NASA SL Critical Design Review

Transcription

1 NASA SL Critical Design Review University of Alabama in Huntsville 1

2 LAUNCH VEHICLE 2

3 Vehicle Summary Launch Vehicle Dimensions Fairing Diameter: 6 in. Body Tube Diameter: 4 in. Mass at lift off: 43.8 lbm. Length: 103 in. Concept L-Class Solid Commercial Motor Rover Delivery Electronic Dual Deployment Fiberglass Airframe 3

4 Vehicle CONOPS Deploy Drogue: 19 seconds 5,429 ft. Powered Ascent: seconds 0 1,190 ft. Deploy Main: 62 seconds 600 ft. Landing: 121 seconds 0 ft. Deploy Rover: Team Command 11/3/2017 USLI PDR 4

5 Vehicle Summary Tracking/Rover Deployment Avionics Rover Piston Main Parachute Recovery Avionics Drogue Parachute Fins (x4) CG 56 in. CP 69 in. Payload Fairing 36 in. Forward Airframe 30 in. Coupler 12 in. Aft Airframe 42 in. 5

6 Upper Airframe Overview Upper Airframe houses the rover, piston ejection system, and GPS tracker 6

7 Nose Cone 6 in. elliptical shape; 6.17 in. OD; ABS Plastic; 3-D printed in-house 1.75 in. shoulder; shear pinned to fairing 0.25 in. Aluminum bulkhead 7

8 Fairing Houses payload and deployment device Fiberglass; 6.17 in. OD, 6 in. ID Shear pinned to nose cone; bolt connection to transition 1/14/2018 8

9 Transition Three piece design, two 3D printed ABS plastic, one 0.5 in. thick aluminum bulkhead Each piece has holes for threaded inserts Held together using ¼-20 and bolts Forward Insert Aft 9

10 Transition Three piece design allows for a 57% reduction in weight Max stress on aluminum bulkhead: ksi Yield stress: 42 ksi 10

11 Transition Coupler Connects 4 in. body tube to the 6 in. fairing U-bolt for recovery harness attachment point Shear pins connect to 4 in. body tube Threaded rod with hex nuts for connection to fairing 11

12 Piston Overview CO 2 Powered 12 gram cartridge Spring driven spike used to release stored gas 60 lbf. test monofilament fishing line used as arming tether for spring Hot wire cuts tether to release spring Two main components: piston head and CO 2 housing 12

13 Piston Head Ejects rover and nose cone Fiberglass coupler with aluminum bulkhead 13

14 CO 2 Housing Houses CO 2 cartridge and release mechanism 3D printed ABS Plastic Allows for easy and quick modification upon testing results 14

15 Piston Configuration CO 2 housing positioned in transition shoulder Mounted to side using 3-D printed brackets and 4-40 bolts Keeps housing fixed during flight 15

16 Aft Subsystem Overview Aft Subsystem Components Fin Can Fin(s) Thrust Plate Recovery Bulkhead & U-Bolt Motor/Motor Case Motor Retention Ring 16

17 Fins Trapezoidal Fin Set (4) Maintain stability G10 Fiberglass Great strength/weight ratio 3/16 thickness Flutter Speed Calculated to be Mach ( ft./sec) 17

18 Fin Interface 4 bolts perpendicular to fin face 6 bolts normal to body tube to hold shape Also used to hold fin can in vehicle Entire assembly can be removed 18

19 Fin Can Assembly Overview Consists of: Fin Can, Motor Retention Ring, Thrust Plate, and Rail Button Press fit nut 19

20 Exploded View of the Fin Can Assembly 20

21 Fin Can 3D Printed in house Material: ABS plastic Purpose: Fin retention and centering of the motor Attached to the body tube using 4-40 bolts which maintain the shape of the Body tube 21

22 Fin Can Dimensions 22

23 Thrust Plate Machined in house Material: 6061 Aluminum Purpose: Transfer motor thrust to the airframe Attached to the fin can using the motor retention bolts Part was added due to concern of shearing the Fin Can while during motor burn 23

24 Thrust Plate Dimensions 24

25 Motor Retention Ring 3D printed in house Material: ABS plastic Purpose: Motor retention Attached to the fin can using the motor retention bolts Aft retention was chosen due to the difficulty of disassembling the forward retention system 25

26 Motor Retention Ring Dimensions 26

27 Motor Selection Aerotech L1420R-PS Best met altitude target Avg. Thrust: lbf. Burn Time: 3.2 sec Motor Aerotech L2200 Aerotech L1420 Aerotech L1520 Altitude 6107 ft ft ft. 27

28 Dimensions OpenRocket Flight Simulation Total length 103 in. Wet mass lbm. CP location in. CG location in inches inches 28

29 Stability Margin Motor Burnout (3.28 cal.) Apogee Initial Stability (2.22 cal.) 29

30 OpenRocket Flight Simulation Attribute Value Apogee (ft.) 5429 Length (in.) 103 Max. Mach Number 0.60 Rail Exit Velocity (ft./s) 60.6 Static Stability (cal.) 2.22 Motor Designation AT L1420R P Thrust-to-Weight Ratio /3/2017 CG 56 in. CP 69 in. USLI PDR 30

31 OpenRocket Flight Simulation Apogee (18.62 sec.) Motor Burnout (3.27 sec.) Main Deploy (62.39 sec.) 31

32 Full Scale Monte Carlo Simulation 1-D method used to verify OpenRocket sim Goal: Determine uncertainty in projected altitude Randomly varies conditions by a percentage drag coeff., vehicle mass, propellant mass, case mass Varied between ±6.25% and ±2.5% Use drag coefficient from subscale flight C d = ,000 flights per simulation 32

33 Full Scale Monte Carlo Simulation Mean: feet Median: feet Std. Deviation: feet Max Altitude: feet Min. Altitude: feet 33

34 Central Subsystem Overview Central Subsystem responsibilities: Coupler between airframes Flight Avionics Ejection System Tracking and Ground Station Recovery System 34

35 Drift Analysis Monte Carlo Drift Model Assumes: Apogee is directly above the launch rail The parachute does not open immediately The drift distance stops once a component lands Horizontal acceleration is solely based on relative velocity Drogue parachute is negligible once the main is fully deployed V wind V relative Wind Speed (mph) OpenRocket Drift Distance (ft) CRW Model Drift Distance (ft)

36 Recovery System Calculations Requirement: No individual section will have a kinetic energy greater than 75 ft.-lbf. upon landing Terminal velocity under drogue: ft./sec. Terminal velocity under main: ft./sec. Vehicle Section Mass (lbm.) Fairing Coupler Aft KE (ft.-lbf.) 36

37 Recovery System Drogue Parachute Deployment: Deployment at apogee Fruity Chute CFC-18 (Cd = 1.5) Shock Cords: 1 inch Nylon (50 ft) Connected between forward motor retention bulkhead in lower airframe and avionics bay housing. Descent speed under drogue: ft/s Main Parachute Deployment: Deployment at 600 ft above ground level Fruity Chute 96 Iris Ultra (Cd = 2.2) Shock Cords: 1 inch Nylon (50 ft) Connected between fairing bulkhead and avionics bay housing. Descent speed under main: ft/s 37

38 Avionics Recovery Avionics Subsystem 2 PerfectFlite StratoLoggerCF altimeters; each with an independent 9V battery and pull pin + SPDT momentary activation switch 4 Safe Touch terminals, e-matches, and black powder charges Full redundancy in avionics and ignition 38

39 Coupler Batteries RBF Switches Screw Terminal Strip Flight Computer Charge Well U-Bolt 12 in. 39

40 Recovery Deployment Avionics Normally Closed SPDT Pull Pin Microswitch Prevents ignition during assembly Helps preserve battery life Primary Drogue charge fired at apogee Secondary fired one second after Primary Main charge fired at 600 ft. Secondary fired at 550 ft. Primary charges contain 4 g. of black powder Secondary charges are 2 g. larger than primary 40

41 System GPS Tracking & Rover Deployment Subsystem CRW will use a custom PCB that contains an Xbee Pro-PRO 900HP RF module, Teensy LC, and MTK3339 GPS Chip Xbee transmits GPS coordinates to a receiver connected to the ground station laptop GPS sentences are parsed and written to file for flight data Rover Deployment Electronics operated via XBee Structure Integration 3D printed mount to secure tracker and deployment electronics PCB within transition section of the rocket Three axis security and battery retention to ensure components are kept intact 41

42 Subscale Design Scaling Factors: Geometry of the design Average Thrust of Motor and Thrust Curve Kinetic Energy 42

43 Subscale Flight Results Successful recovery of all three subscale flights Altimeters ignited the black powder charges at the correct altitudes 43

44 Subscale Flight Results Flight 1- Apogee 2884 ft., some weathercocking Flight 2- Apogee 2323 ft., severe coning Flight 3- Apogee 3165 ft., vertical flight 44

45 Subscale Drag Coefficient Using data gathered from the altimeters, the drag force and coefficient for the vehicle were found A = Area of the exposed section, ft 2 ρ = density of the air, lbm/ft 3 C d = Coefficient of Drag u = Velocity, ft/s m = mass, lbm A = acceleration of the vehicle, ft/s 2 g = acceleration of gravity, ft/s 2 Using a weight of 6.33 lbs, an acceleration of ft/s, an A of ft 2, a ρ of lb/ft 3, and a velocity of ft/s: C d =

46 Payload Diameter: Deployed 16.2 in., Integrated 5.7 in. Rover Length: 14.6 in., Chassis Length: 12 in. 46

47 Payload Integration Rover fits inside the piston, which ejects it from the fairing CO 2 cannister pushes rover through nose cone 47

48 Payload CONOPS 1. The rover is ejected from the rocket 2. Wheels deploy and rover moves 10 ft. 3. Rover stops and deploys solar panels

49 Chassis Stores and protects tray of rover electronics 12 in. x 4 in. x 3 in., Aluminum 6061-T6 Machined from single Aluminum block Connects to motors, electronics tray, and solar panel lid 49

50 Wheel Spokes pulled by springs to expand wheel Wheel hub and spokes CNC milled aluminum Integrated Diameter: 5.7 in. Deployed Diameter: 16.2 in. Spoke 6 in. x 0.5 in. x 0.25 in. 50

51 Stabilizing Arm Used to keep chassis upright during deployment 3D printed ABS 11 in. x 0.25 in. x 0.5 in. Mounts to chassis using a hinge Torsion spring pushes out after deployment 51

52 Strength Check Notes This table details the normal load cases for each structural component The wheel hub is the weakest component but can withstand a 73% load increase Part Load Case Safety Margin Chassis 210 lbf (sidewall) Chassis 210 lbf (base) Wheel Hub 210 lbf (sidewall) Wheel Hinge 105 lbf (each) Spoke 210 lbf (lengthwise force) Spoke 7 lbf (Drive force)

53 Electronics Tray Rover electronics contained inside chassis Tray designed to wire and organize electronics outside chassis Tray lowered into top of chassis once assembled 53

54 Electronics Tray Designed to avoid interference with motors Tray Assembly: 11.6 in. x 3.8 in. x 2 in. 54

55 Rover Lid with Mechanism Lid is closed during rover travel for protection Gear slides top lid out to reveal solar panel in chassis 3D printed ABS lid, gear bought from McMaster-Carr 12 in. x 4 in. x in. when closed 12 in. x 7.25 in. x 0.5 in. when open 55

56 Rover Mass Budget The mass of all components totaled 6.6 lbm. A 6% margin was added to the total weight to account for fasteners and adhesives Component Mass (lbm.) Chassis 2.0 Wheel Assembly 2.4 Lid/Solar Deployment 0.7 Tail 0.2 Electronics 1.3 6% Margin 0.4 Total

57 Payload Power Budget Part Name Current (ma) Voltage (V) Adj Current Duty Cycle (%) Time (hr) Total (mwh) Arduino Mega Camera GPS IMU Pressure/Temp Wheel Motors Lid Motors Radio transmit Radio idle Datalogger Power required Part Name Current (ma) Voltage (V) Adj Current Duty Cycle (%) Time (hr) Total (mwh) Li-Ion Battery N/A Power Supplied Power Supplied mhr Power required mhr Factor of Safety mhr 57

58 Electronics Block Diagram 58

59 Electronics Failure Path Emphasizes dependence of each lower level component on the component above it 59

60 SAFETY 60

61 Safety Overview Training and communication are key to maintain safety and avoid mishaps Priorities in CRW safety program (in order of importance): 1. Safety to personnel 2. Safety to facilities & permanent systems 3. Safety to flight hardware & objective success Established SOP and regulations to maintain safety practices Team is transitioning from designing to manufacturing and testing 61

62 Communication CRW team meets twice weekly Safety briefings are held to update the team with pertinent information All conducted tests have documentation of results and lessons learned Documents and test results are recorded to the team s server for ease of access 62

63 Standard Operating Procedures Philosophy Standardization of processes Address risks and hazards with proper method Creation Based on previous versions In collaboration with team leads to adapt SOP steps to the features and mission needs of the Vehicle and Payload Approval Reviewed and approved by Red team members and faculty advisor Implementation: Use latest version Safety Monitor to ensure strict adherence to steps and safety aspects 63

64 Launch and Assembly Procedures Final rocket assembly procedures for the Full Scale have been developed to fit the design concept Minimum assembly or modification of airframe at field Field operations are limited to subsystem integration and loading of energetics Simulated runs of launch procedures will take place at least one week prior to any launch 64

65 Environmental Factors Factors affecting launch vehicle and payload Sudden high winds Humidity Extreme temperatures Terrain Mitigations established: Minimum exposure to environment Constant monitor of the weather 65

66 Environmental Factors Factors affecting the Environment and Local communities Hot exhaust Landing in trees, difficult terrains Landing on infrastructure and private properties Waste from manufacturing and launches Mitigations Established: Inspection and understanding of launch field Waste collection and proper disposal Constant monitor of wind conditions 66

67 Upcoming Trainings Training Activity Date Red Cross First Aid CPR/AED/FA Completed Basic Emergency Procedures Completed Process Hazard Analysis Completed Safe Testing Procedures Completed Root-Cause Analysis Completed Outreach Safety Procedures Completed Sub-scale Launch Safety Procedures Completed Hazardous Material Handling/Disposal Completed Fire Extinguisher training Completed Workshop Safety Briefings 1/23/2018 System Ground Tests Briefings 1/30/2018 TBD TBD Safety Briefings are held based to relevant safety topics. 67

68 Test Plan Test Plan changes since PDR Completed tests includes the subscale launch and subscale charge test. New tests planned for Rover and Launch Vehicle fairing systems. GPS test is on going to ensure constant compatibility. 68

69 Test Plan Test Number Test Type Test Status T1 Subscale Ejection Charge Test Test has been conducted prior to the subscale flight on Test shows that rocket has to go drogue-less and use only one shear pin on both main and drogue for successful recovery. T2 Subscale Flight Successful launch and recovery Vehicle did not reach initial altitude prediction T3 GPS tracker range and capability/telemetry Tracker currently Exhibit poor performances. Team is currently learning how to trouble shoot issues with tracker. Telemetry test is planned for Feb T4 Fin Can Load Test Test will be planned for the end of January to the early February before the full scale launch. 69

70 Test Plan T5 Rover Piston Deployment test Test will be scheduled in February when the piston is manufactured. T6 Fairing Vibration Test Test is planned for middle to end of February once test articles arrive T7 Faring Drop Test Test is planned for middle to end of February once test articles arrive T8 Fairing Transition Compression test Test will be conducted once FEA results shows doubts in the structures. T9 Rover Operational Test Test will be planned and carried out when rover is constructed. T10 Full Scale Charge Test Test will be conducted approximately one week before the first full scale launch date T11 Full Scale Flight Flight will be held on Feb 17 and 18 70

71 TESTING AND REQUIREMENTS VERIFICATION 71

72 Verification Reports Document template for tracking requirements verification Allows for all 4 methods to be tracked Place to record test procedures, personnel, and results Template is in Critical Design Review Appendix 72

73 General Requirements Verification Expected/In-Progress Verification Reports Review of project plan and procedures Review of all submitted documents, website, and teleconference setup Review of Educational Outreach Reports Demonstration of reusability through full-scale flight 73

74 Test Plan Test Number Test Type Description Test Status T1 GPS tracker o The GPS tracker of the launch vehicle and the Tracker currently range and payload will be tested inside of their respective Exhibit poor capability fairing/compartment. This is to ensure that the GPS performances. can reliable transmit and receive signals. Team is currently o The test will also be conducted in obstacles such as learning how to trees and buildings to reveal the limits of the GPS. trouble shoot o The full test of GPS system performance and reliability will be the subscale launch issues with tracker. o Single component tests (radio, GPS receiver), can be done by a team member without supervision. o Subscale launch tests will adhere to SOP. T2 Electrical Charge on E-matches o o o o Spectrum analysis will be conducted to determine if transmission waves will enter into the avionics coupler and affect the electronic components The tracker can be placed inside the coupler to determine how much transmission power exits. The idea is if excessive power exits the coupler, an excessive amount can enter. Shielding can then be implemented based on the results. This test will require more than one team member. However, Red team members and the mentor will not be required for this type of test. Test is has not been planned. 74

75 Test Plan T3 Altimeter Test o The functionality of the altimeter will be evaluated with Test will be scheduled the Charger Rocket Works altimeter testing container. when altimeter has o Only applied for in-house made altimeters. Third party altimeters like Statologger will not require testing. been created. T4 Ejection Charge Test o This is to experimentally verify the correct amount of black powder to be used in the ejection of the drogue and main parachutes. o o An SOP has to be developed for this test This test is dangerous and only Red Team with the presence of the mentor can conduct the test. Test has been conducted prior to the subscale flight on Test shows that rocket has to go drogue-less and use only one shear pin on both main and drogue for successful recovery. T5 Rover Piston Deployment test o Experimentally verify the functionality of the rover deployment mechanism. o The test requires no pyrotechnics so anyone in CRW can conduct the test. Test will be scheduled in February when the piston is manufactured. T6 Fairing Transition Compression test o Experimentally verify the compression strength of the fairing transition o Only the section in doubt from the FEA results shall printed for test. o Currently planned to be a destructive test T7 Rover Terrain Test o The Rover, once constructed, shall be put through its paces in different terrain conditions (except water and mud). o Test is to verify the spoke wheel design. Test will be conducted once FEA results shows doubts in the structures. Test will be planned and carried out when rover is constructed. 75

76 Full Scale Budget Budget Summary Airframe $ Electronics $ Recovery $ Motors $ Rover Structure $ Rover Electronics $ Total Cost $

77 On the Pad Budget Launch Vehicle Airframe $ Electronics $ Recovery $ Motor $ Rover $ Total $

78 11/3/2017 USLI PDR 78