%
% Assumptions: V_landing = 1.2(V_stall)
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
W_S_landing=.5*Rho_sl*(1/1.2*Vstall)^2*Cl_max;
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
plot(W_S,T_W,'k--')
%
hold on
%
line([W_S_stall W_S_stall],[min(T_W) max(T_W)],'Color','g')
line([W_S_landing W_S_landing],[min(T_W) max(T_W)],'Color','b')
xlabel('W/S');
%
ylabel('T/W');
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
%%%%%%%%%%%%%%%%%
PLotting
%%%%%%%%%%%%%%%%%%%%%%%
text(1.05*W_S_stall,max(T_W),'Stall')
%
text(.8*W_S_landing,min(T_W),'Landing')
%
text(0.5,2,'Climb')
%
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
143
Appendix I – Final Design CAD 3-View
144
Appendix J – Stability Derivatives and Crosswind Script
Wref
rhoref
u0ref
theta0ref
Sref
bref
cref
=
=
=
=
244.95;
0.002378;
84.39049;
0;
= 76;
= 21.5;
= 3.487522;
I = [56.5019144,
93.30773659,
142.9306743,
-3.798490151];
%
%
%
%
Ixx
Iyy
Izz
Ixz
CYbeta = -0.187314;
Clbeta = -0.046917;
Cnbeta = 0.033842;
CYp = 0.102286;
Clp = -0.454969;
Cnp = -0.035389;
Clr = 0.141929;
Cnr = -0.065491;
CYr = 0.106587;
CLalpha = 4.798984;
CDalpha = 0.244;
Cmalpha = -0.661709;
CLalphadot = 0;
Cmalphadot = -3.7;
CLq = 7.691370;
Cmq = -12.739847;
CLM = 0;
CDM = 0;
CmM = 0;
CLdeltae = 0.005487*(180/pi);
Cmdeltae = -0.015113*(180/pi);
CYdeltaa
Cldeltaa
Cndeltaa
CYdeltar
Cldeltar
Cndeltar
=
=
=
=
=
=
-0.000247*(180/pi);
0.006396*(180/pi);
0.000286*(180/pi);
0.001902*(180/pi);
0.000089*(180/pi);
-0.000894*(180/pi);
% Cross wind landing analysis.
% Philip Pesce
% 03/29/08
ref
CASE
= 1;
% CASE 1 deflections for sideslip
% CASE 2 max crosswind
145
dr_max
dr_max
= 25;
= dr_max*(pi/180);
% Max rudder deflection [deg]
% Max rudder deflection [rad]
W_lbs
S
= Wref;
= Sref;
% Weight [lbs]
% Planform Area [ft^2]
rho
= rhoref;
% Density of Air SL [sl/ft^3]
vknot
Vknot
V
v
=
=
=
=
% Crosswind velocity [kts]
B
B_deg
= asin(v/V);
= B*(180/pi);
CW
= 300/((1/2)*rho*V^2*S);
CYB
CYda
CYdr
= CYbeta;
= CYdeltaa;
= CYdeltar;
ClB
Clda
Cldr
= Clbeta;
= Cldeltaa;
= Cldeltar;
CnB
Cnda
Cndr
= Cnbeta;
= Cndeltaa;
= Cndeltar;
10;
35.4;
1.68781*Vknot;
1.68781*vknot;
% Total airspeed [ft/s]
% Crosswind velocity [ft/s]
% Sideslip angle [rad]
% Sideslip angle [deg]
switch CASE
case {1}
Mat1 = [CYda, CYdr, CW;
Clda, Cldr, 0;
Cnda, Cndr, 0];
Mat2 = [CYB; ClB; CnB];
sol = (inv(Mat1))*(-Mat2*B);
da = (180/pi)*sol(1)
dr = (180/pi)*sol(2)
phi = (180/pi)*sol(3)
% Aileron deflection [deg]
% Rudder deflection [deg]
% Bank angle [deg]
case{2}
Mat1 = [CYB, CYda, CW;
ClB, Clda, 0;
CnB, Cnda, 0];
Mat2 = -[CYdr; Cldr; Cndr];
sol = inv(Mat1)*(Mat2*dr_max);
B
= (180/pi)*sol(1)
da = (180/pi)*sol(2)
phi = (180/pi)*sol(3)
% Sideslip angle [deg]
% Aileron deflection [deg]
% Bank angle [deg]
max_v = sind(B)*V*0.5924838 % Max crosswind component [kt]
end
146
Appendix K – Specific Fuel Consumption Script
close all;
clear all;
format long;
RPM
air_min
fuel_min
fuel_hour
ISHP
SFC_Roskam
SFC_est
=
=
=
=
=
=
=
+
+
scale
=
SFC_scaled =
ISHP_scaled =
ISHP_est
=
500:100:8500;
%[rpm]
.00016.*RPM;
%[m^3/min]
.000136.*air_min;
%[m^3/min]
60.*264.172.*6.*fuel_min;
%[lbf/hr]
(202.79.*9.76379906.*RPM)./(792000); %[hp]
fuel_hour./ISHP;
%[lb/(hp*hr)]
1.6085E-16.*RPM.^4 - 1.3328E-12.*RPM.^3 ...
9.1131E-09.*RPM.^2 - 4.0546E-05.*RPM ...
7.0736E-01;
%[lb/(hp*hr)]
SFC_Roskam(1)/SFC_est(56);
%[const]
SFC_est.*scale;
%[lb/(hp*hr)]
fuel_hour./SFC_scaled;
%[hp]
fuel_hour./SFC_est;
%[hp]
figure;
plot(RPM, ISHP, RPM, ISHP_est, RPM, ISHP_scaled);
legend('ISHP_R_o_s_k_a_m','ISHP_e_s_t','ISHP_e_s_t _s_c_a_l_e_d', 2);
title('Power vs. RPM');
xlabel('RPM');
ylabel('Power [hp]');
grid on;
figure;
plot(RPM, SFC_Roskam, RPM, SFC_est, RPM, SFC_scaled);
legend('SFC_R_o_s_k_a_m','SFC_e_s_t', 'SFC_s_c_a_l_e_d', 0);
title('SFC vs. RPM');
xlabel('RPM');
ylabel('SFC [lb/(hp*hr)]');
grid on;
147
Appendix L – Internal Combustion Engine Propeller
148
Appendix M – Electric Engine Propeller
149
Appendix N – Electric Propulsion Battery Requirements
Table 52. Number of Batteries vs. Max Headwind
# in
Weight Weight
mAH
Parallel
(oz)
(lbs)
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26
27
28
29
30
31
32
33
34
35
36
37
38
39
40
6600
13200
19800
26400
33000
39600
46200
52800
59400
66000
72600
79200
85800
92400
99000
105600
112200
118800
125400
132000
138600
145200
151800
158400
165000
171600
178200
184800
191400
198000
204600
211200
217800
224400
231000
237600
244200
250800
257400
264000
72
144
216
288
360
432
504
576
648
720
792
864
936
1008
1080
1152
1224
1296
1368
1440
1512
1584
1656
1728
1800
1872
1944
2016
2088
2160
2232
2304
2376
2448
2520
2592
2664
2736
2808
2880
4.5
9
13.5
18
22.5
27
31.5
36
40.5
45
49.5
54
58.5
63
67.5
72
76.5
81
85.5
90
94.5
99
103.5
108
112.5
117
121.5
126
130.5
135
139.5
144
148.5
153
157.5
162
166.5
171
175.5
180
150
Max Headwind
(no glide)
Max Headwind
(4n.m. glide)
-211.3636364
-80.68181818
-37.12121212
-15.34090909
-2.272727273
6.439393939
12.66233766
17.32954545
20.95959596
23.86363636
26.23966942
28.21969697
29.8951049
31.33116883
32.57575758
33.66477273
34.62566845
35.47979798
36.24401914
36.93181818
37.55411255
38.11983471
38.63636364
39.10984848
39.54545455
39.94755245
40.31986532
40.66558442
40.98746082
41.28787879
41.56891496
41.83238636
42.07988981
42.31283422
42.53246753
42.73989899
42.93611794
43.12200957
43.2983683
43.46590909
-141.6666667
-45.83333333
-13.88888889
2.083333333
11.66666667
18.05555556
22.61904762
26.04166667
28.7037037
30.83333333
32.57575758
34.02777778
35.25641026
36.30952381
37.22222222
38.02083333
38.7254902
39.35185185
39.9122807
40.41666667
40.87301587
41.28787879
41.66666667
42.01388889
42.33333333
42.62820513
42.90123457
43.1547619
43.3908046
43.61111111
43.8172043
44.01041667
44.19191919
44.3627451
44.52380952
44.67592593
44.81981982
44.95614035
45.08547009
45.20833333
Appendix O – Noise Footprint Output from Xrotor
Figure 47. Noise Footprint at Ground of Aircraft at 200 ft AGL at 4500 RPM with 2 Blades
Figure 48. Noise Footprint at Ground for Aircraft at 2000 ft AGL at 4500 RPM with 2 Blades
151
Figure 49. Noise Footprint at Ground for Aircraft at 3000 ft AGL at 4500 RPM with 2 Blades
Figure 50. Noise Footprint at Ground for Aircraft at 200 ft AGL at 4500 RPM with 3 Blades
152
Figure 51. Noise Footprint at Ground for Aircraft at 2000 ft AGL at 4500 RPM with 3 Blades
Figure 52. Noise Footprint at Ground for Aircraft at 3000 ft AGL at 4500 RPM with 3 Blades
153
Figure 53. Noise Footprint at Ground for Aircraft at 200 ft AGL at 6000 RPM with 2 Blades
Figure 54. Noise Footprint at Ground of Aircraft at 2000 ft AGL at 6000 RPM with 2 Blades
154
Figure 55. Noise Footprint at Ground of Aircraft at 3000 ft AGL at 6000 RPM with 2 Blades
Figure 56. Noise Footprint at Ground for Aircraft at 200 ft AGL at 6000 RPM with 3 Blades
155
Figure 57. Noise Footprint at Ground for Aircraft at 2000 ft AGL at 6000 RPM with 3 Blades
Figure 58. Noise Footprint at Ground for Aircraft at 3000 ft AGL for 6000 RPM with 3 Blades
156
Appendix P – Landing Matlab Script
function LandingRoll
% LandingRoll.m outputs distance from touchdown to stop
% Plots distance and velocity vs. time
clear all
close all
clc
warning off
global mu weight alpha S rho gravity AR oswalds Cd0 CL_0 mainWheelRatio Stail
muT CLmax
% ----------- INPUTS -----------------weight=245;
%lbs
mu=0.4; % Main gear rolling friction coefficient. For brakes, use 0.3 to
0.5)_
muT=0.8; %For tricycle, muT=0. For taildragger, muT=0.8 for rubber pad on
concrete
S=76.08;
%ft^2
rho=0.002378;
%slugs/ft^3
gravity=32.2;
%ft^2/s
AR=6.25;
oswalds=0.861;
Cd0=0.0257;
%parasitic drag coefficient
CLmax=1.48;
CL_0=0.3;
mainWheelRatio=.85;
% Percent of weight on main wheels.
Stail=14.3;
% horizontal tail area in ft^2
plotON=1;
% plotON=1 for plotting turned on;
turned off
%======================================
Vstall=sqrt(2*weight/rho/S/CLmax);
V0=1.15*Vstall; %Horizontal Touchdown Speed (ft/s);
%----------------------------timeNeg=500;
stop=0;
for accuracy=1:2
[t,V]=ode45(@rigid,[0,timeNeg],[0;V0]);
for i=1:length(V)
if V(i,2)>0
Vplot(i,1)=V(i,1);
Vplot(i,2)=V(i,2);
Tplot(i)=t(i);
end
if V(i,2)<0
if stop==0
timeNeg=t(i);
stop=1;
end
end
157
plotOn=0 for plotting
end
end
if plotON==1
subplot(2,1,1),plot(Tplot,Vplot(:,1))
ylabel('Distance X, ft');
xlabel('Time t, seconds')
subplot(2,1,2),plot(Tplot,Vplot(:,2))
ylabel('Velocity V, ft/s');
xlabel('Time t, seconds')
end
fprintf('\nTouchdown Velocity: %.1f kts (%.1f ft/s)\n',V0*.5925,V0)
fprintf('\nPredicted Time to Stop: %0.1f seconds\n',timeNeg)
fprintf('\nPredicted Landing Roll: %0.1f feet\n\n',max(V(:,1)))
return
%%%%%%%%
% ---------------------------------------function g = rigid(t,y)
global mu weight alpha S rho gravity AR oswalds Cd0 CL_0 mainWheelRatio Stail
muT CLmax
%----------------------------------------mass=weight/gravity;
%Cl=5*alpha*pi/180+CL_0;
% assumes slope of lift curve for 3D model is 5
V=1.15*sqrt(2*weight/rho/S/CLmax);
Cl=weight/.5/rho/V^2/S;
%alpha= (Cl*(180+CL_0))/(5*pi);
Cd=Cl^2/pi/AR/oswalds+Cd0;
q=-0.5*rho*S;
g = [y(2); q*y(2).^2*Cd/mass-mainWheelRatio*weight*mu/mass+mu/mass*(q*y(2).^2*Cl)-muT/mass*((1-mainWheelRatio)*weight-.5*rho*y(2).^2*Stail*(ClCL_0))];
return
158
Figure 59. Mechanical Failure Method
Appendix Q – Mechanical Failure Model Method
159
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161