FLIGHT READINESS REVIEW TEAM OPTICS
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1 FLIGHT READINESS REVIEW TEAM OPTICS
2 LAUNCH VEHICLE AND PAYLOAD DESIGN AND DIMENSIONS Vehicle Diameter 4 Upper Airframe Length 40 Lower Airframe Length 46 Coupler Band Length 1.5 Coupler Length 12 Nose Cone Length 16.5 Nose Cone Shoulder 4 Total Vehicle Length 105
3 KEY DESIGN FEATURE: FINS Set of 3 fins 3/32 G10 Fiberglass Root Cord: 14 Stability: 3.12 calibers
4 KEY DESIGN FEATURE: NOSE CONE Madcow Rocketry G12 Fiberglass 4:1 Ogive Composite tipped High compressive strength Easy to finish
5 KEY DESIGN FEATURE: AIRFRAME Madcow Rocketry G12 Fiberglass 4 diameter Heat resistant Easy to Finish Lower Airframe: 46 Upper Airframe: 40
6 KEY DESIGN FEATURES: BULKHEADS, THRUST PLATE, AND CENTERING RINGS G10 Fiberglass Thickness of centering rings: 1/8 Eye bolt attached to centering ring and upper bulkhead Thickness of bulkhead: 1/8 Thickness of thrust plate: 1/4 Thrust plate is be milled in classroom using the DaVinci CNC Mill Heat resistant Can withstand the force of ejection
7 KEY DESIGN FEATURE: COUPLER BAY G12 fiberglass coupler and coupler band Diameter: 3.9 Length: coupler band G10 end caps (2) 0.25 thick There is a lip for the inner diameter of the coupler.
8 KEY DESIGN FEATURES: DUAL DEPLOYMENT (DROGUE) Precision LP 18 Parachute Terminal Velocity: 75 ft/s Deployed at apogee with a black powder charge. The ignitor is long enough to reach behind the parachute, so the Drogue is pushed out of the lower body tube. Tethered with a D-link to an eyebolt on the centering ring and the coupler. 30 ft of ½ Tubular Kevlar recovery harness
9 KEY DESIGN FEATURES: DUAL DEPLOYMENT (MAIN) TAC-72 Ripstop Nylon Hemispherical with 6 shroud lines Deployed at 500 ft AGL with a black powder charge. The ignitor is long enough to reach behind the parachute, so the Main is pushed out of the upper body tube. Tethered with a D-link to an eyebolt on the Upper bulkhead and the coupler. 40 ft of ½ Tubular Kevlar recovery harness Terminal Velocity: 24 ft/s
10 KEY FEATURES: DUAL DEPLOYMENT (ELECTRONICS) Two PerfectFlite StratoLogger SL100 altimeters Two 9V batteries The altimeters and batteries are secured onto altimeter sled and into the coupler bay on threaded rods.
11 MOTOR DESCRIPTION K635-RL Cesaroni Total impulse: N Average Thrust: N Burn time: 3.1 s Diameter: 54 mm Length: 19.2 Total Mass: 1768 g Propellant Mass: 1115 g
12 STABILITY MARGIN Static stability margin: 3.12 calibers Stability on rail exit: 3.59 calibers CG location: 67.6 inches from the nose cone. CP location: 80.1 inches from the nose cone.
13 THRUST-TO-WEIGHT RATIO/RAIL EXIT VELOCITY Rail Exit Velocity Rail Exit Velocity: 61ft/s Thrust-To-Weight Ratio Thrust-To-Weight Ratio: 7.42
14 MASS STATEMENTS Vehicle Component Mass (g) Upper Airframe Coupler Lower Airframe (loaded) Total (loaded) 8,774.5 Total (after burn)
15 PARACHUTES Drogue Parachute: 18 Descent Rate: 75 ft/s Main Parachute: 72 Descent Rate: 24 ft/s
16 KINETIC ENERGY AT KEY PHASES Kinetic Energy at Drogue Deployment Section Kinetic Energy (Ft-lbs) Upper Section Coupler Lower Section Kinetic Energy at Landing Section Kinetic Energy (Ft-lbs) Upper Section Coupler 23 Lower Section 67.36
17 PREDICTED ALTITUDE WITH WIND Windspeed (mph) Altitude (feet)
18 PREDICTED DRIFT CALCULATIONS Wind Speed (mph) Max Drift (ft) 0 mph 0 ft 5 mph ft 10 mph ft 15 mph ft 20 mph ft
19 TEST PLANS AND PROCEDURES TEST PLANS AND PROCEDURES Test Objective Success Criteria Methodology Rocksim Flight Simulations Achieve a stable simulated flight, apogee of approx ft, drift within 2500 feet, and landing velocity within KE requirements. Stable ascent, apogee within 15% of 5280 feet, drift distance under 2500 feet, landing velocity under 25 ft/s. Set launch conditions to those in Huntsville. Poor simulated apogee may be solved with a motor change. Poor stability is resolved with fin redesign or ballast. Drift distance or descent velocity resolved with parachute changes.
20 TEST PLANS AND PROCEDURES TEST PLANS AND PROCEDURES Test Objective Success Criteria Methodology Fin Flutter Speed Calculation Test the rigidity of the fin geometry against the forces of flight. A flutter velocity that is greater than the maximum velocity reached by the launch vehicle. Velocity is calculated using the equation in the Apogee Rockets newsletter Issue 29. If a flutter velocity is reached that is less than the maximum velocity during flight, fin geometry will be changed.
21 TEST PLANS AND PROCEDURES TEST PLANS AND PROCEDURES Ejection Ground Test Test Objective Success Criteria Methodology Determine the black powder charge amount required to separate the upper and lower body tubes. Breaking of shear pins and complete separation of both tubes. An online calculator is used to find a preliminary ejection charge amount. The vehicle is packed ready for flight with ejection charges located behind the parachutes to push them out; charges are activated with launch stand from safe distance. If complete separation is not achieved, the test is repeated with a larger charge.
22 TEST PLANS AND PROCEDURES TEST PLANS AND PROCEDURES Test Objective Success Criteria Methodology Subscale Test Launch Validate the launch vehicle design features. Stable flight, apogee within 25% margin of simulations, and successful dual deployment recovery system functionality. The subscale uses the dualdeployment recovery system, altimeter arming system, and construction methods of the full-scale. An unstable flight will result in fin redesign or use of ballast to alter stability margin. Failure of any subsystem will result in a re-flight after resolving the issue.
23 TEST PLANS AND PROCEDURES TEST PLANS AND PROCEDURES Test Objective Success Criteria Methodology Fin and Bulkhead Stress Tests Validate construction integrity. Fins and bulkheads resist applied stresses. Apply stress to each fin and tug on the recovery harness at each point of attachment. If fins or bulkhead attachment points give in to stress, they will be removed and reattached.
24 TEST PLANS AND PROCEDURES TEST PLANS AND PROCEDURES Test Objective Success Criteria Methodology Payload Bench Test Confirm payload functionality. The payload gathers and logs atmospheric data from the photometer. The LED photometer will be assembled with an Arduino, accelerometer, and SD logger and tested outside with the correct program to deduce atmospheric data. If the bench test is unsuccessful the payload design will be redesigned and re-implemented.
25 TEST PLANS AND PROCEDURES TEST PLANS AND PROCEDURES Test Objective Success Criteria Methodology Full Scale Test Launch Validate the launch vehicle design features. Stable flight, apogee within 25% margin of simulations, and successful dual deployment recovery system functionality. The vehicle uses the dualdeployment recovery system, altimeter arming system, and construction methods of the full-scale. Failure of any subsystem will result in a re-flight after resolving the issue.
26 FULL SCALE TEST FLIGHT Rocksim predicted apogee: 1622 feet Achieved apogee: 1603 ft Calculated drag coefficient: 0.7 Stable, nominal flight
27 RECOVERY SYSTEM TESTS The dual deployment system proved successful in the subscale and full scale test flight. Ejection tests were performed prior to flight to test ejection charge size
28 SUMMARY OF VEHICLE REQUIREMENTS VERIFICATION The team will meet all of the requirements in the Statement of Work (SOW) by following the plans outlined in the CDR report. A GANTT Chart is followed to ensure appropriate completion of requirements. Specific procedures can be found in the tests and procedures section
29 PAYLOAD DESIGN AND DIMENSIONS Payload clip: 0.5 by 2 LEDs are soldered onto copper plate circuit board Electronics sled: 6 long on 4 diameter plates Electronics screwed into sled
30 PAYLOAD INTEGRATION Copper LED payload clips are attached to airframe through vertical slits with a screw on either side. Electronics sled fits into payload bay above upper bulkhead and below nose cone which is held in place by three nose screws Output wires from LEDs are routed internally to the electronics sled
31 INTERFACES WITH GROUND SYSTEMS The launch vehicle motor is ignited using a 12 volt DC electrical launch system that returns to the off position when the input is released. Altimeter screw switches on the altimeter sled are armed by hand through the coupler.
32 SUMMARY OF PAYLOAD REQUIREMENTS VERIFICATION The vehicle will carry an atmospheric data monitoring system composed of LED clips analyzing different wavelengths of light absorbed. Data will be processed post-flight to calculate water vapor, optical thickness, and photosynthetically active radiation. The payload does not jettison or contain UAVs, and is fully modular and reusable.
33 Q & A
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