1. Team One Purdue University AAE 451 Project Debriefing 28 April, 2005
Agenda:
Design Concept
Aerodynamics
Structures
Propulsion
Dynamics and Controls
Design Code
Construction
Testing
John W Apostol
Michael S Carpenter
Christopher L Grupido
Jeeyeon Hahn
Daniel J Halla
Douglas K Klutzke
Christopher E LaMaster
Joshua T Mook
Todd M Mostrog
Thomas B Shaw

2. DR&O Requirements Create a Stylish, Remotely Piloted Aircraft
Operable Airspace
Within football field
6-30 ft vertical
Takeoff / Landing
Less than 100 ft
Flight Angles
Climb Angle of 30o
Descent Angle of -5.5o
Velocities
Loiter at less than 30 ft/s
Stall at less than 20 ft/s
Maximum greater than 40 ft/s
Turn
30 ft Radius
Endurance
Greater than 8 minutes
Feedback Controller

3. Design Concept

4. Original Concept Original Concept Areas of Concern
Center of Gravity Location
Low Aspect Ratio Wing
Possible Tail Strike
Aerobatic Capabilities
Marketability/Expense
How do we increase
performance capabilities
without compromising unique
visual appeal?

5. Final Design Final Design Modifications Single-engine configuration
Shortened tail boom
Moved leading edge
Design kept as close to original concept as possible
Achieved higher performance,
possibly aerobatic design that
still maintains a unique shape
and style.
Two motors with pusher props.
Stability Concerns.
Performance Concerns.
Marketability Concerns.
Shorter tail boom length.
Increases static margin.
Reduces take-off problems.
Modifying the planform.
Performance aspects of aerobatic Extra 300 wing incorporated.
Design kept as close to original concept as possible.
Leading edge moved back.

6. Constraints Wing Loading of 0.59 lbf/bhp Power Loading of 20 lbf/ft2
Design Point

7. Overview Aircraft Horizontal Tail Weight = 2.05 lbf Area = 0.68 ft2
Wing
Area = 3.4 ft2
Span = 4.6 ft
Root Chord = 0.97 ft
Tip Chord = 0.52 ft
Fuselage
Length = 2.8 ft
Max Diameter = 0.33 ft
Horizontal Tail
Area = 0.68 ft2
Span = 1.65 ft
Root Chord = 0.47 ft
Tip Chord = 0.36 ft
Vertical Tail (per tail)
Area = 0.13 ft2
Span = 0.52 ft
Root Chord = 0.36 ft
Tip Chord = 0.22 ft

8. Aerodynamics

9. Airfoil Selection Wing: Selig 2091 Tail: NACA 0009
Compatible with Low Reynolds Numbers
Simple Shape
Aerodynamic Performance
Wing: Selig 2091
Tail: NACA 0009

10. How to balance ‘style’ versus aerodynamic performance?
Wing Geometry
Define Wing Geometry
Aspect Ratio
Sweep
Taper
Dihedral Angle
Winglets
How to balance ‘style’ versus
aerodynamic performance?

11. Market Study Wing Geometry
Sampled 25 random people on style parameters
No Winglets
Winglets

12. Market Study Weighted Market Study versus Aerodynamic Performance
Dihedral Angle: 3 degrees
Sweep Angle: 3 degrees
Winglets: 3 inches
Aspect Ratio: 6.2
Taper Ratio: 5

13. Lift Lift slope equation: Maximum Wing Lift Coefficient:
Coefficients for the lift slope equation where calculated using FlatEarth Predator code
CLmax estimation was referenced from Raymer
CLo for the aircraft was determined using FlatEarth
*********************** maybe tabulate CLmax, CLo and CLa into a table for sectional, finite wing, and whole aircraft *********************************************

14. Drag 3-D accounts for build up of all drag components:
Pressure, friction, induced, and miscellaneous
Wing, Tail, Fuselage, Winglets, Landing Gear, and Miscellaneous
Drag Equation:
This plot includes parasite drag that was covered in previous slide. For the aircraft, induced drag is also added as a parameter.
The drag profile is a good estimate of the efficiency of the aircraft. Lower drag for higher values of lift will give a more efficient aircraft
******************************** it may be good to show the point where our aircraft will be operating for the majority of flight *********************************
i.e. Vloiter
like lift, it may be nice to tabulate Cd, Cdo, and Cdi for sectional, wing, and whole aircraft

15. Structures

16. Structural Analysis
For Maximum Load Factor of 2.26 and Weight of 2.0 lbf
Maximum Load = 4.5 lbf
Load Modeling Assumptions
Uses Bending Moment and Euler Buckling Criteria
Number of Stringers of Wing and Fuselage Design
Based on Max Load Case and a Safety Stress Factor
Constant Size Spars
Favorable Effect of Landing Gear Weight Ignored

17. Structural Properties
Wing
0.25 lbf Weight
0.6 in Max Deflection
2 Spars
26 Ribs
2.2 in Separation
2 Stringers
Horizontal tail
0.04 lbf Weight
12 ribs
2.34 in Separation
1 Spar
2 stringers
Fuselage
Weight = 0.05 lbf
Max Moment = 152 in-lbf
0.04 degrees/inch of twist
2 Spars
12 Bulkheads
2.8 inches of Separation
6 Stringers

18. Landing Gear Two 21/16 in diameter foam wheels Fixed to wing ribs
Top half hidden in wing section
Designed for no failure
Forward landing gear in wing
Simple structure attachment
Reduce drag
Aft landing gear on tail
Rudder mounted for ground directional control
Four points of contact
Increases on ground stability 8o
10o

19. Propulsion

20. Constraints Effect of increased climb angle Decreases power loading
25.5 to 18.6 lbf/bhp
0.12 bhp needed to meet DRO
Climb Angle = 20o
Climb Angle = 30o

21. Motor Selection Graupner SPEED 600 7.2V 0.43 lbs (6.88 oz)
0.178 shaft horsepower
0.12 useful horsepower
69% motor efficiency
65% propeller efficiency at loiter
22 amps max
12 amps nominal loaded
7.2 volts
8600 Loaded RPM
Small Propeller Size (No Gearbox)

22. Batteries and Controller
Kokam Li-Poly
2 cell
1250 mAh
7.4 volts
2.92 oz
1.65” by 3.1” by .43”
Castle Creations Griffin-40
40 amp
.9 oz
1.4” by .6”
Flight Time = 15 minutes

23. Propeller Selection Physical Characteristics Small Propeller
Short landing gear
Multiple Propeller Testing
8” by 6” two blade
8” by 4.5” two blade
8” by 6” three blade
Landing Concerns
Prop Clearance if Pitched Forward

24. Dynamics & Controls

25. Center of Gravity Center of Gravity: 0.25*Aircraft Length
Motor
Batteries
Wing
Aileron
Servos
Landing
Gear
Fuselage
Elevator
& Rudder
Tail
Increasing Weight
This is an artist representation of the center of gravity of the aircraft. The farther the blue arrow points down, the greater the weight of that component. The CG was calculated by first identifying the location of each component and it’s corresponding CG. Then the sum of the x-distance times the weight divided by the sum of the weight results in a CG location as a percent of aircraft length.
Center of Gravity:
0.25*Aircraft Length
0.30*Mean Chord
Static Margin = 12.63%

26. Incidence Angles
Wing incidence defined by necessary level flight lift condition
Tail incidence determined by trimming the aircraft to zero moment about the center of gravity
Wing Incidence = 2.4°
Tail Incidence = 5.0°
Relative Tail Incidence to Local Flow = 2.1°
Incidence angles were calculated to ensure the aircraft could attain the necessary lift and would be trimmed in a desirable orientation.

27. Dynamic models Aircraft transfer function
Rate gyro transfer function: 1
Actuator transfer function: 1
Control Law Laplace Transform:

28. Feedback Control Lateral-Directional feedback Poles = 0 0.30452
Negative Gain
Stable at zero gain except for spiral mode which is slightly unstable--- alert pilot to
Can have basically any negative damping, but don’t want anything too large because we don’t want a slow response.
Can have a slightly positive gain, but not too much, or the aircraft will become unstable.
Examine the effects of changing the gain
xxxxxxx
Spiral Mode
Dutch-Roll
Roll
Positive Gain

29. Design Code

30. Design Code User Excel Input File Design Cycle Constraints
Weight Build-up
Main Wing Size
Static Margin
Horizontal Tail Size
Side Slip Derivative
Vertical Tail Size
Aero Performance
Structures
Take Off
Zero AOA Lift
Wing Incidence
Zero AOA Moment
Tail Incidence
Output File
FlatEarth

31. Construction

32. Construction

33. Construction

34. Construction

35. Construction

37. Construction
Torsion bar added

38. Flight Testing

39. “Flight” One

40. Pivoted thrust angle to 3° down
“Flight” One
Pivoted thrust angle to 3° down

41. “Flight” Two

42. “Flight” Two Added 1” to forward fuselage.
Moved batteries forward 1” additional.
Added 3 oz weight to motor mount.

43. Flight Three

44. Final Flight

45. Replacement / Repair Expenses
Expense Summary
First Build Expenses
Total Build Budget Allowed : $150.00
Total Budgeted Build Expense : $138.74
Total Remaining Budget : $ 11.26
Replacement / Repair Expenses
Replacement Expenses : $29.94
Repair Expenses : $19.71
Total Expenses : $49.65

46. Questions / Comments

47. Appendix

48. Final Design
Original Concept
Final Design

49. Style Features Tapered wing planform Provides appealing look
Unique polyhedral wing design
Combines distinctive visual styling cues
Winglets
Provide finished look to polyhedral theme
Hybrid fuselage/boom structure
Creates distinctive profile
Polyhedral tail design
Maintains consistency of aircraft appearance

50. Airfoil Selection Analyze airfoils from NASG database:
Based on operating Reynolds Number

51. Geometry Selection Selig 2091 Simple shape Acceptable t/c ratio
Best balance of aerodynamic characteristics
NACA 0009 chosen for tail
Suitable lift to trim aircraft
Acceptable aerodynamic characteristics
The Selig 2091 was the chosen section. It has a shape that will be easy to manufacture and the thickness ratio allows enough room for equipment to be stored within it near the fuselage. It had a good balance of aerodynamic characteristics with a low drag bucket and a higher Cl,max.
The tail geometry was more concerned with trimming the aircraft. The NACA 0009 provides suitable characteristics for lift and drag. Its symmetric shape will be easy to manufacture and the small thickness will give a smaller weight increase for the rear part of the aircraft.
* The stall characteristics will need to be checked for the tail as it is necessary for the tail to stall after the wing, so the pitch control surfaces can recover the aircraft to a stable flight configuration. The stall characteristics will be modified by changing the aspect ratio of the tail for suitable stall characteristics.

52. Market Study How to balance ‘style’ versus aerodynamic performance?
Sampled 25 random people on style parameters
Sweep Angle
Aspect Ratio
Dihedral Angle
Taper Ratio
Winglets
Weighted Market Study versus Aerodynamics

53. Aspect Ratio Market Study
Original Design
AR = 6.2
44% votes
Modified Design
AR = 7.6
56% votes

54. Aspect Ratio Performance
Advantages
Increases CL
Decreases CDi
Disadvantages
Earlier stall angle
Increased wing weight
Trade study
Used box beam
Corresponds to AR = 6-7
Chose Aspect Ratio = 6.2

55. Dihedral Marketability
Original Design
Θ = 4 degrees
Modified Design
Θ = 0 degrees
76% votes
24% votes

56. Dihedral Angle Polyhedral Wing Design
Dihedral Angle measured using Effective V-Dihedral Method (EVD)
Wing Taper also affects EVD
Self corrects roll
For Aerobatic aircraft typical
EVD = 0°- 2°
Market Study suggests EVD = 4°
EVD = 3.0°
A = -25°
B = 25°
C = 4.5°
x1 = 0.18 (ratio of wing span)
x2 = 0.32 (ratio of wing span)
Chose EVD = 3°

57. Sweep Marketability 48% votes 52% votes Modified Design
Λ = 0 degrees
Original Design
Λ = 3.0 degrees
48% votes
52% votes

58. Sweep Performance Λ Lift Reduction 3 0.3% 5 0.8% 10 3.0% 20 11.7% 40
41.3%
Chose Λ = 3 degrees

59. Taper Ratio Market Study
Original Design
TR = 0.58
Modified Design
TR = 0.8
96% votes
4% votes

60. Taper Ratio Performance
Brings planform closer to an elliptical shape
Desire λ =
Increases AR, decreases area
Chose λ = 0.5

61. Winglets Market Study
No Winglets
Winglets

62. Winglets Market Study
1.5” 3” 6”

63. Chose Winglet Height = 3 inches
Winglets
Penalty?
Winglet Height
Drag Penalty
1.5"
2.4%
3"
4.5%
6"
8.2%
Chose Winglet Height = 3 inches

64. Parasite Drag K = form factor Q = interference factorK Q Cf
Swet
Sref
CDp
Wing
1.21 1
0.0066
7.10
3.50
0.0162
Horizontal Tail
1.19
0.0074
1.70
0.0043
Vertical Tail
1.18
0.0077
0.60
0.0016
Fuselage
1.06
0.0050
1.57
0.0024
Winglets
0.0073
0.24
0.0012
Components -
0.0251
K = form factor
Q = interference factor
Cf =skin friction coefficient
Swet = wetted area (ft2)
Sref = wing plan form area (ft2)
Internal Antenna
Landing Gear Drag:
Tom’s Slide: parasite drag build up calculation
Summation of all components results in total parasite drag
* parasite drag includes the following drag parameters: pressure (form) drag, friction drag and miscellaneous drag (these drags are present in sectional curve)
* miscellaneous drag includes things such as control surface gaps, base area, or other items that were looked over in the other drag calculations
Miscellaneous Drag: 10% CDo
Schlicting Formula

65. Load Factors Load Factor Analysis V-n Diagram for CLmax Turn-Radius
Load Factor of 2.26 for Loiter Speed
Turn-Radius
60o Bank Angle
Load Factor of 2.0
Turn Radius is ft
Vertical Turn / Loop
Considered for Loiter Speed
Pull-Up : ft
Upside Down : ft
Pull-Down : 8.22 ft
Conclusions
Load Factor of 2.26 chosen
Aerobatics Performed at Loiter Speed or Below
Max Speed for Level Dash Only

66. General Construction Medium Grade Balsa
Wing, Fuselage, & Horizontal Tail
Traditional rib and stringer construction
Balsa leading edge
All stringers are 3/16 in2
Vertical Tails
Solid Balsa
Monokote Covering

67. Mission Analysis Warm up/Taxi 45 seconds , 50% power
96 mA-h
Take off roll/Climb
10 seconds, 100% power
49 mA-h
Loiter
4 min turning, 100% power
1174 mA-h
4 minutes straight, 75% power
880 mA-h
Descend and Land
Unpowered
Landing
Warm up/
Taxi
Take off
Climb
Loiter
Descent
TOTAL = 2200 mA-h

68. Parts Location Number Part Motor Batteries Wheels Aileron Servos
Speed Controller
Gyro
Receiver
Tail Servos
Tail Gear

69. Tail Sizing Roskam method implemented in code
Determined By Coefficient Ratio Equation
Static Margin = 12.63% AC CG

70. Class II Tail Sizing Longitudinal X-Plot Directional X-Plot
Xcg Leg Determined from Weight of Design
Xac Leg Determined by Methods in Roskam Part II, Chapter 11
Directional X-Plot
Determined from Yawing Moment Due To Sideslip Derivative Cnβ
For stable aircraft, Roskam’s book sets Cnβ to per degree.

71. Control Surface Sizing
Determined Using Historical Data
Verified adequate in FlatEarth
Rudder
Aileron
Elevator

72. Aircraft Moments Moment about C.G. Equation:
Cm = Cm0 + Cma*a + Cmde*de
Determine elevator deflection for zero
net moment
Cruise Condition:
CL = de = 0o