Critical Design Review University of Alabama in Huntsville - - PowerPoint PPT Presentation

University of Alabama in Huntsville

NASA SL Critical Design Review

University of Alabama in Huntsville USLI CDR 1 1/16/2018

LAUNCH VEHICLE

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University of Alabama in Huntsville USLI CDR 3 1/16/2018

Vehicle Summary ry

• 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

Vehicle CONOPS

11/3/2017 University of Alabama in Huntsville USLI PDR 4

Powered Ascent: 0 – 3.3 seconds 0 – 1,190 ft. Deploy Drogue: 19 seconds 5,429 ft. Deploy Main: 62 seconds 600 ft. Landing: 121 seconds 0 ft. Deploy Rover: Team Command

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Vehicle Summary ry

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

• Upper Airframe houses the rover, piston

ejection system, and GPS tracker

Upper Air irframe Overview

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• 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

Nose Cone

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• Houses payload and deployment device
• Fiberglass; 6.17 in. OD, 6 in. ID
• Shear pinned to nose cone; 10-32 bolt

connection to transition

Fairing

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Transition

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• 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 10-32 bolts

Forward Insert Aft

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Transition

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• Three piece design allows for a 57% reduction in

weight

• Max stress on aluminum bulkhead: 0.712 ksi
• Yield stress: 42 ksi

1/16/2018

• 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

Transition Coupler

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• CO2 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 CO2

housing

Piston Overview

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• Ejects rover and nose cone
• Fiberglass coupler with aluminum bulkhead

Piston Head

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• Houses CO2 cartridge and release mechanism
• 3D printed ABS Plastic
• Allows for easy and quick modification upon

testing results

CO CO2 Housing

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• CO2 housing positioned in transition shoulder
• Mounted to side using 3-D printed brackets

and 4-40 bolts

• Keeps housing fixed during flight

Piston Configuration

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• Aft Subsystem Components

Aft ft Subsystem Overview

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Recovery Bulkhead & U-Bolt Fin(s) Fin Can Thrust Plate Motor Retention Ring Motor/Motor Case

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• Trapezoidal Fin Set (4)

– Maintain stability

• G10 Fiberglass

– Great strength/weight ratio – 3/16” thickness

• Flutter Speed

– Calculated to be Mach 1.947 (2191.57 ft./sec)

Fins

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• 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

Fin In Interface

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Fin in Can Assembly Overview

• Consists of: Fin Can,

Motor Retention Ring, Thrust Plate, and Rail Button Press fit nut

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Exploded Vie iew of f the Fin in Can Assembly

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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

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Fin Can Dimensions

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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

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Thrust Pla late Dim imensions

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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

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Motor Retention Rin ing Dim imensions

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• Aerotech L1420R-PS

– Best met altitude target

• Avg. Thrust: 326.18 lbf.
• Burn Time: 3.2 sec

Motor Selection

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Motor Altitude Aerotech L2200 6107 ft. Aerotech L1420 5429 ft. Aerotech L1520 4329 ft.

1/16/2018 University of Alabama in Huntsville USLI CDR

• Dimensions

– Total length – 103 in. – Wet mass – 43.80 lbm. – CP location – 68.93 in. – CG location – 55.60 in.

OpenRocket Fli light Sim imulation

28 55.60 inches 68.93 inches 1/16/2018 University of Alabama in Huntsville USLI CDR

Stability Margin

29 Motor Burnout (3.28 cal.) Initial Stability (2.22 cal.) Apogee 1/16/2018 University of Alabama in Huntsville USLI CDR

OpenRocket Fli light Sim imulation

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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 8.7 CG 56 in. CP 69 in.

OpenRocket Fli light Sim imulation

31 Motor Burnout (3.27 sec.) Apogee (18.62 sec.) Main Deploy (62.39 sec.) 1/16/2018 University of Alabama in Huntsville USLI CDR

• 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

▪ 𝐷𝑒 = 0.56

– 10,000 flights per simulation

Full ll Scale Monte Carl rlo Sim imulation

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Full ll Scale Monte Carl rlo Sim imulation

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Mean: 5626.31 feet Median: 5617.45 feet

• Std. Deviation: 192.29 feet

Max Altitude: 6463.91 feet

• Min. Altitude: 5010.83 feet

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• Central Subsystem responsibilities:

– Coupler between airframes – Flight Avionics – Ejection System – Tracking and Ground Station – Recovery System

Central Subsystem Overview

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Drift Analysis

Vwind

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• 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

Vrelative Wind Speed (mph) 5 10 15 20 OpenRocket Drift Distance (ft) 17.6 465.8 946.7 1461.9 1995.7 CRW Model Drift Distance (ft) 573.19 1148.9 1741.9 2311.8

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• Requirement: No individual section will have a kinetic

energy greater than 75 ft.-lbf. upon landing

• Terminal velocity under drogue: 112.7 ft./sec.
• Terminal velocity under main: 17.45 ft./sec.

Recovery ry System Calc lculations

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Vehicle Section Mass (lbm.) KE (ft.-lbf.) Fairing 14.35 67.85 Coupler 11.15 52.72 Aft 9.89 46.76

1/16/2018 University of Alabama in Huntsville USLI CDR

• 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: 112.7 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: 17.45 ft/s

Recovery ry System

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Recovery Avio ionics 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

ll re redundancy in in avionics and ig igniti tion

Avionics

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Coupler

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

Recovery ry Deployment Avio ionics

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• 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

1/16/2018 University of Alabama in Huntsville USLI CDR

GPS Tracking & Rover Deployment Subsystem

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System

• 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

Str Structure In 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

1/16/2018 University of Alabama in Huntsville USLI CDR

Subscale Design

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Scaling Factors:

• Geometry of the design
• Average Thrust of Motor and Thrust Curve
• Kinetic Energy

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• Successful recovery of all three subscale flights
• Altimeters ignited the black powder charges at

the correct altitudes

Subscale Flig light Results

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

Subscale Flig light Results

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• Using data gathered from

the altimeters, the drag force and coefficient for the vehicle were found

• Using a weight of 6.33 lbs,

an acceleration of 60.128 ft/s, an A of 0.0533 ft2, a 𝜍

• f 0.0751 lb/ft3, and a

velocity of 396.55 ft/s: – Cd = 0.56

Subscale Drag Coefficient

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▪ A = Area of the exposed section, ft2 ▪ 𝜍 = density of the air, lbm/ft3 ▪ Cd = Coefficient of Drag ▪ u = Velocity, ft/s ▪ m = mass, lbm ▪ A = acceleration of the vehicle, ft/s2 ▪ g = acceleration of gravity, ft/s2

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

Payload

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• Rover fits inside the piston, which ejects it from the fairing
• CO2 cannister pushes rover through nose cone

Payload In Integration

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

Payload CONOPS

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1 3 2

• 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

Chassis

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• 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.

Wheel

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• 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

Stabilizing Arm

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• 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

Strength Check Notes

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Part Load Case Safety Margin Chassis 210 lbf (sidewall) +2.45 Chassis 210 lbf (base) +1.37 Wheel Hub 210 lbf (sidewall) +0.73 Wheel Hinge 105 lbf (each) +5.11 Spoke 210 lbf (lengthwise force) +11.98 Spoke 7 lbf (Drive force) +6.42

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

Electronics Tray

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

Electronics Tray

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• 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 0.375 in. when closed
• 12 in. x 7.25 in. x 0.5 in. when open

Rover Lid id wit ith Mechanism

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• The mass of all components totaled 6.6 lbm.
• A 6% margin was added to the total weight to

account for fasteners and adhesives

Rover Mass Budget

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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 7.0

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Payload Power Budget

57 1/16/2018 University of Alabama in Huntsville USLI CDR Part t Name me Current nt (mA) Voltage e (V) Adj Current nt Duty Cycle e (%) (%) Time me (hr) Tota tal (mWh Wh) Arduino Mega 0.17 5 0.0787 100 2.5 0.984 Camera 350 5 162.037 10 2.5 202.546 GPS 53 3.3 16.194 20 2.5 26.721 IMU 0.35 3 0.0972 100 2.5 0.729 Pressure/Temp 0.36 3.3 0.11 17 2.5 0.154 Wheel Motors 650 12 722.222 20 2.5 4333.333 Lid Motors 360 5 166.667 5 2.5 104.167 Radio transmit 229 3.3 69.972 10 2.5 57.727 Radio idle 44 3.3 13.444 90 2.5 99.825 Datalogger 100 3 27.778 10 2.5 20.833 Power required 4847.01 Part t Name me Current nt (mA) Voltage e (V) Adj Current nt Duty Cycle e (%) (%) Time me (hr) Tota tal (mWh Wh) Li-Ion Battery 2600 10.8 N/A 100 1 28080 Power Supplied 28080 Power er Supplied ed 28080 mHr Power er requi uired ed 28080 mHr Factor tor of Safety ety 5.793 mHr

Ele lectronics Blo lock Dia iagram

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Ele lectronics Fail ilure Path

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• Emphasizes dependence of each lower level component
• n the component above it

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SAFETY

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• 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

Safety Overview

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• 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

Communication

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• 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

Standard Operating Procedures

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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

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• Factors affecting launch vehicle and payload

– Sudden high winds – Humidity – Extreme temperatures – Terrain

• Mitigations established:

– Minimum exposure to environment – Constant monitor of the weather

Environmental Factors

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• 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

Environmental Factors

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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

Upcoming Trainings

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Safety Briefings are held based to relevant safety topics.

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• 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.

Test Plan

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Test Plan

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Test Number Test Type Test Status T1 Subscale Ejection Charge Test ➢ Test has been conducted prior to the subscale flight on 11-19- 2017 ➢ Test shows that rocket has to go drogue-less and use only

• ne 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 10-11 T4 Fin Can Load Test ➢ Test will be planned for the end

• f January to the early February

before the full scale launch.

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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

• f February once test articles

arrive T7 Faring Drop Test ➢ Test is planned for middle to end

• f 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

• ut 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

Test Plan

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TESTING AND REQUIREMENTS VERIFICATION

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• 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

Verification Reports

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• 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

General Requirements Verification

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Test Number Test Type Description Test Status T1 GPS tracker range and capability

• The GPS tracker of the launch vehicle and the

payload will be tested inside of their respective fairing/compartment. This is to ensure that the GPS can reliable transmit and receive signals.

• The test will also be conducted in obstacles such as

trees and buildings to reveal the limits of the GPS.

• The full test of GPS system performance and

reliability will be the subscale launch

• Single component tests (radio, GPS receiver), can

be done by a team member without supervision.

• Subscale launch tests will adhere to SOP.

➢ Tracker currently Exhibit poor performances. ➢ Team is currently learning how to trouble shoot issues with tracker. T2 Electrical Charge

• n E-matches
• 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.

Test Plan

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T3 Altimeter Test

• The functionality of the altimeter will be evaluated with

the Charger Rocket Works’ altimeter testing container.

• Only applied for in-house made altimeters. Third party

altimeters like Statologger will not require testing. ➢ Test will be scheduled when altimeter has been created. T4 Ejection Charge Test

• This is to experimentally verify the correct

amount of black powder to be used in the ejection of the drogue and main parachutes.

• 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 11- 19-2017 ➢ 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

• Experimentally verify the functionality of the

rover deployment mechanism.

• 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

• Experimentally verify the compression strength of the

fairing transition

• Only the section in doubt from the FEA results shall

printed for test.

• Currently planned to be a destructive test

➢ Test will be conducted

• nce

FEA results shows doubts in the structures. T7 Rover Terrain Test

• The Rover, once constructed, shall be put through its

paces in different terrain conditions (except water and mud).

• Test is to verify the spoke wheel design.

➢ Test will be planned and carried out when rover is constructed.

Test Plan

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Full Scale Budget

Budget Summary Airframe $ 1763.11 Electronics $ 334.89 Recovery $ 899.09 Motors $ 1589.96 Rover Structure $ 438.97 Rover Electronics $ 682.34 Total Cost $ 5708.36

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On the Pad Budget

Launch Vehicle Airframe $ 997.81 Electronics $ 167.45 Recovery $ 621.09 Motor $ 259.99 Rover $ 621.00 Total $ 2046.34

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11/3/2017 University of Alabama in Huntsville USLI PDR 78