PDR
Deployable Small UAV Explorer (D–SUAVE)
Customer: Kamran Mohseni
University of Colorado at Boulder
Fall 2006
D-SUAVE Team
1.
2.
3.
4.
5.
6.
7.
8.
David Goluskin*
Michael Lapp*
Burhan Muzaffar*
Jastesh Sud*
Brandon Bobian
Nathan Sheiko
Yoshi Hasegawa
Miranda Mesloh
* Presenting
2
Overview
Objective
Mission Overview
Deployment
Aerodynamics
Structures
Propulsion
Electronics
Project Plan
Major Risk Assessments
Summary
References
Appendices
3
Primary Objective
“To design, fabricate, integrate and
verify a RC controlled UAV capable
of being remotely deployed from the
ARES aircraft and flying a specific
flight pattern.”
4
Feedback

1.
2.
3.
1.
2.
3.
4.
5.
6.
7.
8.
9.
Primary Factors
Requirement flow down
Stability of airplane during deployment
Deployment scenarios affect the configuration selected
Where do we deploy from?
Secondary Factors
What is the expected data transfer or collection rate?
Under what atmospheric conditions the system should be able to
operate
What supports the mass determination?
Required deployment force and acceleration the aircraft must
withstand
Why a Micro aerial vehicle?
What is the detail of the mission?
Overestimating propeller efficiencies
Sketches/pictures of the possible configurations
Who will pilot the vehicle?
Requirements Flow Diagram
Mission
Spatial
Resolution
Lap
Pattern
Turn
Radius
Spatial
Range
Temporal
Readings
Sensors
Resolution per Location
Lap
Distance
Velocity
Lap
Time
Number
Of Laps
Endurance
Carrier
Vehicle
Cruise Cruise Empty Vehicles
Stall
Velocity Drag Mass per Carrier Velocity
Sensor Box Mass,
Vehicle and
Size and Placement Package Mass
Deployment
Velocity
Data Collection Driving Factors
Carrier Vehicle Driving Factors
Requirements
6
Mission Profile
Carrier Vehicle ARES
Cruise velocity
25 m/s
Cruise drag
12 N
Empty wet mass
8.2 kg
Stall Velocity
17 m/s
Maximum
internal payload
3.2 kg
7
5 min/lap, 4 laps
600 m
100 m
Mission Path
8
Requirements
Parameter
Requirement
Motivation
Deployment
Transition to stable cruise
after being deployed at 20 m/s
ARES stall velocity is 17 m/s
Cruise Velocity
12 m/s
5 min/lap, 3600 m/lap
Endurance
25 min
5 min/lap, 4 laps, 5 min for
safe return
Turn Radius
50 m
turn 180° over 100 m
Vehicle and Package Mass
565 g
ARES performance, historical
payload mass fractions
Payload Mass
75 g
Customer-estimated mass of
self-powered payload box
containing temperature,
pressure and humidity
sensors
Payload Dimensions
Not yet determined
Customer-estimated
dimensions of payload box
Payload Location
Bottom of payload box must
be in contact with the
freestream
Atmospheric sensors must
have access to the air
9
Design Process
Lighter than
Air
Fixed Wing
Rocket
Unfolded
Piggyback
Folded in Tube
Folded
Piggyback
Low Aspect Ratio
Conventional
Canard
Helicopter
Flying Wing
Electronics/Comm/Sensors
Glider
Bi-Plane
Propulsion/Power Supply
Vehicle Design
Prop
Launchable/Packageable
Structure
Configuration
Motor
Folding or
Non Folding
Gear
Box
Speed Control
Outrunner
Wing Folding
Material
Stability
Controllability
Dihedral
Control Surfaces
Membrane
Ailerons
Rudder and
Elevator only
Link
Design Alternatives
Packagability
Endurance
Speed
Maneuverabilit
y
Mass
Balloon
Must be
deflated to fit in
tube
Limited by leaks
Controlled by
wind
0-1 m turn radius
~100 g
Airship
Must be
deflated to fit in
tube
13-16 min
typical
~10 m/s
0-1 m turn radius
100-800 g
Glider
Wings fold
2-5 min typical
~9 m/s
2-3 m turn radius
typical
~50 g
Helicopter
Blades fold
5-10 min typical
~13 m/s
0 turn radius
~300-1500 g
Fixed
Wing
Wings fold
15-20 min
typical
15-20 m/s
8-11 m turn
radius typical
~50 g
Rocket
Small size
0.3-0.5 min
typical
Very fast
75-90 m turn
radius typical
~110 g
11
Link
Cruise
Velocity
Endurance
USAFAMAV I
Rollup Wing,
74 cm Span
18 m/s
Wind Tunnel
Verification Only
USAFAMAV II
Folding Wing,
205 cm Span
18 m/s
15 min
UF Flexible
Wing MAV
11 cm span
13 m/s
Never Flown
CMAV
Folding Wing,
50 cm Span
20 m/s
Never Flown
UF Morphing
Wing MAV
61 cm Span
11 m/s
20 min
DART MAV
30 cm Span
18 m/s
Unstable
Acceptable
Packagablity
Unacceptable
Vehicle
Good
Existing Packagable MAVS
*Payload data generally unavailable
12
Deployment
Packaging Methods
Method
Complexity
Drag
Loads on Vehicle
Folding not required
Drag on vehicle
and deployment
mechanism
High aerodynamic
loads on the
vehicle’s wings
during carrier flight
Folded Piggy-back
Wing folding
mechanism required
Drag on compact
vehicle and
deployment
mechanism
Moderate
aerodynamics
loads on vehicle
during carrier flight
Folded in a tube
-Wing folding
mechanism required
-Tube and tube
deployment
mechanism required
Drag on smooth
package only
No aerodynamic
loads on the
vehicle during
carrier flight
Unfolded PiggyBack
14
Vehicle Deployment Methods
Vehicle
Deployment
Method
Parachute
Spring or
Elastic
Dynamic
Pressure
Peak
Acceleration
Complexity
Orientation
Control
Launch Risks
12-14 g’s to slow
to 21 m/s
-Front and back of
tube must open
-Parachute must
release
Parachute force
maintains
orientation after
leaving tube
-Line tangle
-Parachute could
slow vehicle too
much
2-4 g’s to slow to
21 m/s
-Back of tube
must open
-Elastic
mechanism
needed
No orientation
control after
leaving tube
Elastic snag
0.1-0.2 g’s
Front and back of
tube must open
No orientation
control after
leaving tube
Minimal
deceleration while
in the tube
15
Wing Deployment
Folding
Configuration
Required Tube
Diameter
Complexity
Stability
Side
By Side
dtube  2cmax
Simple hinges
No uneven roll
moment
cmax  dtube  2cmax
Angled pivots
or flexible
wings
Uneven roll
moment during
deployment
Overlapped
Wing
Deployment
Mechanism
Deployment
Rate
Complexity
Mass
Stability
Mechanism may be
added to guarantee
symmetrical
deployment
Symmetrical
deployment
guaranteed
Elastic or
Spring-Driven
Dual Pivot
Exert largest
structural loads
Simple
mechanism
Elastic bands,
mounting
hardware, release
mechanism
Servo
Dual Pivot
Servos are not
designed for high
speed
Electronics and
servo linkages
required
Servo is heavier
than a release
mechanisms
Carpenter Tape
Hinges
Provide large
unfolding
moments
Hinges may
need to be made
from composites
Hinges, release
mechanism
Symmetrical
deployment not
guaranteed 16
Tube Drop Verification
Verification Test
Manufacturing
Complexity
Logistical
Complexity
Similarity to
Actual
Drop
Risk of Crash
Low ground
clearance
Car Pole Drop
Build pole system
and attach to car
Find test location
Speed can be
matched, but
ground clearance
is lower
Building Launch
Build launch
system
Find building for
launch
Speed may be
difficult to
reproduce
Moderate ground
clearance
UAV Drop
Integrate tube onto
UAV
Coordinate with
UAV pilot
Identical
High ground
clearance
17
Car/Pole Verification

Vcar = VUAV

Pole height dictated by
drop distance
18
Package Configuration
19
Aerodynamics
Possible Configurations
Low-Aspect Ratio Aircraft
0.1
0.05
3.5
L over D
Wing z-coordinate
4
0
-0.05
-0.1
3
0.1
0
0.2
-0.1
Wing y-coordinate
0.1
0
Assumptions:
1.
Small angle of attack
2.
Thin airfoil
3.
Estimated Parasite drag
Wing x-coordinate
2.5
0
5
10
15
alpha (deg)
21
Possible Configurations
Flying Wing
8
0.1
7
0
6.5
L over D
Wing z-coordinate
7.5
-0.1
0.4
6
5.5
0.2
5
0
0.4
-0.2
Wing y-coordinate
-0.4
Assumptions:
1.
Small angle of attack
2.
Thin airfoil
3.
Estimated Parasite drag
0.2
0
Wing x-coordinate
4.5
4
0
5
10
15
alpha (deg)
22
Possible Configurations
Conventional Aircraft
10
0.1
8
0
L over D
Wing z-coordinate
9
-0.1
0.4
7
6
5
0.2
0
4
-0.2
Wing y-coordinate
-0.4
Assumptions:
1.
Small angle of attack
2.
Thin airfoil
3.
Estimated Parasite drag
0.2
0 Wing x-coordinate
3
0
5
10
15
alpha (deg)
23
Possible Configurations
Canard Aircraft
7.5
6.5
6
0.1
L over D
Wing z-coordinate
7
0
-0.1
0.4
5.5
5
4.5
0.2
4
0
0.4
0.2
-0.2
Wing y-coordinate
-0.4
0
-0.2
Assumptions:
1.
Small angle of attack
2.
Thin airfoil
3.
Estimated Parasite drag
Wing x-coordinate
3.5
3
0
5
10
15
alpha (deg)
24
Possible Configurations
Biplane
8
0.1
7
0
L over D
Wing z-coordinate
9
-0.1
0.4
6
5
0.2
4
0
-0.2
Wing y-coordinate
3
-0.4
0.2
0 Wing x-coordinate
2
0
5
10
15
alpha (deg)
Assumptions:
1.
Small angle of attack
2.
Thin airfoil
3.
Estimated Parasite drag
25
Configuration of Aircraft
Configuration
Packagability
Efficiency
Stability
Manufacturability
Experience
Low-AR
Wings must
bend
High induced
drag
Small moment
arm of control
surfaces
Membrane wing is
simple to construct
No experience
Flying Wing
Wings must
overlap when
they fold
All surfaces
generate lift
Difficult to
achieve pitch
stability
All components must
be inside wing
structure
Flying wing lab
in 3111
Conventional
Wing can sweep
back in multiple
configuration
All surfaces
produce
parasite drag,
but not all
produce lift
-Large
moment arm
for tail
-Flexible CG
placement
Must construct wing,
fuselage, tail and
control surfaces
Design and
modeling
experience in
many classes
Canard
-Longer package
if wings fold back
-Unstable
transition if
wings fold
forward
All surfaces
produce
parasite drag,
but not all
produce lift
Biplane
Small planform
area, but greater
height
Greatest
parasite drag
Similar to
conventional
Similar to
conventional
Must construct wing,
fuselage, canard,
and control surfaces
Must construct two
wings
No experience,
but similar to
conventional
configuration
No experience,
but similar to
conventional
configuration
26
Wing Sizing
Calculate span required to lift vehicle for a given chord:
Guess b
c
Lifting Line
CL
S
no
L
b
12
Span Required for Lift
Maximum Span Allowed by Tube
11
10
14.1 cm
9
AR
Span (cm)
yes
Aspect Ratio for Varying Chords
Span Required and Maximum Span for Varying Chords
150
8.6 cm
L=W ?
100
8
7
6
50
10
12
14
Chord (cm)
16
18
5
9
10
11
12
Chord (cm)
13
14
27
L/D Without Induced Drag
Airfoil Data
Lifting Line Approximation
40
40
35
30
30
20
Aquila
Verbitsky BE50
Clark Y
Cody RobertsonCR001
10
EpplerE387
0
0
2
4
6
8
Angle of Attack [ - deg]
20
15
10
5
0
S 3014
-2
L/D
L/D
25
10
-5
-2
0
2
4
6
8
Angle of Attack [ -deg]
10
28
L/D With Induced Drag
Airfoil Data Corrected for Di
(e = 0.6, AR = 8)
Lifting Line Approximation
15
15
10
10
L/D
L/D
5
5
0
Aquila
Verbitsky BE50
Clark Y
Cody RobertsonCR001
0
EpplerE387
-5
-4
-5
S 3014
-2
0
2
4
6
Angle of Attack [ - deg]
8
10
-10
-4
-2
0
2
4
6
Angle of Attack [ -deg]
8
10
29
Required CL
CL Level Flight
0.72
CL Steady Level Turn
0.7
0.68
C
L
0.66
0.64
0.62
0.6
0.58
0.56
0.09
0.1
0.11
0.12
Chord [m]
0.13
0.14
30
Link
Tail Configurations
Configuration
Packagability
Efficiency
Stability
Experience
Moderate wetted
area
No coupling of
pitch and yaw
control
No experience,
but similar to
conventional
configuration
No experience
T-Tail
Horizontal tail is
not near center
V-Tail
Tail surface
extend up as
well as sideways
Moderate wetted
area
Coupling of pitch
and yaw control
Conventional
Horizontal tail is
near center
Moderate wetted
area
No coupling of
pitch and yaw
control
H-Tail
Vertical tails are
not near center
Small wetted
area
No coupling of
pitch and yaw
control
Computed
controllability in
ASEN 3128
No experience,
but similar to
conventional
configuration
31
Control Surface Configuration
Control Surface
Configuration
Effect on
Deploying Wing
Controllability
Complexity
Manufacturability
Aileron, Elevator
& Rudder
Servos complicate
wing deployment
Good
controllability on
all 3 rotational
axes
4 surfaces
Many parts to
construct
Aileron & Rudder
Servos complicate
wing deployment
3 surfaces
Must integrate
control mechanisms
into wing
Aileron & Elevator
Servos complicate
wing deployment
Poor yaw control
3 surfaces
Must integrate
control mechanisms
into wing
Elevator & Rudder
No servos in
wings
Poor
roll control
2 surfaces
Easy to construct
Poor pitch control
32
Structures
Spar Structure
b/2
MS
MS
α = 12.6 rad/s2
ω average = 3.15 rad/s
Carbon Fiber Tubes
(OD x ID) cm
Shear stress
computed (MPa)
Allowable Shear Stress
(MPa)
Mass (g)
(b/2)
0.318 x 0.188
TBD
65.5
3.4 g
0.584 x 0.366
TBD
65.5
11.4
1.27 x 0.366
TBD
65.5
20
34
Spar Materials
Comparison of Solid Cylider Spars
25
Aluminum
Carbon Fiber
Titanium
Balsa Composite
Mass (g)
20
15
10
5
0
2
4
6
8
10
12
14
16
18
20
Vmax (mm)
35
Wing Materials
maximum deflection (mm)
5
4
3
2024-T3 Al
7075-T6 Al
Aramid-Epoxy
S-Fiberglass-Epoxy
Carbon-Epoxy
Blue Foam
Balsa
2
1
0
0
2
4
radius of gyration (mm)
36
Propulsion
Outrunner vs Gearbox
System
Outrunner
DC Motor with
a Gearbox
Additional
Mass
Efficiency
Complexity
Availability
None
Slightly lower
motor and
propeller
efficiencies
Motor
selection
Available from
local hobby
shops
>10 g
Slightly higher
motor and
propeller
efficiencies,
but ηgb < 0.93
Gearbox
selection and
mounting
Gearheads
are largely
motor-specific
38
Propeller Location
C.G.
Wing
Deployment
Interference
Crash
Consequences
Puller
Easily achieve
stable C.G.
location
None
Could damage
prop/motor
Pusher
Difficult to
move C.G. far
enough
forward
Prop/motor
could interfere
with folded
wings
Less likely to
damage
prop/motor
Propeller
Location
39
Selected Motors
ηprop
Required
Battery
Capacity
(mAh)
Total
Mass (g)
66.70%
52.90%
517
66.4
65.46%
59.48%
470
83.5
Motor
Mass
(g)
Max
Current
(A)
Recommended
Propeller
(in x in)
Back of
Efficiency
Curve?
ηmotor
Komodo
KH220411
24
11
7x3.5
Yes
Model Motors
2208/26
45
11
7x4
Yes
Motor
Assume: ηesc = 95%
Model Motors 2208/26 Efficiency Curve
0.8
0.7
0.7
0.6
0.6
Efficiency
Efficiency
Komodo KH220411 Efficiency Curve
0.8
0.5
0.4
0.5
0.4
0.3
0.3
0.2
0.2
0.1
0
0.2
0.4
0.6
0.8
1
RPM
1.2
1.4
1.6
1.8
2
4
x 10
0.1
0
0.2
0.4
0.6
0.8
1
RPM
1.2
1.4
1.6
1.8
2
4
x40
10
Battery Types
Battery Capacity vs. Mass
350
300
Mass [grams]
250
200
150
100
50
0
0
500
1000
1500
2000
2500
3000
3500
4000
Capacity [mAh]
41
Electronic Speed Control
ESC
Mass
[g]
Max
Continuous
Current
[A]
Castle Creations
Phoenix 10
6
10
Yes
BP
7.1
10
Yes
Jeti
10
12
Yes
BEC?
42
Risk Reduction Experiment
Verify Endurance Requirement
Test entire propulsion system in the ITLL wind
tunnel
Test battery capacity
RC Watt Meter
Verify combined motor/prop and ESC efficiencies
43
Electronics
Electronics Flow Chart
Electronics/Sensors
Options
Power
Source
One
Source
Two
Sources
Real Time Data
Transmission
Requirement
Verification
On Board
Sensors
Recover
Recorded
Data
Radio
Control
From
Ground
Radar
Gun
Direct to
BEC/Servos
Flight
Dynamics
Board
Assistance
Stop watch
and Visual
45
Ground Station Setup
46
Communications Setup
Motor
Battery
ESC/BEC
Servos
Flight Dynamics
Board
Receiver
GS
47
Onboard Sensors
Onboard Sensor
Flight Dynamics
Board
Pitot Tube
Availability
Complexity
Sensor Output
Developed at
CU
-GPS signal
interference
-Software
interface
necessary
-Yields
groundspeed
-Can be used for
autonomous flight
later
Not readily
available
-Rigorous
calibration
-Extremely
susceptible to
damage upon
landing
-Yields airspeed
48
Flight Dynamics Board
Dimensions: 4” x 1.25”
Mass: 30 g
Power/Current: ~1.50 W /125 mA
Components
GPS Receiver
XB-Pro TX
Rate Gyro
Microchip
Pressure Gauge
Cost: ~$500
49
Servos
Maker/Model
Mass
(g)
Torque
(N-m)
Cost
($)
Gear material
System 3000
TS-5 S3K
9.6
.15
21.99
Nylon
Hitec HS-50
6.4
.06
27.48
Nylon
Hitec HS-55
8
.127
19.95
Nylon
Hitec HS56HB
10.7
.06
24.99
Nylon
JR NES 375
Micro Servo
9
.127
40.6
Metal
50
Receivers
PCM 1024
Receivers
Mass
(g)
Current
Draw
(mA)
Voltage
(V)
Cost
($)
Range
(m)
Futaba
R148DP
30.4
14
4.8 - 6
129.99
3600
Futaba
R138DP
50
13
4.8 - 6
99.99
3600
Futaba
R146iP
16.5
13
4.8 - 6
79.99
300
51
Energy Budget
Component
Required
Battery Capacity (mAh)
Motor
517
Flight Dynamics
Board
52
Servos
50
Receiver
6
Subtotal
625
Total
(SF = 1.2)
750
52
Battery Selection
Battery
(LiPoly)
Capacity
[mAh]
Voltage
[V]
Mass
[g]
Rating
[C]
Hyperion LiteStorm VX
800
11.1
62
20
Apogee
800
11.1
65
20
Venom Racing
800
11.1
84
20
53
Mass Budget
Subsystem
Item Description
Aerodynamics
Wing
45-80
Wing Spars
10-15
Fuselage
10-15
Backbone Spar
5-10
Tail
10-20
Wing Deployment Mechanism
20-50
Fasteners
5-10
Dummy Payload
75
Receiver
30
Servos
Electronics
10-20
Flight Dynamics Board
30
GPS Antenna
10
Connectors/Wires
Battery
Propulsion
Mass [g]
Propeller
10-20
62
5-10
Motor
24
ESC
6
Total Vehicle Mass
367 - 487
Remaining Mass for Tube
78 - 198
Total Mass
565
54
Project Plan
Preliminary Cost Analysis
Subsystem
Aerodynamics
Electronics
Item Description
Quantity
Cost ($)
Wing/Fuselage Material
N/A
200
Flight Dynamics Board
1
500
RC Controller
1
300
Onboard Receiver
1
130
Servos
2
25
GPS Antenna
1
30
TBD
TBD
Battery
3
150
Propeller
5
30
Motor
2
150
Li-polymer Battery Charger
1
120
Wattmeter
1
55
ESC
1
50
Tube/Pole
1
200
Connectors/Wires/Other
Propulsion
Deployment
Total (With ~50% Margin)
2500
56
Subsystem Interdependence
Electronics
Ground
Station
Vehicle
Receiver
RC
Transmitter
Propulsion
Wing
Flight
Dynamics
Board
Data
Receiver
Battery
Control
Surfaces
Servos
Speed
Controller
Fuselage
Motor
Swing
Mechanism
Propeller
Tube
57
D-SUAVE
1. Project
Management
2. Systems
Engineering
3. Vehicle
Design
4. Propulsion
5. Electronics
& Comm.
6. Testing &
Verification
1.1.
Organization
2.1. Design
Orientation
3.1.
Configuration
4.1. Motor
Selection
5.1.
Sensors
6.1.
Endurance
1.2.
Schedule
2.2. Subsystem
Coordination
3.2.
4.2. Propeller
Aerodynamics
Selection
5.2.
Controls
6.2.
Velocity
1.3.
Planning
2.3. Project
Feasibility
1.4. Group
Dynamics
3.3.
Lift-to-drag
4.3. Driver
Selection
5.3. Comm.
System
6.3.
Mass
3.4.
Stability
4.4. Battery
Selection
5.4. Power for
Electronics
6.4.
Deployment
3.5.
Wing Area
3.6.
Structure
3.7.
Deployment
6.5.
Propulsion
Work Breakdown Schedule
Gantt Chart
Advisor
Donna Gerren
Advisor
Lee Peterson
Project Manager
Burhan Muzaffar
Aero / Advisor
Michel Lesoinne
Systems Engr.
Nathan Sheiko
Fabrication Engr.
David Goluskin
CFO
Jastesh Sud
Safety Engineer
Yoshi Hasegawa
Deployment
Vehicle Design
Michael Lapp
Nathan Sheiko
Burhan Muzaffar
Michael Lapp
Yoshi Hasegawa
Assistant P.M
Miranda Mesloh
Webmaster
Brandon Bobian
Structure
Electronics
Propulsion
Jastesh Sud
Brandon Bobian
Miranda Mesloh
David Goluskin
Nathan Sheiko
Major Risk Assessments
Risk
Consequence
Plan to Avoid
Vehicle does not transition to
stable flight
Vehicle crashes and must be
repaired
Verify first by hand launch
over soft ground
High drag
Endurance requirement not
met
Minimize drag sources and
model thoroughly
Vehicle unstable
Vehicle design must be
altered
Glide test at early stage and
design movable CG
Insufficient controllability
Control surfaces must be
resized
Ground test RC control and
size surface with SF
Low efficiency propeller
Endurance requirement not
met
Test propulsion system in
wind tunnel
Pole is structurally unsound
Deployment test cannot be
performed
Design pole with large SF
61
Link
Design Choice Summary
Packaged in a tube
Fixed wing aircraft
Conventional configuration
Deploy using dynamic pressure
Wings fold side-by-side
Wings deployed by elastic or springs
Elevator and rudder
Conventional tail
Outrunner motor
Pulling propeller
On-board sensors for verification
62
Acknowledgements
Bill Pisano: Assisted with electronics setup
Dale Lawrence: Helped redefine the project
Donna Gerren: PAB Advisor
Jeremiah Hall: Blade element code for
propulsion
Kamran Mohseni: Project customer
Lee Peterson: PAB Advisor
Michel Lesoinne: Assisted with aerodynamics
numerical analysis
Sedat Biringen: Advised on aerodynamics
numerical modeling
63
References
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
Abdulrahim, Mujahid et. al., Flight Testing A Micro Air Vehicle Using Morphing For
Aeroservoelastic Control<mav.mae.ufl.edu/mujahid/publications/sdm04.pdf>.
Airfoil Data, 18 Sept. 2006, <http://www.nasg.com/afdb/index-e.phtml>.
Askeland, Donald. Essentials of Materials Science and Engineering, Ontario: Nelson, 2004.
Cengel, Yunus. Introduction to Thermodynamics and Heat Transfer, New Jersey: McGraw Hill,
1996.
Corke, Thomas. Design of Aircraft. New Jersey: Prentice Hall, 2003.
Futaba Receivers, 27 Sept. 2006, <http://www.hoverhobbies.com>.
Graphite Store, 13 Oct. 2006 <http://www.graphitestore.com>.
Hall, J., Mohseni, K., Lawrence, D. A., Investigation of Variable Wing-Sweep for Applications in
Micro Air Vehicles. Boulder, CO: University of Colorado, 2005.
Lawrence, Dale A., Mohseni, Kamran, Efficiency Analysis for Long-Duration Electric MAVs.
University of Colorado, Boulder, CO, 2005.
Micro Servos, 18 Sept. 2006 <http://www.modelflight.com>.
Peak Efficiency, <http://www.peakeff.com>.
Schmidt, Louis. Introduction to Aircraft Flight Dynamics. Reston: AIAA, 1998.
Shevell, Richard., Fundamentals of Flight, 2nd Ed., New Jersey: Prentice Hall, 1989.
Tower Hobbies, 18 Sept. 2006 <http://www.towerhobbies.com>.
Vable, Madhukar. Mechanics of Material. New YorK: Oxford University Press, 2002.
Warren F. Phillips, Mechanics of Flight. Hoboken, NJ: John Wiley & Sons, 2004.
64
Appendices
Number of back up slides - 61
Risk assessment
Design Alternatives
Deployment
Aerodynamics
Structure
Propulsion
Electronics
Project Plan
(RA)
(DA)
(DP)
(AD)
(ST)
(PP)
(EC)
(PP)
65
Deployment Risk Assessment
Low
Likelihood
Mid
High
Vehicle does not
transition to
stable flight
Vehicle does not
deploy
Low
Wings do not
deploy
Mid
Consequence
High
66
Link
Deployment Risk Avoidance
Risk
Consequence
Plan to Avoid
Vehicle does not
transition to stable flight
Vehicle crashes and
must be repaired
Verify first by hand
launch over soft ground
Wings do not deploy
Vehicle crashes and
must be repaired
Verify wing deployment
first on ground
Vehicle does not deploy
from tube
Deployment cannot be
verified
Verify tube deployment
first on ground
67
Vehicle unstable
Landing
damages vehicle
High drag
Insufficient
controllability
Low
Likelihood
Mid
High
Aerodynamics Risk Assessment
Insufficient lift
Low
Mid
Consequence
High
68
Aerodynamics Risk Avoidance
Risk
Consequence
Plan to Avoid
High drag
Endurance requirement
not met
Minimize drag sources
and model thoroughly
Vehicle unstable
Vehicle design must be
altered
Glide test at early stage
Insufficient controllability
Control surfaces must
be resized
Ground test RC control
and size surface with SF
Insufficient lift
Wing must be resized
Size wing with SF
Landing damages
vehicle
Vehicle must be
repaired
Design for low stall
speed
69
High
Manufacturing
difficulties
Likelihood
Mid
Aeroelasticity
impairs
performance
Low
Structure Risk Assessment
Wing
deployment
damages wing
structure
Low
Mid
Consequence
High
70
Structure Risk Avoidance
Risk
Consequence
Plan to Avoid
Manufacturing difficulties
Delay in schedule and
increase in costs
Use well-known
techniques and budget
time for mistakes
Aeroelasticity impairs
performance
Aerodynamic
performance impaired
Design for maximum
allowable deflections
rather than yield strength
Wing deployment
damages wing structure
Vehicle must be repaired
and wing structure must
be redesigned
Design structure with SF
and ground test
deployment
71
High
Propulsion Risk Assessment
Likelihood
Mid
Insufficient thrust
Low
Low efficiency
Insufficient
battery capacity
Low
Mid
Consequence
Overheating
High
72
Propulsion Risk Avoidance
Risk
Consequence
Plan to Avoid
Low efficiency
Endurance
requirement not met
Test propulsion system
in wind tunnel
Insufficient thrust
Velocity requirement
not met
Test propulsion system
in wind tunnel
Overheating
Motor and electronics
damaged
Monitor motor
temperature during
wind tunnel testing
Insufficient battery
capacity
Endurance
requirement not met
Confirm battery
capacity on ground
73
Electromagnetic
noise due to
motor
Poor electrical
connections
Low
GPS signal
interference
Likelihood
Mid
High
Electronics Risk Assessment
Unable to
receive real-time
data
-Transmission
range exceeded
-Insufficient
servo power
RC signal
interference
Low
Mid
Consequence
High
74
Electronics Risk Avoidance
Risk
Consequence
Plan to Avoid
Poor electrical
connections
Control and sensor
failure
Ground test all possible
connections
GPS signal interference
Velocity requirement not
verified
Insulate antenna from
below and ground test
Electromagnetic noise
due to motor
Control and sensor
failure
Ground test electronics
with motor running
Transmission range
exceeded
Loss of control and
possible crash
Ground test range
Insufficient servo power
Insufficient control
Design with SF and
verify in wind tunnel
Unable to receive realtime data
Velocity cannot be
verified in real time
Develop ground station
software and ground test
Insufficient control
Use a high quality
transmitter and receiver
and ground test
RC signal interference
75
Configuration of Aircraft - (DA)
Configuration
Packagability
(40%)
L over D
(25%)
Stability
(20%)
Manufacturing
(15%)
Weighted
Total
(100%)
Low-AR
5
1
1
5
3.2
Flying Wing
4
3
2
4
3.35
Conventional
3
5
4
2
3.55
Canard
2
2
3
3
2.35
Biplane
1
4
5
1
2.55
76
Design Alternatives – (DA)
Packagability
(40%)
Aircraft
Length
Endurance
(20%)
Time Aloft
Speed
(20%)
Maneuverability
(10%)
Speed
Turn Radius
Overall Mass
(10%)
Weighted
Total
Mass
Balloon
2
5
1
4
5
2.9
Airship
2
4.5
3
4
3
3
Glider
4
2
2.5
3.5
2
3.05
Helicopter
4
3
4
4.5
3
3.75
Fixed Wing
4
4.5
4.5
3
2
3.9
Rocket
5
1
5
1
5
3.8
Best
77
Available Design - (DA)
Aircraft
Overall Mass
Maneuverability
Packagability
Endurance
Speed
Mass (kg)
Turn Radius (m)
Length (m)
Time Aloft (min)
Speed (m/s)
Balloon
0.10
0.00
collapse
limited by leaks
0.00
Airship
0.1-0.8
0.00
collapse
15.00
10.00
0.50
2.64
0.80
3.00
9.00
0.3-1.5
0.00
0.80
12.00
13.00
Fixed Wing
0.50
10.00
0.80
15.00
15-20
Rocket
0.11
81.63
0.47
0.43
50.00
Glider
Helicopter
78
Deployment Calculations – (DP)
Option
Peak
Force (N)
Peak
Complexity
Acceleration
(# Parts)
(Gs)
Parachute
58.11
13.94
3
Spring
12.50
3.00
2
Elastic
9.50
2.28
2
Track
4.85
1.16
4
79
Force – (DP)
80
Force – (DP)
81
Stall Speed – (DP)
82
Aerodynamics Analysis Model – (AD)
Span
Chord
Taper
Twist
Induced
Drag
Parasite
Drag
Spanwise CL
Distribution
Define Geometry
Calculate
Circulation
Wing Loading
Drag
Lift-to-Drag
Airfoil
Data
Flight Performance
Characteristics
Stability
Derivatives
Dynamic
Derivatives
Twist Angle – (AD)
45
40
L over D
35
30
25
20
15
10
-20
-10
0
10
twist angle (deg)
20
84
Configuration of Aircraft – (AD)
Flying Wing vs Conventional
Focus on Longitudinal Controllability
(Fast pitching behavior of the aircraft at the trim condition)
Assumption: No Control Effect        
,
M  I y
M  Q S cCm
 c
Cm  Cm   
 2V



 Cm   Cmq 



 c 


I y  Q S c
 C m  C mq   Q S c C m   0
 2V 
Eq.(1)
85
Configuration of Aircraft – (AD)
Flying Wing vs Conventional
Cmq
 lH  CL 
 2.20VH  

 c    H
Cm
 lH  CL 
 2.20VH    

 c    H
Flying wing: No damping terms in Eq. (1)
Conventional Aircraft: Damping terms in Eq. (1)
Conventional Aircraft is more controllable.
86
Airfoil Data – (AD)
1.2
1
CL
0.8
Aquila
Verbitsky BE50
Clark Y
Cody RobertsonCR001
0.6
0.4
Drela Dae51
EpplerE387
0.2
Wortmann Fx60100
Martin Hepperlemh32
RG15
0
S 3014
-0.2
-4
-2
0
2
4
6
Angle of Attack [ - deg]
8
10
87
Drag Polar – (AD)
1.4
1.2
1
0.8
0.6
Cl
Aquila
Verbitsky BE50
Clark Y
Cody RobertsonCR001
0.4
EpplerE387
0.2
S 3014
0
-0.2
-0.4
0.01
0.02
0.03
0.04
Cd
0.05
0.06
88
Package Drag and Size Budgets – (AD)
Tmax 
W0
Wstores

L / D 0 L / D stores
L / D 0  Dstores
L / D stores D0
Dp 
CD 
(Assuming range is not a limitation when carrier
vehicle is operating at maximum thrust )
(Assuming stores affect drag but not lift)

m0
T 
 1
2  ma.c.  m p 
Dp
QS
 K C f
(1)
(Assuming parasite drag is the only drag
on the stores)
Equation (1) implicitly relates maximum length to diameter
Span is related to length, and chord is related to diameter
Remin  60,000
(No airfoil data for Reynolds numbers below 60,000)
89
Package Drag Budget – (AD)
Tmax 
W0
Wstores

L / D 0 L / D stores
L / D 0  Dstores
L / D stores D0
Dstores  D0
(Assuming range is not a limitation when carrier
vehicle is operating at maximum thrust )
(Assuming stores affect drag but not lift)
W0
m0
 D0
Wstores
mstores
Dstores  D0  D
mstores  ma.c.  m p


m0
D  T 
 1
m m

p
 a .c .

Dp 
D
2
90
Package Size Budget – (AD)
CD 
Dp
QS
 K C f
S  Dl 
D 2
2
K  f  Dl 
Cf 
0.455
2.58
log 10 Re 
Re 
(1)
(Assuming parasite drag is the only drag
on the stores)
(Assuming package is a cylinder)
(Shevell Fig. 11.4, which assumes Mach 0.5)
(Assuming turbulent boundary layer)
Vc

Equation (1) implicitly relates maximum length to diameter
91
Vehicle Design – (AD)
Lift-to-drag ratio for Varying Sweep
19.5
19
V = 10 [m/s]
V = 16 [m/s]
18.5
18
L/D
17.5
17
16.5
16
15.5
15
10
20
30
40
Sweep [deg]
50
60
70
92
Tail Design – (AD)
SVT  CVT
bW SW
lVT
cW SW
S C  CC
lC
S HT  CHT
cW SW
lHT
SV Tail  SVT  S HT
Historical Data of Vertical and Aft Horizontal Tail Coefficients
Aircraft Type
CVT CHT
Sail Plane
0.02
0.50
Homebuilt
0.04
0.50
Single Engine General Aviation
0.04
0.70
93
Tail Design
b
Assumption: Length of the aircraft is l Aircraft  W  cW
2
SW  0.068  0.095 m 2
l Aircraft  0.505  0.575m
bW  0.85  0.95 m
lVT  lHT  0.303  0.345 m
cW  0.08  0.1 m
94
Tail Design
Tail Areas
Tail
Area (m2)
SVT
0.0067 – 0.0119
S HT
0.0079 – 0.0157
SV Tail
0.0146 – 0.0276
S H Tail
0.0142 – 0.0268
95
Tail Design
10
8
0.1
7
0
L over D
Wing z-coordinate
9
-0.1
0.4
6
5
4
0.2
0
3
0.4
-0.2
Wing y-coordinate
0.2
-0.4
0
2
0
2
4
6
alpha (deg)
8
10
Wing x-coordinate
96
Tail Design
10
0.1
8
0
7
L over D
Wing z-coordinate
9
-0.1
0.4
6
5
0.2
4
0
3
-0.2
Wing y-coordinate
0.2
-0.4
0
Wing x-coordinate
2
0
2
4
6
alpha (deg)
8
10
97
Tail Design
10
8
0.1
7
0
L over D
Wing z-coordinate
9
-0.1
0.4
0.2
6
5
4
0
3
0.4
-0.2
Wing y-coordinate
0.2
-0.4
0
Wing x-coordinate
2
0
2
4
6
alpha (deg)
8
10
98
Tail Design
10
0.1
8
0
7
L over D
Wing z-coordinate
9
-0.1
0.4
0.2
0
5
4
-0.2
Wing y-coordinate
6
0.2
-0.4
0
Wing x-coordinate
3
2
0
2
4
6
alpha (deg)
8
10
99
Potential Wing Materials
Material
u
(MPa)

(g/cm3)
E (GPa)
Cost ($/m2)
2024-T3 Al
483
2.78
73.1
55.43
7075-T6 Al
572
2.81
71.7
101.50
Aramid-Epoxy
(60% fiber vol.)
1379
1.44
75.8
43.01
S-Fiberglass-Epoxy
(60% fiber vol.)
1510
2.05
53.1
34.40
Carbon-Epoxy
(60% fiber vol.)
1241
1.55
145
64.54
Blue Foam
0.12
0.021
0.021
16.14
Balsa
25
0.2
4.1
18.70
100
Wing Materials – (ST)
Material
 u (MPa)  (g/cm3) E (GPa)
2024-T3 Al
483
2.78
73.1
1
1
7075-T6 Al
572
2.81
71.7
1.17
1.02
Kevlar+Epoxy
(unidirectional)
1355-1980
9.06-7.54 44.0-60.0
0.86-1.51
1.66-1.11
S-glass+Epoxy
1355-1980
9.32-7.96 34.7-45.1
0.84-1.43
2.11-1.62
Carbon(HS)+Epoxy
1648-2449
9.13-7.66 81.9-120
1.04-1.84
0.89-0.61
Carbon(HM)+Epoxy
700-932
9.17-7.72 146-222
0.44-0.70
0.50-0.33
Blue Foam
0.12
0.021
0.021
0.03
3481
Balsa
25
0.2
4.1
0.72
17.8
 M1 


M2 
  c1 


  c2 
HS = High Strength
HM = High Modulus
High Strength Carbon Fiber + Epoxy is Best Material
101
Wing Materials
Cantilever beam with a uniformly distributed load
Maximum deflection occurs at the free end is
 max
w l 4 W l 3  Al 4 g  l 4 g 1




8EI 8EI
8EI
8E r 2
r is the radius of gyration
W=wl
102
Turn radius & Model – (ST)
103
Spar Deployment Stresses
104
Spar Deployment Stresses
105
Propulsion Power – (PP)
Steady Level Flight
http://www.teachengineering.com/collection/cub_/lessons/cub_images/cub_air
planes_lesson02_fig1.jpg
106
Propulsion – (PP)
Motors vs Total Mass
110
100
Total Mass [g]
90
80
8x4
7x3.5
6x4
70
60
50
40
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
Motors
107
Matlab vs MotoCalc – (PP)
Battery Capacity vs. MotoCalc and Matlab
700
Battery Capacity [mAh]
Matlab
MotoCalc
600
500
400
Custom CDR
Single Hotwind
Hacker A20 26M
Model Motors
2208/26
Komodo
KH220411
108
Propulsion Analysis Software
Input
D and P
Gear Ratio
Pout prop= 4.43 W
Jeremiah Hall’s
Blade Element
Code
Prop RPM’s
η prop
Pout motor= Pout prop / η prop
Repeat for 30 Selected Motors
Output
Total Mass
Motor Efficiency
Function
η motor
(Motor+Battery)
Curve Check
Battery Capacity
Pout batt= Pin motor / η esc
Pin motor= Pout motor / η motor
109
Motor and Battery Mass
(PP)
Total Mass for Selected Motors
100
Battery
Motor
Total Mass [g]
80
47.4
38.5
60
42.4
43.9
40
42
20
45
24
23
0
Custom CDR
Single Hotwind
Hacker A20 26M
Model Motors
2208/26
Komodo
KH220411
110
Thermodynamic Analysis – (PP)
Forced Convection
Ploss  hA(Ts  T )
Assuming: T∞=25 °C
Ts=68 °C
111
Electronics Mass and Cost
Budget – (EC)
Component
Mass (g)
Cost ($)
RC Control Receiver
30
130
Servos
20
25
Flight Dynamics Board
30
500
GPS Antenna
10
30
10-20
20-50
100-110
695-725
Connectors/Wires/Other
Total
112
Work Breakdown Schedule –(PP)