NASA Student Launch Competition CORNELL ROCKETRY TEAM. Preliminary Design Review Presentation Maxi-Mav Participant

Transcription

1 CORNELL ROCKETRY TEAM Presentation Maxi-Mav Participant NASA Student Launch Competition

2 LAUNCH VEHICLE

3 LAUNCH VEHICLE SUMMARY Length: 72 in (6 ft) Weight: 182 oz (11.8 lbm) Motor: Cesaroni Technologies J380 Recovery: Drogue Parachute (20 in diameter) Deployed at apogee PEM Parachute (30 in diameter) Deployed at 1000 ft. Main Parachute (42 in diameter) Deployed at 500 ft. Tracking: Main AV Bay above Avionics Tube Commercial Altimeters: Secondary AV Bay above Booster Section

4 VEHICLE DESIGN D = 4 C G = 41.3 C P = Nosecone PEM PEM Avionics Avionics Tube Booster Section Rivets Shear Pins Main AV PEM Parachute Second AV Drogue J380 Motor Main

5 MASS STATEMENT Containing Total Mass Component (oz) Nosecone 17.8 PEM Tube 46 PEM Avionics 14.4 Avionics Tube 29.5 Booster Tube 74.8 Launch Vehicle 182 Current Thrust to Weight Ratio: 7.5:1 Maximum Mass Growth = 50% (17.0lbs) Thrust to Weight Ratio 5:1

6 RECOVERY Section (Independent) Initial Weight (oz) Final Empty Weight (oz) PEM PEM Tube Avionics Tube Booster Tube Total Section Parachute Parachute Diameter (in) Deployment Altitude (ft) Rocket Body Drogue 20 apogee PEM PEM Rocket Body Main

7 RECOVERY Both the PEM and Booster avionics bays include two commercial altimeters: a Marsa54L and Perfectflite Statologger. The Booster-avionics bay will only contain commercial electronics and be shielded from all other devices. The Marsa54L in this section will report altitude by a series of beeps. Blast charges are currently sized at 1.5 and 2 g, but ground testing will be performed to verify the amount of energetics. Terminal Blocks Charge Canisters 4 Main Screw switches Drogue 9

8 KINETIC ENERGY DRIFT Component Mass (oz) Velocity (ft/s) Kinetic Energy (lbf-ft) PEM Assembly Tethered Rocket Body Component Windspeed (mph) Drift (ft) PEM Assembly Tethered Rocket Body

9 RISK DEFINITION Level of Risk High Moderate Low Minimal Level of Management Highly Undesirable Undesirable Acceptable Acceptable Severity Probability Catastrophic Critical Marginal Negligible A - Frequent 1A 2A 3A 4A B Probable 1B 2B 3B 4B C - Occasional 1C 2C 3C 4C D Remote 1D 2D 3D 4D E - Improbable 1E 2E 3E 4E

10 RISK DEFINITION Severity Definitions A condition that can cause: Description Personnel Facility/Equipment Range Safety Project Plan Environmental Safety and Health 1 Catastrophic Loss of life or a permanent disabling injury. 2 Critical Severe injury or occupational related illness. 3 Marginal Minor injury or occupational related illness. 4 Negligible First aid injury or occupational related illness. Loss of facility, systems or associated hardware that result in being unable to complete all mission objectives. Major damage to facilities, systems, or equipment that result in partial mission failure. Minor damage to facilities, systems or equipment that will not compromise mission objectives. Minimal damage to facility, systems, or equipment. Operations not permitted by the RSO and NFPA 1127 prior to launch. Mission unable to proceed. Operations not permitted by the RSO and NFPA 1127 occur during launch. Mission suspended or laws and regulations are violated. Operations are permitted by the RSO and NFPA 1127, but hazards unrelated to flight hardware design occur during launch. Operations are permitted and executed in accordance with the RSO and NFPA Delay of mission critical components or budget overruns that result in project termination. Delay of mission critical components or budget overruns that compromise mission scope. Minor delays of noncritical components or budget increase. Minor delays of noncritical components. Irreversible severe environmental damage that violates law and regulation. Reversible environmental damage causing a violation of law or regulation. Mitigatible environmental damage without violation of law or regulations where restoration activities can be accomplished. Minimal environmental damage not violation law or regulation.

11 RISK DEFINITION Description Qualitative Definition Quantitative Definition A Frequent High likelihood to occur immediately or Probability is > 0.1 expected to be continuously experienced. B Probable Likely to occur or expected to occur 0.1 probability > 0.01 frequently within time. C Occasional Expected to occur several times or 0.01 probability >0.001 occasionally within time. D Remote Unlikely to occur, but can be reasonably probability > expected to occur at some point within time. E Improbable Very unlikely to occur and an occurrence is not expected to be experienced within time probability

12 Hazard Cause Effect Motor Cato (catastrophic failure) Parachute deployment failure Altimeter failure Improper motor manufacturing Altimeter failure Electronics failure Snag in parachutes Parachutes snag on shock cord Failure in electronics Failure in programming Sections fail to separate Black powder charges fail Altimeters fail Motor retention failure Design of retention failure Assembly of retention failure Motor is too powerful for designed retention Launch vehicle is destroyed and motor has failed. Moderate explosion Launch vehicle will not lose enough speed as it descends and will strike the ground at a high velocity Parachutes will fail to deploy (see above) Sections will fail to separate (see below) No data collection Parachutes will not deploy (see above) Motor falls out of booster section while propelling body forward and launch vehicle fails to achieve 3,000 ft altitude Mitigation Careful measurement of shroud lines and shock cord for appropriate lengths and safety factor Methodical packaging of parachutes No excessive altimeter/electronics use on days of flight Checking and testing of altimeter programming before days of flight. Calculate the correct amount of black powder needed for each blast charge. Measure black powder using scale Retention rings will be machined using designs from Solidworks to assure proper dimensions. Robust material such as aluminum will be used to assure the integrity of the design Verification Test the altimeters the night before days of flight for programming failures. Test flights to verify appropriate lengths of shrouds and shock cord. Launch field verification of altitudes in addition to previous verifications of altitude on and before flight days Ground testing to be done on the full-scale launch vehicle before all test flights. All test flights will serve as verification The launch vehicle will be launched multiple times with this design implemented. This design will also be simulated using software such as ANSYS. Pre- RAC 1D 3D 3D 3D 3E Post- RAC 4D 4D 4D 4E

13 Hazard Cause Effect Pre-RAC Mitigation Verification Post- RAC PEM parachute gets tangled with other section Launch vehicle is too heavy for AGSE Shock cord gets tangled due to being too long. Parachute is not deployed correctly PEM is not fully deployed Launch vehicle body is too heavy for previously sized parachutes Inadequate weight The launch vehicle will measured or tolerated not be able to be by the AGSE team launched. No factors of safety for launch vehicle weight Unaccounted for growth Construction Failure to follow errors result in directions and to be failure(bad precise epoxying, careless measurements, etc) Not enough time Failure to create a for adequate testing precise timeline The launch vehicle is not durable in launch, deployment and landing Imprecision in the launch vehicle design and less verification of design 3C 3D 4E 4E Proper calculations and measurement for sizing Use a deployment bag Keep AGSE Team informed to make sure they incorporate a factor of safety in design Keep launch vehicle within projected weight margin Plenty of test launches and ground testing. Weighing each component and testing on the launch pad to gauge its suitability for launch. Educate those doing epoxy work about Launches and drop tests proper use, safety and application and ground tests. Supervision by those more experienced using these methods Supervision by those who have experience with composites Create a rigorous timeline and ensure Comparison of progress everyone stays on schedule with timeline schedule. Make due dates at least three days in advance for deliverables 3E 4D 4D 4E

14 Hazard Cause Effect Pre-RAC Mitigation Verification Post- RAC Launch vehicle Explosion/uncontai ned debris Unrecovered Launch vehicle Launch vehicle Explosion/uncontai ned debris Motor malfunction Tube shatters Bad construction Tracking system failure Unpredictable flight path Motor malfunction Tube shatters Bad construction Potential harm to bystanders environment. Fuel/fiberglass may be spread over launch area Harm to environment, launch vehicle parts will be left in the field where wildlife may interact with it Potential harm to bystanders environment. Fuel/fiberglass may be spread over launch area 1D 2C 1D Multiple subscale and full scale launches No unwanted debris should expose potential failures before may be seen on test launch day. Ground testing will mitigate many launches. construction errors. Launch vehicle will be designed with a structural factor of safety because fiberglass will be used Multiple subscale and full scale launches will be used to make sure tracking system is functioning and flight path is predictable. OpenRocket design will simulate the stability of design. Launch vehicle must be recovered every launch. Multiple subscale and full scale launches No unwanted debris should expose potential failures before may be seen on test launch day. Ground testing will mitigate many launches. construction errors. Launch vehicle will be designed with a structural factor of safety because fiberglass will be used 2D

15 PAYLOAD EJECTION MECHANISM (PEM)

16 DESIGN RATIONALE The PEM shall accept and seal the payload. The body of the PEM will be the forward section of the rocket. This entire section will be ejected at 1000 feet during decent. The hinge will be a torsional spring that will want to close the system shut. The system will remain open through the use of a folding bracket. This will keep the system open even against a gust of wind. The payload will be received by a net system, which will pull down on the pivoting section of the folding bracket. Once the bracket has folded the torsional spring will force the door closed.

17 DESIGN RATIONALE CONT. Magnets in the hatch and body will pull the door closed, and a latch mechanism will seal the payload into place Sealing the door will cause copper conductive strips to complete a circuit, which will send a signal to the gas system to fill airbags which will secure the payload in place.

18 LATCH DESIGN 6 6 Number Part Name 1 Bending Portion 2 Locking Portion 3 Epoxy Portion 4 Magnets 5 Inner Locking Tube 6 Door Number Component Name 1 Door 2 Conductive Copper Tape 3 Inner Locking Tube R 800 kohm C 10 microf Capacitor V 12V battery S 12V Solenoid Valve MOSFET MOSFET Semiconductor

19 PERFORMANCE CHARACTERISTICS AND METRICS OF LATCH AND CIRCUIT DESIGN A simple latch mechanism was chosen to close the PEM door due to its simplicity and reliability proven by the many uses in everyday applications. Magnets are used to ensure that the latch is pulled shut. Once shut, a circuit will be activated that opens a solenoid valve and inflates airbags, securing the payload in the rocket. This circuit is programmed to produce a 2 second delay to reduce the risk that the airbags inflate before the door latch activates. Characteristic Door Closes The Circuit functions with the closing door The solenoid valve opens with a delay The airbags are fully inflated Door is able to be opened Metric The door of the PEM closes securely and will not open freely from simulated launch conditions The closing of the door will be repeated multiple times, and the circuit will be checked to make sure that the correct voltage and current is supplied to the solenoid each time. The valve is observed to be in an open position when attached to a power source after a short delay The circuit stays connected and continues to keep the solenoid valve open, filling up the airbags to full capacity The PEM Door can be easily opened after simulated launch conditions

20 MANUFACTURE AND INTEGRATION OF LATCH AND CIRCUIT DESIGN The latch will be 3D printed. The latch will be secured to the PEM door using epoxy on the Epoxy Portion. The Epoxy Portion will be 1.2 cm long and curved to fit the 1.5 radius rocket door. There will be a part of the latch that is flexible, dubbed the bending portion. This will be 0.2 cm thick in order to allow for the flexibility while still maintaining structural integrity from the 3D printing. The locking portion of the latch will attach to an inner locking tube that will be fit inside the rocket body snuggly to provide end support for the door when it closes and a location for the latch to lock on to. The door and locking tube will have a cut out section to make room for the latch without taking up too much room of the airframe. The ends of the door will fit snuggly against the rest of the airframe, therefore the incision on the door will be enough to make room for the latch and a bit more to provide easier access to open the PEM. The inner locking tube will stick out about 1 cm from the rest of the airframe body. The inner locking tube will also have bored out holes for the two magnets to fit in. The magnets will be located on the inside portion of the door, attached with epoxy, and will fit in the holes bored out on the inner locking tube when the door is closed. Below the bored out holes will be a thin membrane of balsa wood that will have another opposite pair of attracting magnets epoxied to them. The battery will send current through the resistor and capacitor, charging the capacitor to full voltage. After the voltage of the capacitor passes the Voltage threshold of the MOSFET semiconductor, the gate will open, allowing the battery to power the solenoid valve. The time for the gate to open should be about 5 seconds or so. The components will be mounted on a perforated prototyping board, and this board will be cut down to size in order to save space. The components will be soldered together on this board.

21 AIRBAG SYSTEM Once the payload is inserted, the airbag system will hold it in a central position in the rocket so that it does greatly change the center of mass of the rocket. There will be a two-way, normally-closed, solenoid valve which will be switched on by the door closing circuit, allowing gas from a CO 2 cartridge to flow. Between the solenoid valve and the CO 2 cartridge will be a pressure regulator so that the pressure being fed into the airbags can be kept below a certain level in order to maintain a 4:1 factor of safety for all components. The gas will feed through a manifold, and into four separate hoses, which will fill four airbags surrounding the payload. The relief valve after the solenoid valve will also serve to keep the pressure low enough to maintain a 4:1 factor of safety. The CO 2 cartridge, solenoid valve, and manifold will be positioned in the nosecone of the rocket, and the airbags themselves will surround the walls of the Payload Ejection Mechanism.

22 AIRBAG SYSTEM Characteristic Airbags are sufficiently filled Solenoid valve opens Solenoid valve opens quickly The airbags inflate quickly Airbags do not leak or burst CO 2 cartridge opens fully Parts arrive on time All parts needed are obtained Hoses do not leak or burst Relief valve controls pressure Pressure regulator controls pressure reaching the system at any given time Metric Payload does not move more than 0.15 inches in any direction once airbags have been fully inflated The valve is observed to be in an open position when attached to a power source There is no more than a 5-second lag between when the solenoid valve is attached to a power source, and when it opens There is no more than a 20-second lag between when the solenoid valve is attached and when the airbags begin to inflate When left for thirty minutes at the pressure they will be at during performance, the pressure inside does not deviate by more than 10 PSI When attached to the solenoid valve, and opening device in between, gas finishes flowing out of solenoid valve after no longer than 15 seconds Construction of the system can begin by December The full system can be constructed without any missing parts When left for thirty minutes at the pressure they will be at during performance, the pressure inside the airbags does not deviate by more than 10 PSI If the pressure of the system is raised above its set cracking pressure, the relief valve should release gas until the pressure is at or below the set cracking pressure No more than 100 PSI ever reaches the solenoid valve

23 AIRBAG SYSTEM The system will be made up of a CO 2 cartridge, a pressure regulator, a solenoid valve, a relief valve, a manifold, hoses, and airbags. The cartridge, pressure regulator, solenoid valve, relief valve, manifold, and hoses will be purchased separately, and connected, likely using threading adaptors. The airbags will be made out of bicycle tubes. The tubes will be cut into four strips 5.5 in length. One end of each tube will be closed, and the other will attach to a hose. The cartridge, pressure regulator, solenoid valve, relief valve, manifold, and hoses will be mounted inside the nosecone of the rocket, the hoses leading out, feeding the airbags. The solenoid valve will be connected to a circuit that will complete when the door to the Payload Ejection Mechanism closes, and will open at this point CO 2 cartridge 2 Pressure regulator 3 Solenoid valve 4 Relief valve 5 Manifold 6 Airbags 6

24 AMOUNT OF CO 2 DESIRED The volume of the air bag was determined as a fraction of the volume of the entire bicycle tube, a torus, it was a section of. The diameter of the tube of the rocket will be padded such that its inner diameter is 3. The diameter of the payload is Therefore, the space between the wall of the PEM and the payload in any direction is 3.00" 0.75" = 1.125" 2" and so the desired width of the airbags is about The volume of a torus is 2π 2 ab 2, where a is the distance from the circular cross-section to the center of the torus, and b is the radius of the circular cross-section. In this case, the circumference of the torus is 27. Therefore, a is = " 2π 2 because this number represents the radius from the center of the torus to the inner diameter, plus half of the diameter of the tube. The volume of the torus is therefore 2π = in 3 The payload compartment is 6 long, and the payload is 4.75 long. To safely secure the payload, the airbag will be 5.5 long. Treating the segment of torus as an arc of a circle (looking at the shape two-dimensionally), the angle of the arc height 5.5 is θ = 2 arcsin 5.5/ = rad

25 AMOUNT OF CO 2 DESIRED, CONT. Therefore, the volume of one airbag is So the volume of all four airbags is π = L = in 3 4X0.0952= L Choosing a desired pressure of 4 ATM in each airbag, the approximate number of moles of CO 2 needed is calculated (assuming room temperature): PV = nrt = n(0.0822)(293) The molar mass of CO 2 is So in mol, the mass of CO 2 is n = mol = g = g So it appears that a 12 g CO 2 cartridge is approximately the right size. These numbers are approximations. In reality, the pressure inside the airbags will be lower than 4 atm because the pressure is shared with the other components of the system, not just confined to the airbags.

26 PEM CAD MODEL

27 Hazard Cause Effect Pre RAC Mitigation Verification Post RAC Solenoid valve does not open Insufficient power, or other malfunction that cannot be anticipated Airbags do not inflate, and payload is not supported 2C Find the correct solenoid valve for the given power source Testing should show that the solenoid valve always opens under the expected conditions 2E Airbags leak Insufficient strength, or excessive pressure Airbags deflate, and payload is not supported 2C Seal airbags properly, use proper hoses and connections. Do not exceed pressure rating for tubes Testing the system should show that all parts are strong enough for their roles, and that the relief valve is functioning properly, not letting the pressure exceed expected pressures 2E Airbags burst Insufficient strength, or excessive pressure There is no pressure to hold the payload in place 2D Use the proper amount of CO2, and use a 3-way solenoid valve for relief Testing the system should show that all parts are strong enough for their roles, and that the relief valve is functioning properly, not letting the pressure exceed expected pressures 2E CO 2 cartridge does not open Malfunction/flaw of opening device The airbags do not inflate, and the payload is not supported 2A Use an attachment that has a needle inside to pierce the metal film sealing the CO 2 canister and allow the gas to flow Testing should show that the CO 2 cartridge always opens (gas flows into airbags) under the expected conditions 2E

28 Hazard Cause Effect Pre RAC Mitigation Verification Post RAC Parts take a long time to order/arrive Long shipping time Testing and implementation are delayed, and working speed has to be increased 4B Working quickly and efficiently to select and order parts ahead of time The mitigation will be deemed successful if all necessary parts have been ordered, on time, and their shipping times are not too long. 4C Parts break before the project is complete and need to be replaced Parts are not strong enough for their assigned tasks Having to order new parts costs extra money and delays completion of the project 3C Because several parts may be damaged or used up during testing phase, this should be accounted for in ordering extra parts such as tubes and CO 2 cartridges Budgets should reflect the possibility of replacing parts 4C Hoses leak/burst Insufficient strength/excessiv e pressure There is not sufficient pressure in the airbags to hold the payload in place 2C Find hoses that are properly threaded and fit with one another, and have a high enough pressure rating for their task. Relief valve relieves excess pressure if it occurs Testing the system should show that all parts are strong enough for their roles, and that the relief valve is functioning properly, not letting the pressure exceed expected pressures 2E

29 SPACE SYSTEM

30 SPACE CLAW DESIGN Accurately captures payload from ground Transports payload onto SPACE ramp for insertion into launch vehicle -Plastic construction decreases mass without sacrificing operational strength Operation: The two arms of the claw actuate by piston motion, ultimately driven by the force of the nitinol wire. These arms close around the payload to secure it. The payload rests on the bottom scoops.

31 CLAW DESIGN SELECTION Decision Matrix Option 1 Option 2 Option 3 Option 4 Option 5 Low size/mass Tightness of grip Simplicity of design Ease of integration into SPACE Structural strength Results: Option 1 Option 2 Option 3 Option 4 Option 5

32 OPTIONS FOR ACTUATION OF THE CLAW Nitinol Wire Nitinol wire is a shape memory wire made of an alloy of nickel and titanium. While cold it can be deformed and upon heating it returns to a preset shape. This characteristic means that it can be used for actuation by heating it with current from a power source. Electromagnets Electromagnets are made of closely packed spun wire that electricity is run through. When the electricity passes through the wire it creates a magnetic field that can be stopped by ending the electric current through the wire. This could be used to actuate a system by placing two electromagnets on opposite sides of something that needs to be pulled together. Motors Motors provide rotational motion instead of linear motion and may not work well in a Martian environment.

33 CONSIDERATIONS FOR CHOOSING AN ACTUATION MECHANISM Reliability is a key factor in the claw actuation mechanism. It needs to work perfectly in the harsh Martian environment. Control over the speed and movement of the claw is also very important. The power requirement needs to be small.

34 ACTUATION METHOD SELECTION Decision Matrix Nitinol Wire Electromagnets Motor Reliability Control Power Requirement Results: Nitinol Wire Electromagnet Motor

35 FINAL DECISION ON ACTUATION MECHANISM Nitinol wire would work better than electromagnets or motors because it is the most reliable option and gives more control over the exact movements of the claw. Based on the design of the claw it requires a single upward force to contract and this would be best achieved with nitinol wire and simple springs to provide the opposing force to either contract the nitinol. Electromagnets are more difficult to manufacture or buy with the exact specifications needed to actuate the claw. There is also twice as much room for error as two magnets would be needed to operate the claw. A motor would not work well either since it would have too high an RPM to move the claw in a controlled manner. In addition a motor would require the conversion of rotational motion to linear motion, which would result in many more moving parts than the nitinol wire setup would have. The best option for the nitinol wire is to buy preset compression springs. Shaping nitinol wire is a complicated process which would decrease the reliability. In addition, a compression spring that open the claw instead of a tension spring that holds it closed would reduce power requirements and drastically decrease the amount of time the wire would need to be in it s active state. Holding the claw shut with conventional springs will increase the reliability and feasibility of this design.

36 CONSIDERATIONS FOR CHOOSING A MOTOR We want a large torque, so that the motor doesn t stall. We re going to assume the weight of the cables, claw, and payload is 2 lbs, and that the radius of the pulley is 1 inch. To successfully rotate this, we re going to need 32 oz*in of torque. With a factor of safety of three, we re looking to find something with around 100 oz*in of torque. We do not want a motor with an RPM that s uncontrollably high, otherwise we risk damaging some part of the crane or rocket. A DC motor should be sufficient. They are cheap, simple, and widespread, making them easy to work with.

37 PULLEYS VS. GEARBOX The idea of using pulleys to mechanically assist the motor in lifting the crane was deemed impractical. Motors that are capable of lifting the payload without the need of assistance from a pulley exist, and are approximately the same cost as a less powerful motor. That would mean that purchasing a pulley would be more expensive than any saving gained from having a weaker motor. In addition, the pulley would now add unnecessary weight and complexity to the crane, which is a bad idea overall. A gearbox would be preferable to a pulley system. A gearbox would allow us greater control over the output torque and RPM, and would be a simple attachment to the front of the motor.

38 MOTOR CHOICES Decision Matrix Option 1 Option 2 Option 3 Torque RPM Cost Reliability Results: Option 1 Option 2 Option 3

39 FINAL DECISION FOR THE MOTOR The final motor comes with a gearbox included, which improves simplicity. With the gearbox, the output torque is 110 oz in, and the output RPM is 350. This will quickly and reliably get the claw up to the top of the crane, which is important in a timed challenge. It is a cheap option, and was recommended to us by another project team on campus, so we know that it will be reliable.

40 Hazard Cause Effect Pre RAC Mitigation Verification Post RAC Exposure to voltage. Exposure to heat. Personal exposure to moving parts. Movable ramps getting stuck. Burning through of nitinol. Contact with frayed wires being used to power motor, light, or nitinol. Short-circuiting of motor overuse of motor. Moving parts in the machinery can cause injury if they come into contact with a body. Either due to too much friction in ramp or lack of lubrication. Too much voltage is applied to the nitinol. Slight shock, and possible first aid injury. Slight burn, and first aid injury. 4D 4D All wires will be bought new, all solder joints will be covered either with tape or surrounded by another material (i.e. a container). Prior to working with the motor, after a long period of heavy use, the motor will be given at least a minute to cool down prior to being touched. The wires to the motor will be checked prior to usage of the motor. First aid injury. 4D During operation of the SPACE system, all team members will not be allowed to stand within 3 of the system. Damage to rocket. Loss of the nitinol for the rocket. 3C 4C During operation the ramps must be monitored. If any jamming occurs, the operation musts be stopped. Extra precaution will be taken by the electrical team to ensure that too much voltage is not applied. Prior to working with the equipment, all electricity supplies (i.e. batteries) will be disconnected. If the motor has been used at its highest settings for more than one minute, a leader will check that all other members do not touch the motor prior to a minute passing. Prior to usage the motor wires will be inspected. Any leads present will be required to ensure that this rule is followed for personal safety. Prior to the testing and working of any of these components, the GS must make a kill-switch in the case that the ramp gets stuck. When in operation, no personnel will be allowed by the SPACE system. When being set up, the GS team will be advised to be extra careful about the application of too much voltage due to a careless mistake. 4E 4E 4E 4D 4D

41 LAUNCH PAD

42 ASSEMBLY Major subassemblies include: Tripod structural support, leveling, prevent tip over. Masses to weight down the tripod to prevent tip over will be quantified by CDR A-frame structural support, attachment for separation mechanism Pivot ASSY interface between launch rail and damper system Launch Rail ASSY dynamic portion of LP Counterweight shift CG to allow actuation Launch Rail ASSY Pivot ASSY Counterweight Sep. Mech. A-frame 40 SPAC E Tripod

43 TRIPOD AND A-FRAME Designed to hold the launch rail in the horizontal position until the command is sent from the ground station confirming payload insertion. Made of aluminum tubing and 80/20, but will need additional mass to satisfy the requirement that it will not tip over in sustained 20 mph surface winds. Preliminary analysis shows that all stress in critical members have a safety factor against yield of at least 3. This will be confirmed by a FEM model in ANSYS by CDR. Each leg is equipped with leveling feet to level the launch vehicle and space with respect to ground.

44 GRAVITATIONAL ASSIST DESIGN The trade space for lifting the rail from the horizontal to 5 degrees from the vertical away from the crowd. An motor on the end of the rail, a motor attached to a pulley system, and a using gravity were considered to actuate the rail to the deployed positon. Using gravity was selected because of cost, the size of the motor needed to lift the rail, and added design complexity. Gravity lift the pad by having the CG of the launch rail ASSY on the counterweight side of the launch pad. This will provide the necessary torque to raise the rail. After being raised, the rail will hit a stopper plate that is set to 5 degrees from the vertical. A door latch system will be added to lock the pad in the vertical position.

45 TORSIONAL DAMPER SIZING Feasibility of the gravity actuated LP requires a rotational damper that can slow down the actuation of the pad to the vertical by a damper system. Linear dampers were not considered because of increased complexity of converted rotational to linear motion. In order to determine a range of damping coefficient required to achieve a linear velocity at the end of the counterweight of less than 2 mph, a MATLAB simulation was created to simulate the system as a torsionally damped pendulum with the following second order differential equation: θ + c I ZZ θ + lm sysg I ZZ sinθ = 0 Damping coefficient in the range of 18 to 20 Nm/rad/s should be sufficient to create a critically or over damped system assuming the system mass has an uncertainty of 50%. The adjustable Kinetrol LA Vane Dashpot on the bottom right is currently being tested to see if it satisfies this constraint. According to the specifications of the damper, this is within its operational range.

46 SEPARATION MECHANISM Nylon rope holds launch rail in equilibrium Rope severed using Nichrome wire High resistance wire capable of burning through nylon Moment imbalance on launch rail causes rail to move into vertical position Two Nichrome wires on separate circuits with separate power sources act as a redundancy system

47 SEPARATION MECHANISM Nichrome wire circuit closed when payload door is closed System is turned off once the launch vehicle is in its vertical orientation Further testing needed to mature design Nylon rope diameter Wire gauge Input Voltage

48 A-FRAME STRUCTURE Independent structure designed to support housing mechanism and SPACE subsystem Made out of 1 square aluminum tubing and custom aluminum machined joints

49 LP LP System Hazard Cause Effect Pre-RAC Mitigation Verification Nichrome wire does not cut through nylon rope of separation mechanism. Launch pad torsional damper is not properly sized for the system. Power source to nichrome wire fails Nichrome wire slips off of nylon rope. Chosen damper has too much damping. Chosen damper has too little damping. Separation mechanism fails and launch rail not moved to vertical position. Launch rail rotates at a high velocity or does not rotate. 2D 3C The nichrome wire setup will consist of two nichrome wire coils each with their own independent power source. The chances of both power sources failing is low. Having two separate nichrome wires coiled around the nylon rope means the separation mechanism is still functional if one slips off. The chances of both slipping off is low. Use an adjustable torsional damper that is capable of providing a damping coefficient that bounds the desired value. Purchase and test multiple damper options. Possible gearing of damper. Post- RAC Test torsional damper properties with a calibrated test step up prior to being attached to the competition launch pad. LP Launch pad counterweight is not large enough to actuate the pad to the vertical. Counterweight does not have sufficient mass to place the CG of the rail to the counterweight side of the pivot. System fails to actuate from the horizontal to the vertical. 3B Determine the optimal mass of the counterweight using MATLAB simulation of a torsionally damped pendulum. Make mass adjustable to allow system to be tuned on the field.

50 System Hazard Cause Effect Pre-RAC Mitigation Verification Post- RAC LP Launch rail will not be able to support the mass of the counterweight. 1 80/20 launch rail yields from the applied load. Launch rail plastically deforms and cannot complete the mission objective. 4D Perform static analysis for bending of the rail under the expected loads. Qualify the flight unit by performing a two times expected competition load. Create a full system FEM model. Stress calculations modeling the rail as a cantilever beam with the projected counterweight mass yields a safety factor of 2.5 against yield. LP LP LP Structural components yield beyond operational limits and need to be replaced. Launch pad actuates before the separation command is executed resulting in scoring penalty and possible injury to personal. Launch pad moves or tips over because of sustained winds up to 20 mph and wind gusts up to 35 mph. Structural members not designed for operational loads. Nylon rope tethering the launch pad in the horizontal position fails or system is misused. Tripod launch pad does not have sufficient weight to avoid undesirable movement during operation. Launch pad will not 4D operate as intended or require repair. Unexpected transition from the horizontal to vertical position. Damage to the entire competition segment or unsafe launching conditions. 2E 1E Reinforce rail with structural doublers. Perform static analysis on all critical components. Create a full system FEM model to identify possible failure modes prior to fabrication. Electronic circuit designed to not initiate separation mechanism until payload insertion. Launch personal understand the operation of the system and avoid excessive contact. Perform a static analysis and design for a safety factor of at least three against tipping over in wind gusts up to 35 mph. Stress calculations for key components have been performed to design components with a safety factor or 3 against yield.

51 IGNITER SYSTEM

52 SYSTEM REQUIREMENTS The igniter system (IGS) will autonomously install the igniter into the rocket motor, placing the igniter at the top of the fuel grain. The IGS will be controlled by the ground station.

53 DESIGN SELECTION: CRITERIA Criteria Mass Size Maneuverability Durability Reusability Complexity Location Cost Description Total mass of all system components Total size occupied by all system components Ability of the system to move the igniter system in and out of the motor Ability of the system to withstand rocket motor exhaust Ability of the system to fully function within two hours of the previous launch with minor to no repairs Number of moving components and the degree of coupling present between components Position of the motor relative to the launch vehicle Cost of the system

54 DESIGN SELECTION: OPTIONS The critical design feature is the mechanism to move the igniter. Options considered: Motor/Pulley System Stepper Motor Linear Actuator

55 DESIGN SELECTION Critical Factors Against Design Selection: The size of the stepper motor design option exceeded the available space on the launch pad rail between the blast deflector and the rocket The lack of a rigid attachment fixture between the dowel for the igniter and the motor did not allow for alignment between the center of the fuel grain and the dowel to be maintained when the launch pad was rotated for horizontal to vertical Motor/Pulley Stepper Motor Linear Actuator Mass Size Maneuverability Durability Reusability Complexity Location Cost SUM Scale: 1=Undesirable 2=Average 3=Desirable

56 SYSTEM OVERVIEW

57 SYSTEM OVERVIEW Component Number Component Name 1 Linear Actuator 2 Copper Rod 3 Dowel 4 Teflon Tube 5 Aluminum Block 6 Bottom Plate 7 80/20 Rail Fastener 8 ¼ -20 Threaded Rods 9 ¼ -20 Hex Nuts

58 LINEAR ACTUATOR The linear actuator will create motion purely along the axis of the rocket motor Rigidly fixed to the launch pad Copper rod translates motion of linear actuator to motion of dowel Will be positioned on the launch pad such that when it is fully extended the dowel will be at the top of the fuel grain IGS when the linear actuator is fully retracted Max loading = 200 lbs. Stroke length = Retracted length = IGS when the linear actuator is fully extended

59 PERFORMANCE CHARACTERISTICS Performance Characteristic Linear actuator stroke length Description Evaluation Metric Verification Metric The linear actuator will extend a distance equal to the stroke length listed by the manufacturer The linear actuator will have a stroke length of inches The linear actuator will be extended to its full length and the stroke length will be measured using calipers Linear actuator extension speed The linear actuator will insert the igniter into the motor in under 1 minute The linear actuator will have an extension speed of at least 12 inches per minute The linear actuator will be extended to its full length. The stroke length will be measured using calipers and an approximate time for full extension will be measured using a stopwatch.

60 INTERFACES Launch Pad The linear actuator will be bolted onto the launch rail using 80/20 rail fasteners and pre-existing holes on the linear actuator structure Ground Station Control of the linear actuator will be done through the ground station The ground station will be able to move the linear actuator in both directions, start, pause, and re-start using a series of buttons or switches Ground Launch System The igniter will be wired to the RSO ground launch system only

61 VERIFICATION PLAN Requirement Insert the igniter into the rocket motor Place the igniter at the top of the fuel grain Design Feature(s) to Satisfy Requirement A linear actuator will move the dowel into the rocket motor Positioning of linear actuator prior to movement Verification Plan Test the linear actuator to verify movement of the system Test the linear actuator to determine the stroke length and position the system such that the igniter is one stroke length away from the top of the fuel grain Status Pre-testing Pre-testing

62 System Hazard Cause Effect IGS IGS IGS Failure of linear actuator system. Failure of linear actuator motorground station electrical connection. Failure of igniter to ignite the rocket motor. Jamming inside the actuator device. Breakage of mechanical linear actuator components. Short circuit. Open circuit. Wire breakage. Faulty igniter. IGS Ignition failure. Placement of the igniter too far away from the top of the fuel grain. Linear actuator cannot move the igniter into the rocket motor and launch does not occur, equipment damage. 2D Pre- RAC HAZARDS/RISKS Linear actuator is not 2D supplied with power and cannot move the igniter into the rocket motor, therefore launch does not occur. Rocket motor ignition does not occur. Rocket motor does not ignite. 2D 2C Mitigation The linear actuator will be a commercially manufactured device. It will be tested to ensure proper function. It will be rigidly attached to the launch pad to prevent damage from launch pad rotation. All electrical connections will strong and robust. All electrical circuits will be inspected prior to use. Redundant circuitry will be used. Strain relief wire sections will be used. Use commercial igniters with maximum reliability rating. The length of the rocket motor will be determined by contacting the motor vendor. The linear actuator will have a sufficient stroke length, the dowel will be sufficiently long, and the dowel will be placed at the correct initial position such that, when the linear actuator is fully extended, the igniter is at the top of the fuel grain. An electrical circuit will record the position of the linear actuator and confirm that it has fully extended before ignition can occur.

63 System Hazard Cause Effect Pre-RAC Mitigation IGS Exposure of linear actuator to high voltage/current. Power surge. Improper circuitry. Linear actuator is damaged and cannot function, therefore launch cannot occur, equipment damage. 2D Surge protectors will be placed around all critical circuits. All circuits will be redundant and robust. Testing will be conducted with non-flight hardware to confirm correct circuitry. IGS Separation of copper rod from dowel or linear actuator. Failure of attachment mechanisms. Igniter cannot be moved into rocket motor and launch cannot occur. 2D The dowel inserted into the copper rod and then glued using Contact Adhesive to secure the joint. The copper rod will be mechanically fixed to the linear actuator to prevent motion. IGS Deflection of dowel prior to igniter insertion. Lack of stiffness in the dowel. Dowel cannot be inserted into rocket motor and launch does not occur. 2D The dowel will be made of a material that will have minimal deflection. A correct length dowel will be tested to determine that the defection experienced allows for proper insertion into the motor. IGS Deflection of dowel during linear actuator motion. Lack of stiffness in the dowel. Dowel scrapes or jams against the fuel rain and cannot be inserted properly, damage to motor, improper ignition. 2D The dowel will be made of a material that will have minimal deflection. A correct length dowel will be tested to determine that the deflection experienced allows for proper insertion into the motor.

64 GROUND STATION

65 REQUIREMENTS Breakdown of Requirements Master control for starting, stopping, and pausing all procedures Safety light indicating AGSE power turned on, flashing at frequency 1 Hz and solid while AGSE paused Electronic tracking device for recovery Reliable wireless communication to Raspberry Pi to enable payload ejection and tracking Modular electrical systems for instance, separate tracking devices for launch vehicle and untethered payload and recovery system independent of any payload electrical circuits. User friendly GUI minimizing operational failure during mission

66 VERIFICATION PLAN AND STATUS Requirement Solution Verification Plan Master control for starting, pausing, and/or stopping all autonomous procedures and subroutines. Safety light indicating AGSE power turned on, flashing at frequency 1 Hz and solid while AGSE paused Electronic tracking device for recovery Ground Station Control GSCM described in previous section, consisting of the key and on switch, pause switch, and killswitch. Safety lights added to both Ground Station Control GSCM and main panel of GS GUI such that will be clearly visible during launch operations to both control GSCM and GUI operators Use BigRedBee transmitter, attenuator, and radio to enable triangulation Install beeping sound module on recovery components to allow location by ear when nearby Receive GPS packets using BigRedBee transmitter and onboard GPS module to locate components by latitude and longitude Receive GPS packets using XBee-PRO 900HP modules and Adafruit MTK3339 GPS module Test control module using power supplies similar to those used by the AGSE. Run tests using a power supply similar to that of the AGSE, as explained above, and ensure that safety lights illuminate and blink appropriately Track rockets using both GPS and triangulation with BigRedBee and XBee transmitters at test launches. Subsystem will be driven to several different locations and tracked at each.

67 VERIFICATION PLAN AND STATUS Requirement Solution Verification Plan Reliable wireless communication for payload ejection enable and tracking Modular electrical systems for instance, separate tracking devices for launch vehicle and untethered payload and recovery system independent of any payload electrical circuits User friendly GUI minimizing operational failure during mission Limit wireless communication modules; run wires between ground subsystems, and only use XBee module for payload and launch vehicle Experimentation to enhance confidence in wireless communication reliability Compress GPS and auxiliary data before transmission Use MARSA54L altimeter to trigger payload ejection mechanism upon reading the specified altitude Develop prototypes and conduct early testing to mitigate risks or communications failure. Develop visually effective user interface. Train mission control operator through test flights and simulation. XBee wireless communication will also be tested at a test launch as well as separating the XBee modules across several different distances (i.e., driven around and at the Cornell Plantations) Run simulations on altimeter, and verify payload ejection at test launches by logging altitude at ejection Each subsystem will be tested individually, as described above. The GS GUI will be tested using data from an Open Rocket Simulation (as the old GS GUI currently does), as well as during test launches.

68 SPACE MOTOR CONTROL DIAGRAM Voltage will be sent to the CLAW of the SPACE. This will open the CLAW since it starts in the CLOSED Position The Arduino activates one of the two pins (D2 or D3) shown in the diagram above to supply current to the motor. The motor s rotations will send periodic HIGH signals back to the Arduino which will be counted to know when to stop supplying current to the motor The motor is halted once the threshold count has been met by setting the pin on the Arduino to LOW again. The voltage supplied to the CLAW will be set to a LOW. This will force the CLAW to clamp on to the payload. The CLAW will be used to transport the payload. Current will be supplied to the motor in the reverse direction, by activating the other of the two pins from the one used in 3.2. This will force the bidirectional motor to spin in the other direction, pulling the CLAW and the payload up. The motor wil-l be moving until threshold rotation count - when the CLAW is above the ramp.

69 CONTROL INTERFACE WITH AGSE AND LAUNCH VEHICLE A hardwired Ground Station Control Module (GSCM) will be developed to start, stop, and/or pause AGSE procedures and subroutines. This GSCM will have multiple switches that will be capable of starting the AGSE subroutines as well. This GSCM will feature a set of LEDs that will serve as indicators for the completion of a subroutine. There will be LEDs to indicate the status of procedure (paused/active), the success of the payload storage, the achievement of 5 verticality, and also the successful placement of the igniter. The last light will turn ON to indicate a successful launch.

70 GROUND STATION GUI Real-time trajectory plotting and tracking Interface for XBee wireless communication Subsystem status display (e.g., separated sections) and control (e.g., payload ejection enable)

71 Hazard Cause Effect Pre- RAC Mitigation Verification Post- RAC Kill switch failure Communicatio ns failure between switch and launch vehicle due to faulty/severed connection Unable to abort launch vehicle start sequence in an emergency launch vehicle launches before all settings are ready, including correct vertical launch orientation 1D Use a high quality, low fail rate switch Connect launch vehicle via high quality and reliable connection Connect redundant secondary kill switch as an added precaution Do a test launch and run kill procedure to ensure proper functionality of the system 1E May lead to injuries Igniter failure Arduino/Raspb erry Pi hardware or software failure Communicatio ns failure in counting steps taken by stepper motor Limited to moderate damage to personnel health, equipment or environment 2C Inspect the Arduino/Raspberry Pi prior to insertion into the on-board electronics system Run test cases to verify software correctness Test communications with ground station Have a redundant Xbee handling communications in a second channel Inspect hardware onsite Routine update and inspection of software Check robustness of communications during launch 2D Malfunction of accelerometer and/or barometer on altimeter Damage during launch, faulty device Payload ejected too early, too late, or not at all Parachutes deploying too early, too late, or not at all 3D Keep altimeter out of direct sunlight to prevent barometer malfunction Have backup altimeters on hand Test all altimeters on launch day to ensure proper function Have trial runs to test integrity of accelerometers and barometers in launch test all altimeters on launch day for functionality 4E

72 Hazard Cause Effect Pre- RAC Mitigation Verification Post- RAC GPS receiver not functioning Damage during launch Device malfunction Invalid data being transmitted Unable to track or locate launch vehicle via GPS 2C Use multiple GPS modules for redundancy Use backup radio tracker Use high quality GPS module(s) with low failure rates Run test runs of GPS module and radio tracking module in launch vehicle to test integrity of module Test-run known route to check for data integrity 3E Arduino and/or Raspberry Pi programmed or wired incorrectly Running the wrong version of the program Mixing up pin connections Electronic controls operate incorrectly or not at all Data processed incorrectly or not collected 2C Have multiple people check for mistakes in the connection based on a schematic Multiple, independent human checks and verification of system based on design from a successful test 2D Interference impedes wireless communication channels between XBees on the launch vehicle and Ground Station Bad weather conditions and terrain inhibits transmission Surrounding magnetic field causes interference Other sources, such as radio stations, cause interference Limited tracking functionality increased difficulty locating the launch vehicle and sending data between the launch vehicle and Ground Station 3B Test the XBees in the location and environment anticipated for launch day Use multiple XBees for redundancy Utilize signal amplifiers to emit higher amplitude transmission Test the XBees in the location and environment anticipated for launch day 3D

73 Hazard Cause Effect Pre- RAC Mitigation Verification Post- RAC Arduino and/or Raspberry Pi system crash or freeze Overheating board Improperly terminating loops or recursion in the code Halted or terminated computing functions during operations 3B Implement a final inspection by running a test case program before launch Code the program running on the device to be as simple as possible and prevent programming styles that may take up too many system resources Run a trial program before launch and ensure communication of results to ensure that proper data is evaluated 3D Pause switch not functioning Communication between Ground Station and launch vehicle is suffering interference Unable to pause launch vehicle start sequence in case of a minor malfunction 2D Use a high quality model with a low failure rate Make secondary redundant physical pause switch Run trial communication with launch vehicle to ensure communication of pause command 3E Altimeter loaded with wrong program or wired incorrectly Human error Payload ejected too early, too late, or not at all Parachutes deploying too early, too late, or not at all 3D Institute a checklist and final inspection Run test flights with the same altimeter to be used on launch day Test run altimeters to verify that system is functioning properly 3E

74 Hazard Cause Effect Pre- RAC Mitigation Verification Post- RAC Power lost for onboard electronics Short circuit Leaking battery Electrical connections broken Severed communications between all components on the launch vehicle and between the launch vehicle and Ground Station Payload and parachute fail to deploy Unable to track launch vehicle 2C Secure all wires and electrical connections with tape, glue, or pins during construction Implement a checklist to ensure all electrical connections are secure Inspect all wiring before launch for incorrect placement and short circuits Implement a checklist to ensure a fresh battery is always placed in the launch vehicle Simulate low power situation Insert a spare power source and verify that it can provide the power needed to run all systems 3D Test and label batteries before and after each use Implement a backup battery Not enough power supplied to onboard electronics Damaged/destroyed wires Damaged batteries Battery leakage Uncharged batteries Short circuit Glitched communications between all components on the launch vehicle and between the launch vehicle and Ground Station Payload and parachute fail to deploy Unable to track launch vehicle 3C Add an extra battery Use a high quality battery that lasts longer on a single charge or use Implement a checklist to ensure a fresh battery is placed in the launch vehicle Test and label batteries before and after each use Ensure that wires are all secured Run a pre-launch check to ensure that all batteries are in working condition Check all wires are in working condition Install backup battery/power source 4D

75 FULL AGSE SYSTEM ASSEMBLY Fully integrated AGSE CAD model. A prototype of the Launch Pad will be completed prior to CDR to ensure all interfaces are fully defined.