1. Critical Design Review
AAE451 – Team 3
Project Avatar
December 9, 2003
Brian Chesko
Brian Hronchek
Ted Light
Doug Mousseau
Brent Robbins
Emil Tchilian

2. Aircraft Name
Avatar
av·a·tar - n <chat, virtual reality> An image representing a user in a multi-user virtual reality.
Source: The Free On-line Dictionary of Computing

3. Introduction Walk Around Design Requirements and Objectives Sizing
Propulsion
Aerodynamics
Dynamics and Controls
Structures
Performance
Cost
Summary
Questions

4. Aircraft Walk Around Wing Span = 14.4 ft Wing Chord = 2.9 ft
A/C Length = 10 ft
T-Tail – NACA 0012
Pusher
Internal Pod
Tricycle Gear
Low wing – Clark Y

5. Design Requirements & Objectives
Maximum weight < 55 lbs
Cruise speed > 50 ft/sec
Stall speed < 30 ft/sec
Climb angle > 5.5°
Operating ceiling > 1000 ft
Flight time > 30 minutes
Payload of 20 lbs in 14”x6”x20” pod
Carry pitot-static boom
Spending limit < $300
T.O. distance < 106 ft (~60% of McAllister Park runway length)
Rough field capabilities
Detachable wing
Easy construction

6. Constraint Diagram

7. Propulsion

8. Ref. www.towerhobbies.com
Chosen Engine
O.S. Max 1.60 FX-FI
RPM
1,800-9,000 RPM
2.08 lbs
Fuel Injected
Ref.

9. Chosen Propeller 4-blades
Zinger 16X7 Wood Pusher Propeller
16 inches in diameter with 7 inch pitch
4 blades
Ref.

10. Ref. www.towerhobbies.com
Chosen Fuel Tank
Fuel tank chosen is:
Du-Bro 50 oz. fuel tank
Available from Tower Hobbies
Located at the C.G. of aircraft
Good for up to 32 min. of flight time (when completely full).
Ref.

11. Takeoff EOM Integration
Thrust
Drag + Rolling Friction
Position [ft]
Velocity [ft/s]
Velocity vs. Position at Takeoff
Takeoff Distance Within Constraint

12. Max Velocity Maximum Velocity Thrust Thrust/Drag [lbf] Drag
Flying Velocity [ft/s]
Thrust/Drag [lbf]
Maximum Velocity
Thrust
Drag

13. Aerodynamics

14. Wing Dimensions
Prandtl’s Lifting line theory used for aerodynamic modeling of the lifting components
Input parameters: AR, a0, aL=0, a.
Lifting Line Model Gives CL, CDi at prescribed a
CDvisc found using Xfoil which was used to obtain CD = CDi+CDvisc
5° Dihedral

15. Clark Y Airfoil has low drag over range of interest
Airfoil Selection
Region of Interest
Clark Y
Clark Y Airfoil has low drag over range of interest

16. Airfoil Selection Section Drag Coefficient Cd
Section Lift Coefficient Cl
Section Drag Coefficient Cd
Angle of Attack (AOA)
Section Lift Coefficient Cl

17. Wing Stall Performance
CL needed = 1.19
Wing without flaps reaches CL at a=13°
Wing stall possible
Wing with 15° flap deflection reaches CL at 11°
Required CL CL
Angle of Attack (degrees)
Flaperons necessary to meet stall requirements

18. Wing Performance
Required CL at stall CD CL

19. Drag Build Up At Cruise Component CD Drag Wing 0.018 2.6 lbf Fuselage
0.0045
0.6 lbf
Horizontal Tail
0.0043
Vertical Tail
0.0017
0.04 lbf

20. Wing Operating ParametersCL
a (of wing)
Flaperon Deflection CD
L/D
Stall
1.19
11°
15°
0.119 10
T/O
0.989 8°
0.084 12
Cruise
0.44
2.8° 0°
0.018 24

21. Dynamics and Controls

22. Center of Gravity & Aerodynamic Center
Aircraft Center of Gravity is 3.2 ft from nose.
Calculated from CAD program Pro-E
Aircraft Aerodynamic Center is 3.7 ft from nose.
Position where pitching moment of aircraft doesn’t change with angle of attack
Calculated using Lift from Wing and Horizontal Tail
Aerodynamic Center
Center of Gravity

23. Aerodynamic Center of Aircraft
Static Margin
Desired Static Margin is 15% - 20%
Dependent on C.G. and A.C. location
Static Margin is 15%
Contributes to
Horizontal Tail Sizing
Aerodynamic Center of Aircraft
Static Margin = 20%
Static Margin = 15%
Center of Gravity

24. Horizontal Tail Sizing
Tail sized based on desired static margin for static stability and take-off rotation ability
 double-dot should be at least 10 deg/sec2
Ref. Roskam, Airplane Flight Dynamics
Area
12 ft2
Span
6 ft
Chord
2 ft
2 ft
6 ft

25. Vertical Tail Sizing
Value of yawing coefficient due to sideslip angle should be approximately = 10e-4
Tail area should be ~2 ft2
Ref. Roskam, Airplane Design
Area
2 ft2
Span
1 ft
Chord
2 ft
2 ft
1 ft

26. 5° dihedral is a good compromise
Dihedral Angle
Recommendations
Survey of Roskam data on homebuilt & agricultural low-wing aircraft: ~5°
“Wing and Tail Dihedral for Models” - McCombs
RC w/ailerons (for max maneuverability, low wing): 0-2° EVD (Equivalent V-Dihedral ≈ dihedral)
Free Flight Scale model low wing: 3-8° EVD
5° dihedral is a good compromise

27. Control Surface Sizing
Sizes calculate from traditional lifting device percentages.
Ref. Roskam, Airplane Design
Flaperon
Elevator
Rudder
Chord
0.58 ft
0.6 ft
Inboard Position
0.95 ft
0.2 ft
0.1 ft
Outboard Position
7.2 ft
3 ft
1 ft
0.6 ft
0.58 ft
0.9 ft
6.25 ft
0.6 ft
2.8 ft

28. Trimming
Incidence of Horizontal Tail calculated from trimmed flight during cruise (0 Angle of Attack)
Analysis set
incidence at
-2

29. Structures

30. Wing Spar Design 2 Spar Design (at .15 & .60 chord): Resist Bending
Assuming 5-g loading
53 lbf weight
Safety factor of 1.5
Resist Torsion
Less than 1o twist at tip under normal flight conditions
Spar Results:
Material of Choice: Bass or Spruce Wood
Front Spar:
3.6” high (based on airfoil)
0.37” thick (0.73” at root)
Rear Spar:
3” high (based on airfoil)
0.16” thick (0.25” at root)

31. Longitudinal Beam Design
Resist Bending from:
20 lbf payload
Horizontal tail loads
Resist Torsion from:
Rudder deflections
Prop wash over tail
Beam Results:
Material of Choice: Bass or Spruce Wood
Beam Dimensions:
3” high
0.25” thick
8” between the beams

32. Tail Structures Foam core with carbon fiber shell
Horizontal and vertical tails comprised of carbon fiber w/ foam core
Possible to make two foam cores, and cure entire tail at one time
Control surfaces just need to be cut out of tail structure
Tail spars allow attach points and transfer load to beams

33. Rear Gear Design Blue lines represent pin joints
Black tie-downs absorb energy from landing
Up to a 33 ft/sec “crash” from 5 feet high
Need 18” relaxed length tie-down
Square aluminum tube transfers landing load to tie-downs and surrounding structure
1” x 1” x 0.065” thick – 6063-T6

34. Front Gear Design Aluminum Bolt Elastic Band & Nylon Bolt
Provides pivot for gear (does not break)
Elastic Band & Nylon Bolt
Elastic Band Absorbs some energy from landing
Nylon bolt breaks during hard landing
Front Gear Aluminum Tube
Designed not to break
Designed not to bend
Al tube:
1” x 1” x 0.065” thick 6063-T6

35. Ref. www.towerhobbies.com
Other Odds and Ends
Covering for Wing:
Coverite 21st Century Iron on Fabric
0.34 oz/ft2
Stronger, and resists tears better than MonoKote
Covering for Fuselage:
Fiberglass
Either mold or foam core
Not conductive – won’t interfere with internal electronics
Ref.

36. Final Weight Estimate

37. Performance

38. Aircraft Performance
(with 2.2lbf fuel)
90 ft/sec

39. Cost

40. Airframe Cost

41. Electronics Cost

42. Propulsion Cost

43. What Purdue Will Pay For This Project
Total Aircraft Cost
What Purdue Will Pay For This Project

44. Total Aircraft Value TOTAL AIRCRAFT VALUE = $106,341.15
Total Aircraft Value = (Engineering Pay) + (Cost) + (Value of Already Possessed Parts)
Engineering Pay = hr x $100/hour = $82,375
Aircraft Cost = $13,966.15
Value of Already Possessed Parts = $10,000
Micropilot = $5,000
Carbon Fiber & E-Glass = $5,000 (estimate)
TOTAL AIRCRAFT VALUE = $106,341.15
What Purdue Would Pay to Outsource This Project

45. Summary

46. Summary – Internal View
Internal Pod
Camera View

47. Summary – 3-View

48. Summary -Major Design Points
Aircraft Description
Aspect Ratio = 5
Wing Span = 14.4 ft
Wing Area ~ 42 ft2
Aircraft Length = 10 ft (not including air data boom)
Engine = 3.7 hp O.S FX-FI – Fuel Injected
Weight = 53 lbf
Aircraft Configuration
T-Tail
Low Wing
Pusher
High Engine
Tricycle Gear
Internal Pod

49. Questions?

50. References (I)
[1] MATLAB. PC Vers Computer Software. Mathworks, INC
[2] Raymer, Daniel P., Aircraft Design: A Conceptual Approach, AIAA Education Series, 1989.
[3] Roskam, Jan., Airplane Flight Dynamics and Automatic Flight Controls. Part I. DAR Corporation, Kansas
[4] Gere, James M., Mechanics of Materials. Brooks/Cole, Pacific Grove, CA. 2001
[5] Tower Hobbies. 9 December
[6] XFoil. PC Vers Computer Software. Mark Drela
[7] Niu, Michael C., Airframe Structural Design, Conmilit Press Ltd. Hong Kong
[8] Halliday, et al., Fundamentals of Physics, John Wiley & Sons. New York
[9] Roskam, Jan, Airplane Design (Parts I-VIII), Roskam Aviation and Engineering Corp. Ottawa KS
[10] Kuhn, P., “Analysis of 2-Spar Cantilever Wings with Special Reference to Torsion and Load Transference”. NACA Report No. 508.
[11] McMaster-Carr. 9 December
[12] Pro/ENGINEER. PC Release PTC Corporation.
[13] Roskam, Jan., Methods for Estimating Stability and Control Derivatives of Conventional Subsonic Airplanes. Publisher Jan Roskam. Lawrence, KS

51. References (II)
[14] Zinger Propeller. 9 December
[15] McCombs, William F., “Wing and Tail Dihedral for Models”, Model Aviation. Dec

52. Appendix

53. SIZING

54. Cruise Speed

55. Stall Speed

56. Climb Angle

57. Ceiling

58. Endurance

59. Takeoff

60. Landing Distance

61. PROP

62. Appendix
OS 1.60 FX-FI
Consistency: The Fuel Injection system constantly supplies the correct air/fuel mixture to the engine, regardless of speed, altitude, or attitude.
Recommended is a cc fuel tank that allows approximately 10 to 12 minute flights. = 30 min. with 50 oz. tank.

63. AERO

64. Aerodynamic Modeling Solving Prandt’s equation Substituting:
Prandtl’s Lifting line theory used for aerodynamic modeling of the lifting components
Solving Prandt’s equation
Substituting:
Equation to solve:
Main Results
CL = πAR*A1*(α- αLo)
System of N equations with N unknowns (Solve N  N matix)
Take N different spanwise locations on the wing where
the equation is to be satisfied: 1, 2, .. N; (but not at the tips, so: 0 <  < )
The wing is symmetrical  A2, A4,… are zero
Take only A1, A3,… as unknowns
Take only control points on half of the wing: 0 < i  /2

65. Choice of main wing airfoil
From lifting line with Initial parameters:
Rectangular planform, 1000 ft
a0 = 2pi,
αL0 = 0,
AR = 5;
W/S = 1.28 (from sizing)
CL =
Cl distribution found at cruise
Cl varies :0 to 0.58
Taking into account the Cl variation above, the need of an airfoil with a drag bucket at the specified Cl’s
Xfoil utilized for different foils at the above conditions

66. Clark Y Airfoil Drag Bucket location fits best
Airfoil Selection
Region of Interest
Clark Y
Clark Y Airfoil Drag Bucket location fits best

67. ClarkY foil Xfoil runs of ClarkY foil at cruise and take-off
Cruise: αL= -3.5deg
Takeoff no flap: αL= -3.8deg
Takeoff 10deg flap: αL= -7deg
Takeoff 15deg flap: αL= -7.8deg
In lifting Line Equation:
a0 – updated depending on condition
αL - updated according to above

68. Flaperons necessary to meet stall requirements
Stall Performance
CL needed = 1.19
Wing without flaps reaches CL at 13 deg aoa
Wing stall possible
Wing with 15 deg flap deflection reaches CL at 11 degrees
Required CL
Flaperons necessary to meet stall requirements

69. Stall Performance Drag Calculation
CDtotal = CDinduced+CDvisc
CDinduced – from Lifting line
CD visc – integrated at the found Cls
Required CL
CD = at required CL

70. Cruise Performance CL needed = 0.44 CL achieved at 2.8 deg
Total Lift produced = 57lbf
Total Drag = 2.6 lbf, L/D =21

71. Operating Parameters Stall 1.19 11 deg 15 deg 0.119 10 T/O 0.989CL
Aoa
Flap Deflection CD
L/D
Stall
1.19
11 deg
15 deg
0.119 10
T/O
0.989
8. deg
0.084 12
Cruise
0.44
2.8 deg
0 deg
0.018 24

72. D & C

73. Center of Gravity Center of Gravity of Aircraft
Weight of Horizontal Tail changes with area
Note: 0.44 lbs/ft2 based on aircraft sizing code

74. Aerodynamic Center
Aerodynamic Center as a function of Horizontal Tail Area
Roskam Eq 11.1
Raymer Fig 16.12

75. Takeoff Rotation Equation
This sizing based on angular acceleration during take-off rotation
Ref. Roskam 421 book, pg
Variable definitions found in above reference

76. Yaw Moment due to Sideslip
Vertical Tail sized from Coefficient of Yaw Moment due to Sideslip
Roskam Eq 11.8
Vol 2
Due to Wing and Fuselage:
Roskam Eq 10.42
Vol 6

77. Ref. McCombs, William F. “Wing and Tail Dihedral for Models.”
Dihedral Angle
EVD = A + kB
A = 0°
k = f(x/(b/2)) = 0.98
B = EVD / k ≈ EVD
A=0° B X CL
Ref. McCombs, William F. “Wing and Tail Dihedral for Models.”

78. Dynamics Short Period Mode Pole -14.391 ± 1.0079i Natural Frequency
(rad/s)
Damping Ratio
Phugoid Mode
Pole
± i
Natural Frequency
(rad/s)
Damping Ratio
Dutch Roll Mode
Pole
± i
Natural Frequency
(rad/s)
Damping Ratio
Spiral Mode
Pole
Roll Mode
Pole
Ref. Purdue University AAE565, Matlab Predator Code

79. STRUCTURES

80. What Materials to Use
Titanium
Bass / Spruce

81. Ref. www.towerhobbies.com
Material Properties
Titanium = difficult to obtain
Wood = not difficult to obtain
Ref Forest Products Laboratory Wood Handbook
Ref.

82. Twist Constraint (<1o)
Ref. Kuhn pg. 49
Where T = Torque (in-lbf)
L = Length (in)
l = f(B0, A0) (ref. Appendix)
A0 = f(E, I) (ref. Appendix)
B0 = f(G,J) (ref. Appendix)
E = Young’s Modulus (psi)
I = Moment of Inertia (in4)
G = Torsional Stiffness (psi)
J = Polar Moment of Inertia (in4)
Assumptions:
Small Deflections
Spars & Ribs Carry all Torsion
Span ~ 14.4 ft
Chord ~ 2.9 ft
Safety Factor = 1.5
G-Loading = 5.0
Weight = 53 lbs
Ref. Gere

83. Twist at Tip

84. Twist at Tip (Zoom) Chosen Front Spar = 0.73” thick
Chosen Rear Spar = 0.25” thick
(note, this doesn’t include the step)

85. (based on span-wise lift distribution)
Deflection at Tip
a (in)
L (in)
Load (lbf)
Ref. Gere pg. 892
Where Load = Weight*SF*G-loading (lbf)
L = Length (in)
E = Young’s Modulus (psi)
I = Moment of Inertia (in4)
Assumptions:
Small Deflections
NO TORSION
Span ~ 14.4 ft
Chord ~ 2.9 ft
Safety Factor = 1.5
G-loading = 5.0
Weight = 53 lbs
For this design:
a ~ 3 ft or 36 in
(based on span-wise lift distribution)

86. Chosen Spar Configuration
Deflection at Tip
Chosen Spar Configuration

87. (based on span-wise lift distribution)
Is Stress too High?
Load (lbf)
Ref. Gere pg. 323
a (in)
Where M = Weight*SF*G-loading*a (in-lbf)
y = Maximum Dist from Neutral Axis (in)
I = Moment of Inertia (in4)
L (in)
Assumptions:
Span ~ 14.4 ft
Chord ~ 2.9 ft
Safety Factor = 1.5
G-loading = 5.0
Weight = 53 lbs
For this design:
a = 3 ft or 36 in
(based on span-wise lift distribution)

88. Max Tension Stress

89. Max Compression Stress

90. Ref. www.towerhobbies.com
Covering
Traditional Monocote may not be strong enough for these large aircraft
Coverite 21st Century Iron on Fabric is stronger, and resists tears much better
0.34 oz/ft2
Approx. 2 lbs for entire wing
Ref.

91. Summary Main Wing Spruce or Bass wood Front Spar Rear Spar
0.73” thick by 3.6” high
Rear Spar
3/8” thick by 3” high h t

92. Rear View of Tail
NOTES
Torsion can effectively be reduced with appropriate beam spacing
Bending can be reduced by increasing moment of inertia of beams (not spacing)
Some torsion is inherent, torsion can not be negated as it could in wing
Side force from V-stab creates torsion effect on beams
Downward force from H-stab creates bending moment on beams

93. Deflection at Tip (Rear of Tail)
Load (lbf)
Ref. Gere pg. 892
L (in)
Where Load = (lbf)
L = Length (in)
E = Young’s Modulus (psi)
I = Moment of Inertia (in4)
Assumptions:
Small Deflections
Safety Factor = 1.5
G-loading = 3.0
Rectangular Beams
Current Known Values:
L = 6.2 ft
Load ~ 8 lbf
Moment of inertia of rectangular beam:
I (in4) = (t)(h3)/12
t and h shown on next slide

94. Deflection at Tip (Rear of Tail)
Green = spruce
Black = bass h
h=2 in t
h=3 in

95. Deflection at Tip (Rear of Tail)
Green = spruce
Black = bass h
h=2 in t
h=3 in
Required t ~0.55 in

96. Landing Gear Placement (I)
θ = tipback angle =
Landing gear placement based on
guidelines found in Raymer

97. Landing Gear Placement (II)
γ = overturn angle =
Landing gear placement based on
guidelines found in Raymer

98. Easily Obtainable Square Tubing
Ref.

99. Buckling of Rear Gear Load L Load For Rear Gear: L ~ 15.3 in
Ref. Gere pg. 763 L
Where L = Length (in)
E = Young’s Modulus (psi)
I = Moment of Inertia (in4)
A = Cross Sectional Area (in2)
Load
Assumptions:
Pinned-Pinned Column
1st Mode Buckling
No Eccentricity
For Rear Gear:
L ~ 15.3 in

100. Compressive Failure of Rear Gear
Load L
Ref. MIL-HDBK-5H: 3-255
Where Load = (Weight)(S.F.)(Gloading)
A = Cross Sectional Area (in2)
Load
Assumptions:
Weight = 53 lbf
Gloading = 10
S.F. = 1.5
Aluminum 6061-T6
No Buckling

101. Smallest easily obtainable tubing: 1” x 1” x 0.062”
Stress on Rear Gear
Smallest easily obtainable tubing: 1” x 1” x 0.062”
t=0.062”
t=0.125”

102. Great, what about the bungee?
Consider worst reasonable landing situation
Moving at (1.1)Vstall
5 feet above ground
Aircraft falls out of the sky
Can the bungee absorb the energy associated with this landing?

103. Great, what about the bungee?
Assumptions:
Weight = 53 lbf
Vstall = 30 ft/sec
Altitude = 5 ft
Don’t want x to exceed 3 inches (beyond initial stretch) on landing

104. What Spring Constant is Needed?
Required k ~ 3.75 lbf/in
1/k ~ in/lbf

105. What is the Spring Constant?
Relaxed Length ~18 inches

106. How Big is the Bolt? Diameter of nylon bolt = 0.5 in Load Reaction
Ref. Gere pg 900
Load
If load = (Weight)(S.F.)(Gloading) = 795 lbf
Reaction = 1770 lbf (instantaneous)
Need cross sectional area of bolt to be in2
Diameter of nylon bolt = 0.5 in
Reaction
Assumptions:
Weight = 53 lbf
Gloading = 10
S.F. = 1.5
3.1”
6.9”

107. PERFORMANCE

108. Endurance Avg. Engine Fuel Consumption = 45.455 mL/min
Endurance = Fuel / Consumptionfuel
Avg. Engine Fuel Consumption = mL/min
Endurance = 30 min

109. Since this is RC, assume almost instaneous cruise conditions
Range
Since this is RC, assume almost instaneous cruise conditions
L/D = 19
Cbhp = 1.5 lb/hr/bhp
Prop eff = .67
Fuel Frac = 1.043

110. Minimum Flight Velocity
Velocitymin= ft/sec
Weight = 53 lbf
CLmax = 1.19
q =1.067 lbf/ft^2

111. Rate of Climb Vv= 7.5 ft/sec D = 6.5lbf hpengine = 3.7 hp W = 53 lbf
Prop Eff = .3

112. Maximum Velocity Maximum Velocity Thrust Thrust/Drag [lbf] Drag
Flying Velocity [ft/s]
Thrust/Drag [lbf]
Maximum Velocity
Thrust
Drag

113. Climb Angle
Vv = 7.5 ft/sec
V = 33 ft/sec