NASA SL Preliminary Design Review

Transcription

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

2 Mission Summary Design, fabricate, test and fly a rocket and payload to 1 mile in altitude Deploy a rover upon landing to autonomously travel and unfold solar panels Conduct STEM outreach with students *Throughout the presentation, all dimensions are in inches 2

3 VEHICLE DESIGN 3

4 Vehicle Summary Launch Vehicle Dimensions Fairing Diameter: 6 in. Body Tube Diameter: 4 in. Mass at lift off: 39.7 lbm. Length: 96 in. Concept L-Class Solid Commercial Motor Rover Delivery Electronic Dual Deployment Fiberglass Airframe 4

5 Vehicle System Locations Tracking/Rover Deployment Avionics Rover Piston Main Parachute Recovery Avionics Drogue Parachute Fins (x4) CG 51 in. CP 63 in. Payload Fairing 36 in. Forward Airframe 24 in. Coupler 12 in. Aft Airframe 41 in. 5

6 Vehicle CONOPS Deploy Drogue: 19 seconds 5,282 ft. Powered Ascent: seconds 0 1,050 ft. Deploy Main: 50 seconds 600 ft. Landing: 100 seconds 0 ft. Deploy Rover: Team Command 6

7 Flight Simulation OpenRocket Sim: A 1-D in house Monte Carlo simulation will be used to verify results Results will also be compared to flight tests for verification Attribute Value Apogee (ft.) 5282 Length (in.) 96 Max. Mach Number 0.56 Rail Exit Velocity (ft./s) 55.7 Static Stability (cal.) 2.0 Motor Designation AT L1520T - P Thrust-to-Weight Ratio 8.7 CG 51 in. CP 63 in. 7

8 Simulation Results Apogee of approximately 5282 ft. at 19 sec. Motor burnout at approximately 1050 ft. at 3.2 sec. Burnout at 3.2 sec. Apogee of 5282 ft. (19 sec.) 50 sec. Main deploy (600 ft.) 8

9 Stability Analysis Stability of 2.06 cal. at rail exit Calculated with no wind conditions Stability of 2.74 cal. at motor burnout Maximum Stability: 2.74 Takeoff Stability:

10 UPPER AIRFRAME Nose Cone Payload Fairing Transition Forward Body Tube 10

11 Objectives Forward System Overview Protect and deploy the payload House assembly for tracking vehicle location Transition upper airframe to payload fairing Payload Piston Avionics Bay 11

12 Nose Cone and Fairing 6.0 in. 2.0 in. 6.0 in. 3D printed High Strength ABS Ejected with rover deployment Room to store ballast for stability No electronics housed inside Shear pin interface Bulkhead at base 6 in. ellipsoid shape 2 in. shoulder 24.0 in. Responsible for housing the rover and rover deployment system Filament wound fiberglass 12

13 Piston Overview Used to deploy rover from fairing Spring driven spike punctures cartridge Spring released by hotwire upon command; redundant arming Plunger Ø 6.0 in. Cylinder Ø 6.0 in. Machined from aluminum Powered by 8 or 12 gram CO 2 cartridge Plunger tethered to base Standard Operating Procedure in development 13

14 Fairing Transition Aerodynamic transition between upper airframe and fairing, load path supplemented with aluminum insert 3-D printed with ABS plastic, single piece design Threaded rod in tension connecting to aft bulkhead to built in forward coupler 14

15 Fairing Transition Problems with ABS single piece design Mass: 4.7 lbm Complicated FEA Structurally weak without aluminum insert Aluminum insert could pose manufacturing difficulties Other options considered: Aluminum brace with direct bulkhead connection, purely aerodynamic cover 15

16 CENTRAL SUBSYSTEM 16

17 Central Subsystem Overview Central Subsystem responsibilities: Primary coupler between airframes Flight Avionics Ejection System Tracking and Ground Station Recovery System 17

18 Coupler U Bolt (2 Places) Aluminum Bulkheads Stratologger CF Altimeter (2 Places) 9V Battery (2 Places) Switch/Pressure Equalization holes (2 Places) 9 in. 1 in in. All-Thread (2 Places) 3D Printed Avionics Sled 1 in. Switchband Black Powder Housing (4 Places) 18

19 Avionics Recovery Avionics Subsystem 2 PerfectFlite StratoLoggerCF altimeters; each with a 9V battery and SPDT momentary activation switch 4 Safe Touch terminals, E-matches, and black powder charges Full redundancy in avionics and ignition 19

20 Recovery Deployment Avionics Normally Closed SPDT Pull Pin Microswitch Prevents detonation during assembly Helps preserve battery life Primary Drogue charge fired at apogee Secondary fired one second after Primary Main fired at 600 ft. Secondary fired at 550 ft. Primary charges are roughly 4 g of black powder Secondary charges are 2 g larger than primary 20

21 GPS Tracking Subsystem System CRW will reuse a previously designed PCB that contains an Xbee Pro- PRO 900HP RF module, and an Antenova GPS Chip PCB will includes traces for all relevant connections including battery sources. Xbee transmits GPS coordinates to a receiver connected to the ground station laptop. Tests will be performed prior to the full scale launch to verify operation success Structure Integration 3D printed mount to secure tracker and its essentials within the transition section of the rocket. Three axis security and battery retention to ensure components are kept in tact 21

22 Recovery System Drogue Parachute Deployment: Deployment at apogee Fruity Chute CFC-18 (C D = 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: 62.2 ft/s Main Parachute Deployment: Deployment at 700 ft. above ground level Fruity Chute 60 in. Iris Ultra (C D = 2.2) Shock Cords: 1 inch Nylon (50 ft.) Connected between fairing bulkhead and avionics bay housing. Descent speed under main: ft/s Open Rocket Simulation between 0 and 20 mph winds showed a maximum drift at 15 mph of about 1,700 ft. 22

23 Recovery System Calculations Required that each individual section will have a maximum kinetic energy of 75 ft-lbf For initial calculations, a conservative estimate of 75 ftlbf was used for the heaviest section KE = 1 2 mv2 m = mass of the section, lbm v = velocity, ft/s The largest independent section is 15 lbm, so the safe descent speed was determined to be 17.9 ft/s D = 8mg πρc D v 2 D = diameter of parachute, ft. m = mass of vehicle, lbm g = force of gravity, ft/s 2 ρ = density of the air, lbm/ft 3 C D = Coefficient of Drag v = previously calculated velocity, ft/s Minimum Diameter must be 93.3 inches 23

24 Load Path (Drogue and Main) 3 2 The load in 1, 2, and 3 are causing tension under Drogue and Main. Shock cord applies load to eyebolt in the coupler bulkhead. The load in 4 is transferred through the all thread and down to the motor casing then back up the tube. 1 4 represents force due to drag represents the force due to mass 24

25 AFT SUBSYSTEM 25

26 Aft System Objectives Objectives/Responsibilities Fin Design Optimize dimensions and materials for flight stability Centering Ring/Thrust Plate Carry load path from the vehicle Centering and fin integration ability Forward/Recovery Retention Provide method for recovery attachment Carry thrust through the vehicle via forward retention 26

27 Aft System Components Design Overview Through the wall design/slotted body tube Slots allow for fin mounting integration G10 Fiberglass fins attached with seven 4-40 bolts per fin Fins will be mounted to centering ring 3-D printed centering ring/fin mounting bracket Can be removed from body tube for repair/inspection Aluminum Forward/Recovery retention bulkhead Uses U bolt for recovery system Motor case tapped to allow for forward retention Trapezoidal Fin(s) (4) Forward/Recovery Retention Bulkhead Secondary Centering Ring Motor/Motor Casing Fin Can/Centering Ring Thrust Ring 27

28 Motor Selection Other motors considered: L1150 Too little total impulse L850 Too slow off the rail L1390 Too much total impulse Aerotech L1520R-P Specifications Motor Designation L1520T-P Apogee 5,282 ft. Stability 2.0 cal. Ballast 51 in. Diameter 75 mm. (3 in.) Length 25.7 in. Propellant Mass 8.0 lbm Total Impulse 835 lbf.-s Max Acceleration 289 ft./s 2 Velocity off the Rail 55.7 ft./s Burn Time 2.5 sec 28

29 Motor Retention Forward Retention Bulkhead Screwed onto top of motor Recovery retention is fixed on U-bolt 3.9 in. diameter 0.5 in. thick Aluminum Fixed to body tube with four ¼-20 screws 29

30 Fin Can Requirements fulfilled by part: motor centering, fin mounting, thrust takeout from motor Material: 3D printed high strength ABS plastic Location: inserted in the bottom of the aft body tube Fin Can 30

31 Fin Can Dimensions 31

32 Secondary Centering Ring Purpose: align motor as it is inserted into the rocket Bolted to the aft body tube using 4-40 bolts Material: Polycarbonate 32

33 Fin Design Trapezoidal Fin Design Allows more freedom in fin design Adjust fin shape to shift CP Fin Dimensions 8 in base 3.5 in height with extended base for body tube insertion Seven holes allow integrated mounting to centering ring located inside body tube Rounded leading edge Fin Material G10 Fiberglass Will be fabricated/designed in house Fin Mounting Fins mounted through the body tube to centering ring Replaceable upon breakage/damage Flutter speed Calculated to be mph (Mach 1.88) 33

34 Fin Material Fins made out of G-10 fiberglass This material was chosen for its high strength to weight ratio Tensile Strength: Crosswise: 38 ksi Lengthwise: 45 ksi Flexural Strength: Crosswise: 65 ksi Lengthwise: 75 ksi Flexural Modulus: Crosswise: 2400 ksi Lengthwise: 2700 ksi Compressive Strength: 65 ksi Its density is lbm/in^3. 34

35 Fin Retention Each fin mounted with seven 4-40 bolts; normal to fin face Four sets of ten 4-40 bolts normal to body tube surface used to maintain body tube shape under motor thrust 35

36 Subscale Rocket 3.0 in Subscale Rocket 6.0 in Full-Scale Rocket Approximately half-scale 4 in. body in. body 6 in. fairing 3 in. fairing Mach in. motor 1.5 in. motor 96 in. length 49 in. length 9.0 G 15.2 G 36

37 PAYLOAD DESIGN 37

38 Payload Summary Objective: Design an autonomous rover that will deploy from the interior of the rocket, move a minimum of 5 ft. away from the rocket, and deploy solar panels The rover s design consists of a rectangular chassis, two expandable wheels, and a stabilizing arm The rover measures temperature, pressure, location, and transmits this data with images to a ground station 38

39 Rover Assembly The tail will be wrapped around the chassis while inside the fairing. Rover will be kept collapsed passively by the fairing. The collapsed diameter is 5.7 in with 0.15 in of clearance. 39

40 Rover Assembly Rover wheels will expand to in. diameter when deployed Wheels rotate independently. Allows for steering via differential Lid will slide open via linear gear driven by a DC motor Solar panel will increase its effective area from 0 to 100% Solar panel will charge battery for distance extension 40

41 Rover Chassis Trade Study Aluminum Unibody 3D Printed ABS Unibody Aluminum Base/3D Printed ABS Walls Ease of Manufacturing Strength to Weight Environmental Protection Total Score Aluminum Unibody selected Highest strength to weight design Resistant to drastic changes in temperature Least deflection under load protects motors and electronics 3D Printed ABS Unibody is secondary selection Will be used if aluminum unibody is too difficult to manufacture 41

42 Rover Chassis Design The chassis will be milled out of a single block of 6061-T6 aluminum The chassis will house all electronics The drive motors will be mounted directly to the sidewalls The tail will be mounted to the bottom of the chassis 42

43 Rover Chassis Stress Analysis The chassis can sustain a 30G (210 lbf) load to the sidewall, simulating a load from the wheel during adverse deployment conditions (left) The chassis can sustain a 30G (210 lbf) load to the base, simulating loading from inside the rocket upon landing (right) 43

44 Rover Tail Trade Study Ease of Manufacturing 18 inch Measuring Tape (Wrapped around) 11 inch Sideways Hinged Aluminum Tail 5 2 Strength 3 5 Tail Length 5 3 Total Score Wheel Rotation Counter moment from tail Measuring Tape selected Ease of manufacturing Results in a longer tail and moment arm Sideways Hinged Aluminum is secondary selection Will be used if measuring tape fails integration and deployment tests 44

45 Main Goals for design: Expanding wheels 6 in. diameter constraint while inside rocket > 6 in. diameter desired for handing terrain Chosen Design: Umbrella wheel Desired for handling terrain Trade Study: Wheel Design All designs similarly decent in other categories Telescoping Wheels Foam Umbrella wheel Cost Design Complexity Low Risk of Damage Terrain Effectiveness Total

46 Main Considerations: Pushing wheels out of rocket without taking damage Ease of manufacturing wheel shapes Chosen Material: Aluminum Highest strength while maintaining low weight Easiest to manufacture wheels Trade Study: Wheel Material Aluminum ABS Polycarbonate Cost Design Complexity Weight Strength of Material Total

47 Wheel Design Current Chosen Design: Umbrella Wheel lbm 5.7 in. diameter wheel expands to 14 in. diameter wheel Linear extension spring for compression and expansion Keeps compressed while in rocket, expands naturally once out Spring located on the exterior, pulls in to bring spoke vertical Rod used for assembly of main wheel to spokes 47

48 Wheel Design Main Wheel Made of Aluminum 6061 T6 Eight notches for eight spokes, holes for attaching spokes with rod 48

49 Wheel Design Spoke 6061 T6 Aluminum 0.75 in. extrusion for grip with expanded wheel Circular piece for attaching to wheel base 49

50 6061 T6 Aluminum Attaches to wheel base Wheel Design Motor Mount 50

51 Main Wheel Can withstand 120 lbf before yielding Load: Pushed out by piston, no more than few pounds Will likely be more distributed to entire wheel base 51

52 Spoke Can withstand 35 lbf before yielding to 40 ksi Max Stress 15 ksi Full weight of rover and motor torque 52

53 Motor Mount Can withstand 900 lbf before yielding Load: Pushed out by piston, absorbed by other parts Max load by piston no more than a few pounds 53

54 Solar Deployment Sliding solar panel lid utilizing remote servo gear Solar panels will remain static Solar panels recharge battery Gear System Hinge Simplicity 3 4 Functionality 5 3 Weight 2 2 Cost 1 1 Total Score Two different designs considered for solar panel deployment mechanism Gear system lid Hinged lid Lid with gear system was selected Hinge mechanism would be harder to close once opened Gear system would be easier to bring the cover back over the solar panels 54

55 Rover Mass Budget Component Mass (lbm) Chassis 2.5 Wheel Assembly 1.4 Lid/Solar Deployment 1.0 Tail 0.1 Electronics % Margin 0.6 Total 7.0 The mass of all components totaled 6.4 lbm. A 10% Margin was added to the total weight to account for fasteners, adhesives, and design changes 55

56 Battery Trade Study Three different batteries considered 3x Li-Ion in series 4x CR123a Surefire in series 8x Energizer Recharge Power in series Trade studies conducted by rating each battery s benefits on a scale of 1 5 Li-Ion was selected based on criteria Li-Ion CR123A Surefire Energizer Recharge Power Plus Power Capacity Weight Safety Reusability Power Density Total

57 MCU Trade Study Arduino Mega Arduino Uno PCB with ATMega 2560 Beaglebone Raspberry Pi 3 Clock Speed I/O Pins Operating Voltage Power Draw Complexity Volume Mass Cost Total

58 Component Selection Component Selection Features MCU Arduino Mega (7 12) Vin I2C, SPI, UART, GPIO 16 MHz IMU Temperature and Pressure Sensor Motor Adafruit LSM9DS0 Adafruit BMP280 Cytron DC Geared Motor SPG30-300K Accelerometer Gyroscope Magnetometer 3 Axis I2C, SPI Press range: ( ) hpa Temp range: (-40 85) C SPI, I2C 0.8" x 0.7" x 0.1" 12 V At load 410 ma Stall torque 1.18 Nm Mass: 160 g Brushed 58

59 Selection cont. Component Selection Features Solar Cell GPS Radio Lid Motor DC/DC converter OSEPP Monocrystalline Solar Cell Adafruit MTK3339 X-Bee PRO NMB Technologies PPN7PA12C1 LM3671 Buck Converter 100mA 5 V 4 x 3 x 0.2 5V 20 ma 10 Hz updates -165 dbm sensitivity 28 mile range (with high gain antenna) 900 MHz Data rate 200 kbps UART, SPI 5V DC Brushed lbm 3.3 V output 600mA draw 0.6" x 0.4" x 0.1" Camera ArduCam CMOS OV VGA 3.3V supply needed 59

60 Power Budget Required battery capacity = I ma V V DC Time hr Efficiency 11.1 V Component Current (ma) Voltage (V) Time (hr) Duty Cycle (%) Efficiency (%) Necessary Capacity (mahr) Arduino Mega Pressure/ Temp IMU Wheel Motors Lid Motors Radio Camera GPS Voltage Regulator Required Capacity (mahr) Available Capacity (mahr) Safety Factor

61 Component Block Diagram 61

62 Payload Software Flow Diagram Remove RBF; Payload detects launch via acceleration Takes acceleration data throughout flight, calculates changes Once acceleration is zero for several iterations, waits for deployment signal from ground station Rover transmits acknowledgement, waits for confirmation signal Receives confirmation signal; Delays 30 seconds Supplies power to motor, begins taking temperature, pressure, and IMU data Get position data via GPS and accelerometer; Sample 2 times per second Transmit data back to ground station, save on board to eeprom If position change by a certain margin, back up, turn motor, begin moving again Once distance traveled, deploy solar panels, end data collection Measure battery voltage Transmit data back to ground station, save on board to eeprom 62

63 REQUIREMENTS COMPLIANCE 63

64 Requirements Compliance Plan All requirements, both USLI and derived, will be complied with, and verified using the following methods The requirements may be found in the PDR Document Inspection Nondestructive/passive examination of the system No numerical data collected Design components present, use of checklists, follow safety guidelines Analysis Calculation of performance prior to any physical testing Completely theoretical based on expected performance Simulation software, FEA, hand calculations, CAD Demonstration System verification through repeatable exhibition of the design feature Pre-determined pass/fail criteria Parachute deployment, repeat flight tests, capability to launch within an hour Testing Demonstration of system with known input and output values Numerical data feedback as well as demonstrative verification Static motor fire, flight test with altimeters, recovery location tracking 64

65 NAR and FAA Compliance Test launches will only occur at NAR or TRA sponsored launch events. Only the mentor is allowed to handle rocket motors The rocket will use an L motors and will not exceed the impulse limit set by NASA 65

66 Launch Vehicle Verification Recovery Ejection Coupler Strength Motor Thrust/ Load Path Simulation/ Aerodynamics Overall System Performance Will the parachutes eject properly with the planned explosives? Will the rocket buckle at the coupler under max. thrust? Is the load path sufficiently strong/how does the motor behave when fired? Is the simulation accurate/is the rocket stable? Does the rocket reach the expected altitude/does every component work properly? Multiple ground ejections of each component Apply calculated moment to coupler Static fire of the motor, measuring thrust through the high-risk loadpath components Launch a subscale version of the rocket multiple times Launch the full-scale (final) rocket multiple times 66

67 General Requirements Compliance Most of the General Requirements are fulfilled through inspection of the schedule and design documents The TRA Mentor s (Jason Winningham) credentials have been confirmed Outreach will be demonstrated through the Outreach Reports The team will demonstrate the ability to teleconference during the review Rocket rail launch capability, reusability, and readiness will be demonstrated at the test flight 67

68 SAFETY 68

69 CRW Safety Commitment Training and Communications are key Weekly Safety Briefings on relevant current activities Create Hazard analysis and Standard operating procedures Team work and proper supervision are how risks and hazards can be minimized No team member shall work alone when manufacturing and testing the rocket and its components. CRW members double and triple check each other s work to ensure that all steps of manuals and standard procedures are followed Supervision from experienced mentors and staff ensures all procedures are done correctly. 69

70 ATF, DOT, and NPFA Compliance Rocket motors, e-match, igniters are purchased by the mentor or appropriate PRC staff with the proper license to ensure legality and compliance. Motors will be stored in Type 2 Magazine and transported in Type 3 magazines. 70

71 Safety Plan Hold weekly Safety Briefings with the entire CRW team Each sub-team will designate a Safety Representative to work with the Safety Officer Aid in Hazard and failure mode analysis for their respective sub-section of the rocket A Component Description Sheet will be created for each component used in the rocket Analyze failure modes Track evolution of the component to aid in verification process CRW has identified the required success criteria and a method of verification for each (as outlined in the PDR report) A Test Plan has been created based on the verification of all identified success criteria (as outlined in the PDR report) 71

72 Safety Representatives Bao H. Safety Officer Davis H. Launch Vehicle Lead Andrew W. Payload Lead The Safety Officer will be responsible for the overall safety outlined by the SLI Handbook The Launch Vehicle lead and the Payload lead will be responsible for the reliability and risk assessment of their systems. 72

73 Safety Briefings and Trainings Training Activity Date Red Cross First Aid CPR/AED/FA 10/13/2017 Basic Emergency Procedures 10/17/2017 Process Hazard Analysis 10/18/2017 Safe Testing Procedures 10/24/2017 Root-Cause Analysis 10/24/2017 Outreach Safety Procedures 11/7/2017 Sub-scale Launch Safety Procedures 11/14/2017 Hazardous Material Handling/Disposal 11/21/2017 Fire Extinguisher training 11/21/2017 TBD TBD The Red Team have completed training for First Aid and CPR/AED Additional training content will be added based on relevance to the stages in the development cycle. 73

74 Launch and Assembly Procedures The Test Plan and Verification Processes will be used to optimize the final design, assembly, and launch procedures Final rocket assembly procedures have been developed to fit the design concept Any changes to the design that require updating the assembly or launch procedures will be coordinated through the team safety officer Simulated runs of all procedures will take place at least one week prior to any launch 74

75 Published Information For the convenience of all team members, the following items will be located on the CRW team website: Material Safety Data Sheets Operators Manuals CRW Safety Regulations Safety Briefing slides Standard Operating Procedures The Safety Officer will work to keep this information relevant and up to date 75

76 PROGRAM MANAGEMENT 76

77 Work Breakdown Structure Payload Mechanical Structure Wheel Design Chassis Design Vehicle fabrication Electrical Design Component Selection Schematic Development Software Rover software Ground Station CRW Launch Vehicle Aft Motor Selection Fins Lower Body Tube Simulation Central Avionics Recovery Forward Upper Body Tube Nosecone Payload Fairing Management Website Updates Outreach Coordination Schedule and budget tracking Requirements Verification Interface management Safety Risk Identification and Analysis Mitigation Strategy Development Safety Briefing Manufacturing and Testing supervision 77

78 Schedule Schedule Philosophy Work around finals and Winter Break Internal deadlines 2 weeks ahead of NASA deadline for all documents Identify backup dates for critical test launches Upcoming Events Launch Opportunities: Nov 18, Dec 16, Jan 20, Feb 17 CDR Internal due date: Dec 22 78

79 Budget/Funding Summary Launch vehicle- two subscales (6 flights) and two full scales (6 flights) - $5,760 Payload- two fully operation rovers -$1,010 $750 margin for shipping/unexpected expenses Proposed to ASGC and UAH Propulsion Research Center for funding and Recovery 22% Motors 50% Rover Frame 4% Rover Electronics 11% Airframe 13% Rover Frame Rover Electronics Airframe Motors Recovery 79

80 Outreach Girls Science and Engineering Day Before project started, but good practice 80 middle school girls participated FIRST Robotics Boy Scout STEM Winter Camp Invited to teach space, robotics, and maker culture Science Olympiad at UAH, February

81 Web Presence Website updated and reformatted to highlight current content while preserving 2017 team documents Facebook and Instagram kept current Press release posted 81

82 Questions 82

83 Picture Credits NASA USLI Wikipedia Portrait_by_Curiosity_Rover_Arm_Camera.jpg NASA Thrustcurve.org Wonderfulengineering.com Professionalgrantwriter.org National Association of Rocketry Youtube Emotionalhealth.net 83

84 Component Mass (g) Number per rover Total Mass (g) Total Mass (lbm) Arduino Mega BMP SPG30 geared motor Appendix A: Electronics Mass Budget Motor Shield LSM9DS Solar cell Camera MHz Xbee Brushed DC motor Battery GPS Voltage regulator Total mass