1. Critical Design Review David Akerman, Jen Getz, Greg Goldberg, Zach Hazen, Jason Patterson, Benjamin Reese December 4, 2006 PRV (Peregrine Return Vehicle)

2. 2 Presentation Outline Project Objectives and Overview System Architecture Design Elements –Mechanical Design Elements –Electrical Design Elements –Software Design Elements Integration Plan Verification and Test Plan Project Management Plan Questions

3. 3 Requests for Action Flutter Analysis / Control Gains –Open (as of 12/4/06) Manufacturing difficulties –Closed Federal Aviation Administration Requirements –Closed

4. 4 SYSTEM ARCHITECTURE

5. 5 Objective Overview Objective: To provide the Colorado Space Grant Consortium with a reusable vehicle that can return student built science payloads to a selected target.

6. 6 Requirements Overview Combined weight of Vehicle, EOSS telemetry beacon, and payload must not exceed 26 lb –Payload weight 8.3 lb. –EOSS telemetry beacon weight 2.7 lbs –Vehicle structure and subsystems must weight < 15 lbs Vehicle must carry five, 4.7-inch cubical student-built science payloads, weighing 1.65 lb each. –Vehicle must have the necessary volume to accommodate payloads, subsystems, and internal structure. Ground Impact velocity must not exceed 15 ft/sec –Vehicle structure must be durable and resilient to withstand heavy shock loads –Parachute touchdown required (no runway available) Vehicle must be able to land within ¼ mile of an intended target chosen prior to launch.

7. 7 System Design

8. 8 Subsystems –Avionics –Payload Structure –Airframe –Parachute Deployment System –Thermal Control System System Design

9. 9 System Design: Avionics Rack Houses –Auto-Pilot MP2028g –Power regulation board –Thermal control board –Video Overlay Board Provides structural support Easy access for removal

10. 10 System Design: Payload Structure Payload Structure –Secure payloads during flight –Supply support for payload weight of 8.2 lbs. –Provides mounting for avionics rack –Supports front spar

11. 11 System Design: Airframe Provides stiffness Supports and protects payload Provides mounting points for control surfaces

12. 12 System Design: Parachute Deployment System Parachute Deployment System –Deploys parachute using pyrotechnic charge –Triggered by auto-pilot or pressure sensors –Touchdown velocity of 15 ft/s

13. 13 Proto-Peregrine Results Flight Behavior of Flying Wing Stable Configuration Adequate Control Response (Low Altitude) Manufacturing Experience Material Selection

14. 14 MECHANICAL DESIGN ELEMENTS

15. 15 Design to Specification Fuselage –Designed to withstand up to 10 g loads in dive pullout manuver. –Designed to withstand parachute deployment. –Designed to survive an impact of 15 ft/s and be re-usable. Wings –Must withstand 10 g-loads (260 lbf) in pullout manuver. Payload Bay –Designed to acomodate five, 105.4 in 3 cubes (4.73” on all sides). –Support a combined payload mass of 10 lb. –Field of view through the fuselage for each box. –Nadir-pointing in ascension phase Parachute Deployment System –Descent rate of 15 ft/s –Pyrotechnic Parachute Deployment Note: The payload will be contained within the fuselage

16. 16 Drawing Tree

17. 17 Weight Budget ObjectBudgeted Weight (lbs)Current Weight (lbs) Carbon Fiber Spars and Ribs1.51.3 EPP Foam3.5 Payload8.3 EOSS Package2.7 Payload structure1.751.73 Avionics1.5 Skin (Thin Plastic / C.F.)?? Parachute1.751.5 Parachute Mechanism11 Misc (Glue, Tape, etc)?? Total Weight22.0021.53 Remaining Weight4.004.47

18. 18 Overall System and Subsystems

19. 19 EOSS Payload Tracking beacon and flight data collector Required for launch by FAA Provided by EOSS Contains: –Alinco DJ-C5 dual band credit card radio (144 – 148MHz) –GPS Receiver –Basic Stamp Processor –TinyTrak 3.0 APRS encoder –Balloon cutaway device Weight: 2.7 lbs Dimensions: 10''L x 5.6"W x 2.5"H Cannot be taken apart

20. 20 *Formula from Knacke, T.W. Parachute Recovery Systems Design Manual, 1992. Parachute Deployment System Sizing

21. 21 Parachute Deployment System

22. 22 Airframe Design

23. 23 Payload Structure Purpose –Structure to hold payloads –Distribute bending load from the spar –Structure coupled with front spar to mitigate the force exhibited in pullout maneuver –Give the ability for the autopilot and avionics bay to be situated over the glider C.G. Material –Top of Structure made of rigid PVC –All other surfaces and components made of Aluminum 6061 Dimensions –16.06”L x 9.76”W x 5”H Projected weight –1.75 (lbs) Faces attached using #2-56 screws Structure bonded to the inside of the center section

24. 24 Glider Structure *Note: Dimensions in inches

25. 25 Glider Center Section

26. 26 Wings

27. 27 ELECTRICAL DESIGN ELEMENTS

28. 28 Design to Specification Avionics –Autopilot Must control a 20-lb UAV Must withstand high G-loads (parachute deployment, high- G turns/pull-outs) –Controls/Servos Must provide the torque necessary to apply aerodynamic forces at high airspeed (Mach 0.4) –Recovery System Must slow vehicle to safe touchdown speed 15 ft/s or 10.23 mph Must be reusable > 90% proven reliability Must operate independently –Power Supply Must be able to provide reliable voltage and current to Power Distribution Board Must be able to provide 3 A- hrs at 8-14 VDC for avionics excluding servos Servo battery must be able to provide 3.3 A-hrs at 4.8 VDC –Power Distribution Board Needs to provide appropriate voltage and current to different components

29. 29 Overall Electrical System

30. 30 Electrical Subsystems Auto-Pilot System Balloon Release System Parachute Deployment System Servo Controls Thermal Control System Auto-Pilot System

31. 31 Auto-Pilot Provides control for servos Triggers parachute deployment Power: 11.1 VDC 3300 mAh Li-Po Battery MicroPilot On Board Components: Trimble Lassen SQ GPS Receiver 3-axis Accelerometer Barometric Pressure Barometric Airspeed

32. 32 Balloon Release System Current EOSS System –NiChrome wire cut away device Relay connected to standard EOSS package Signal sent by EOSS closes a relay causing the NiChrome wire to burn through the nylon Parachute Cord

33. 33 Parachute Deployment System PIC16F84 allows for user to program deployment altitude. Also performs all logic functions Accepts trigger from either pressure sensors or auto-pilot Trigger causes detonation of pyrotechnic charge Power: 10 VDC Independent battery

34. 34 Servo Controls Futaba FTM0035 –2x Servos Control Elevons –Input from Auto-Pilot –Power: 4.8 VDC supplied from separate servo battery –Torque 89 oz-in –Speed 0.24 sec/60º

35. 35 Thermal System Provides thermal control for all electronics Temperature of avionics must be maintained above 0ºF Ambient temperature at 92,000 ft is -70ºF Must keep above 0ºF Controlled to +/- 2ºF Power: 5.8 VDC 8 Watts

36. 36 SOFTWARE DESIGN ELEMENTS

37. 37 Software Break Down Flight simulation Aerodynamic design Autopilot setup Autopilot simulation Autopilot conditions feed test

38. 38 Flight Simulation U of Wyoming Wind Data Range 40 mi Altitude 90,000 ft Vehicle L/D range Simulation Range covered L/D required to cover range given Time to target Dive trough Jetstream from 45,000 ft to 16,000 ft Heading set towards target Loops dive angle to optimize range Weight Area Cd InputsOutputs

39. 39 Design Analysis Range Requirement: 40 miles w/winds Vehicle L/D: 5-7 necessary predicted at PDR Aspect Ratio Oswald’s Efficiency Factor Zero Lift Drag (parasite)

40. 40 Design Analysis AR Peregrine C Lcruise C D0 L/D required 40.5~0.027 Assuming the above Values, an efficiency of e > 0.4 should provide necessary L/D “Real” Sailplane: e ~ 0.95, typical: e ~0.8

41. 41 Performance Predictions C D0 ~ 0.2-0.3 –Trade Study, Estimations C L cruise = 0.48 –Mission Simulation AR = 4.51 –Aircraft Geometry e = 0.84 –Vortex-Lattice, Trade Studies L/D ~ 9.8 - 12.2

42. 42 Design Analysis Vehicle Must Survive all Flight Regimes and Be Controllable Derived from PDR Requirements Flutter SafetyBending StrengthControl Gains

43. 43 Analysis Limitations Theoretical Flutter Analysis –CFD Model –Representative Wind- Tunnel Model –Flutter Comparison Practical Flutter Prevention –Flight Testing –Autopilot Protection

44. 44 Design Analysis Flutter and Control Response are both dependent on Dynamic Pressure –Dynamic Pressure for a given glider shape in a steady glide depends only on the glide angle. Terminal dive testing will reveal if flutter will occur at any altitude

45. 45 Design Analysis

46. 46 Vehicle Stability Static Margin = 10% –Chosen from experience, trade studies –Set with wing placement Measure of Directional Stability = 5% –Chosen from experience, trade studies –Set with winglet sizing

47. 47 Auto-pilot Setup PC interface Control gains are found Moments of inertia identification routine Define pitch, yaw, descent rates, etc… Autopilot Setup Autopilot is mounted Control gain schedule Wait to climb GPS Navigation Fly to Approach Deploy parachute Flight path programming Launch site GPS lock Initialize Auto-pilot integration software Servo Configuration Elevon controls Defines flying wing Elevon mixing Moments of inertia routine Vehicle tilted in specific angles and let to settle Rates definition Max and min rates desired Pitch, yaw, roll, descent Angle definition Pitch, yaw, roll

48. 48 Autopilot Testing Simulation Mission Simulation Software Included with MP2028 –Horizon mp Provided by Micropilot Specifically designed for Micropilot MP2028 flight simulations Allows Atmospheric and wind data as a simulation input

49. 49 Recovery System Software

50. 50 INTEGRATION PLAN

51. 51 Assembly Flow Diagram

52. 52 Assembly Diagram: Avionics

53. 53 Assembly Diagram: Airframe

54. 54 Payload Structure Assembly

55. 55 Assembly Diagram: Parachute Deployment System

56. 56 Functional Test Plan Master Plan –Avionics Flight Data Recording Flight Control/Navigation System (FCS) Power Distribution Thermal Control –Airframe Payload Bay Complete, “empty” aircraft –(nothing installed except payload bay) –EOSS Beacon –Parachute Deployment System

57. 57 Avionics Testing: Flight Data Recording Video Recording System: –Test DVR capture/playback. –Verify interfaces between Camera, Autopilot Overlay board, and DVR. –Test data overlay and storage. Data Acquisition/Analysis –Recover data from onboard data logger via RS-232 serial port. –Comparison between acquired data and model.

58. 58 Auto-pilot: –Simulate flight situation to verify proper function –Integration test Power Distribution Board: –Connect board to all components, verify proper function of each component Batteries: –Test full discharge time of cold batteries ~ -40F –Test the thermal output of batteries Thermal Control System –Cold-test Avionics Compartment at a temp of ~ -40F in flight condition –Determine time and temperature at equilibrium Ensure equilibrium T is within limits for all components Avionics Testing

59. 59 Mechanism testing EOSS Telemetry Beacon –Physical compatibility test with other systems in the area (PDM, PLB, airframe) –Radio Frequency interference (RFI) test Parachute Deployment System –Canopy test –Pyrodex (Ejection Charge) test –Aircraft Integration Test

60. 60 VERIFICATION AND TEST PLAN

61. 61 Verification and Test Plan Total Vehicle Weight < 26 lb –48 hours prior to launch: Weigh all student payloads, determine ideal (balanced) placement in payload bay Weigh EOSS beacon package Weigh vehicle both empty and loaded (Ready-to- fly configuration)

62. 62 Touchdown Velocity < 15 ft/sec –3 months prior to launch: Parachute will be tested on a dummy 26-lb load (water jugs, etc) to verify that touchdown velocity is actually < 15 ft/sec –Extra time allows for rearrangement of deployment system if necessary –If testing budget allows, we will deploy the parachute via manual RC control at ~1000’ AGL on the full-scale (Flight) model, WITHOUT the autopilot installed. –If no chute deployment, can intervene via RC. Verification and Test Plan

63. 63 Landing Accuracy –Vehicle is NOT required to return to launch site –Landing sites will be in rural Eastern Colorado and will be selected in terms of accessibility. –Landing accuracy will be tested on ½ and full- scale craft operating under autopilot control Program landing site, then drop/fly from highest possible altitude under autopilot control with RC pilot standing by in case of emergency Bungee launch, air drop (helicopter, skydive aircraft) Verification and Test Plan

64. 64 Parachute test –Car test Force at the velocity Cd at the velocity Tension on string Parachute Deployment Test –Test ignition system under different temperatures and ambient conditions –Launch parachute while on a flat spin –With autopilot failure Structural Test –Static loading –Dynamic loading Balloon Release System Test –Temperature testing Verification and Test Plan

65. MANAGEMENT PLAN

66. 66 Organization Chart

67. 67 Work Breakdown

68. 68 Project Risk 1.Power System Failure 2.Parachute Failure 3.Difficulties in Auto-pilot Programming 4.Unrecoverable Flight Situation Including Flutter 5.Auto-Pilot Failure 6.Loss of GPS Signal 7.Electronics Malfunction

69. 69 Manufacturing Schedule

70. 70 Test Schedule

71. 71 Monetary Budget

72. 72 Special Needs and Facilities Balloon Launch Site –Provided by EOSS and Colorado Space Grant Consortium FAA Approval –Given as long as flight includes EOSS package Provides Real Time Telemetry to the FAA

73. 73 QUESTIONS

74. APPENDIX I: SYSTEM ARCHITECTURE

75. APPENDIX II: MECHANICAL DESIGN ELEMENTS

76. 76 Avionics Rack

77. 77 Parachute Deployment System

78. 78 Parachute Deployment System

79. 79 Payload Structure

80. 80 Drawings

81. 81 Drawings

82. 82 Drawings

83. 83 Drawings

84. 84 Drawings

85. 85 Drawings

86. 86 Drawings

87. 87 Drawings

88. 88 Drawings

89. 89 Drawings

90. 90 Drawings

91. APPENDIX III: ELECTRONIC DESIGN ELEMENTS

92. 92 Recovery System

93. 93 Thermal Control System

94. 94 Auto-Pilot System

95. APPENDIX IV: SOFTWARE DESIGN ELEMENTS

96. 96 PID Loops Autopilot Navigation

97. APPENDIX V: INTEGRATION PLAN

98. 98 Payload Capacity: 5 x 4.7-in cubical payloads, 1.65 lb each –1 week prior to launch: Student payloads will be individually fit-checked in a mockup of an individual “payload slot” All payloads will be fit-checked together in the vehicle payload bay to check for interference Must be able to close and seal the payload bay with all five payloads installed in ready-to-fly configuration (switches, hatches, buttons, etc) Verification and Test Plan

99. APPENDIX VI: VERIFICATION AND TEST PLAN

100. 100 Aircraft Integration Test –Determine best mounting arrangement of loaded launch tube with complete airframe –Ensure that riser lines are securely mounted to as many spars as possible, and have a free path of travel during chute ejection and inflation Parachute Deployment System

101. 101 Parachute Deployment System Canopy Test –Drop-check with 26-lb ballast and ~15-ft riser lines, estimate time and vertical drop distance to canopy inflation, check oscillation characteristics –Compare actual (measured) vs. predicted (15 ft/sec) touchdown velocity under canopy Pyrodex (Ejection Charge) test –Verify that a 1-gram charge can pop the chute (tie- stowed to keep folding intact for testing) at least 3 feet out of the launch tube (to avoid line tangling) Increase charge size as required, if necessary –Check for damage to the tube, investigate properties of ABS, PVC, and rocket-tube cardboard (Aluminum?)

102. 102 Complete Aircraft (Mock payloads, no avionics) –Static (+ and -) G-load testing on the wings Sandbags loaded slowly, tip deflection measured Testing to 10 G (Static load of 260 lb, test will be aborted if signs of trouble are noted, and will not AT ALL proceed beyond 260 lb) –Shock load testing “Belly-flop” drop test from 6 feet (arm’s length) and 10 feet (ladder) onto simulated expected landing terrain (grass, hard dirt) –Inspect for damage Parachute “yank” test on main spar area –Drop test with parachute riser lines connected to a fixed beam –Will simulate g-loads encountered during parachute deployment and test skin-spar-wing foam-payload bay interaction Airframe Testing

103. 103 Avionics Testing: FCS Autopilot-Software: –Set up and start basic “.fly” file in HORIZON –Watch the virtual mission take place in real time (on-screen instruments) –Connect aircraft to HORIZON –Repeat above test, verify that servos are moving properly during the test. (Correct elevon mixing should be displayed) –Check HORIZON vs STK-8 predicted and actual behavior

104. 104 Avionics Testing: FCS Autopilot-Navigation/Gain Adjustment (Ground Launch, AFTER complete aircraft integration): –Fly in straight line, max distance in direction of launch –90-degree “L” turn, max distance after one right-angle turn –U-turn after launch, max distance away from launch direction –“Z” turn, max distance after two opposite right angle turns –Fly in straight line, circle about a waypoint (Ideally a good thermal) for as long as possible –Pull-up maneuver –Adjust feedback loop gain and NP location as necessary (moving batteries fore/aft)

105. 105 Autopilot- Hardware: –Measure and inspect MP2028 card to determine: Suitable mounting to avionics rack Best routing of cables/wires/pitot-static lines Best way to protect the autopilot from: –Physical Damage (Landing and Parachute Deployment) –Radio Frequency Interference (if applicable) –Electrostatic Discharge (ESD) from EPP foam body Avionics Testing: FCS

106. 106 Test (same for both batteries): –Charge time from zero to full –Full-discharge time under constant load at 65F (light bulb, heater, etc) –Full-discharge time of cold battery (dry ice equilibrium temperature, ~ -40F) This will help determine the maximum mission duration at altitude, and quantify our battery safety factor Avionics Testing: Batteries

107. APPENDIX VII: MANAGEMENT PLAN

108. 108 Table of Contents System Architecture Mechanical Design Elements Electrical Design Elements Software Design Elements Integration Plan Verification and Testing Management Plan Appendix I: System Architecture Appendix II: Mechanical Design Elements Appendix III: Electrical Design Elements Appendix IV: Software Design Elements Appendix V: Integration Plan Appendix VI: Verification and Testing Appendix VII: Management Plan