University of Michigan Flight Readiness Review Project Wolverine: Butterfly Valve Drag Variation for Mile High Apogee 2011-2012 University Student Launch Initiative Department of Aerospace Engineering University of Michigan 3012 Francois-Xavier Bagnoud Building 1320 Beal Avenue Ann Arbor, MI 48109-2140
Table of Contents 1. Summary... 5 1.1 Team Name and Location... 5 1.2 Administrative Staff... 5 1.3 Launch Vehicle Summary... 5 1.3.1 Vehicle Size... 5 1.3.2 Motor Choice... 5 1.3.3 Rail Size... 5 1.3.4 Recovery System... 5 1.4 Science Experiment... 5 2. Changes Made Since Proposal... 6 2.1 Changes Made to Launch Vehicle... 6 2.2 Changes Made to Science Experiment... 6 2.3 Changes Made to Activity Plan... 6 3. Vehicle Criteria... 7 3.1 Design and Construction... 7 3.1.1 Structural Elements... 11 3.1.2 Electrical Elements... 12 3.1.4 Flight Reliability Confidence... 14 3.1.5 Test Data and Analysis... 15 3.1.6 Workmanship and Mission Success... 16 3.1.7 Safety and Failure Analysis... 16 3.1.8 Full-scale Launch Test Results... 24 3.1.9 Mass Report... 25 3.2 Recovery Subsystem... 26 3.2.1 Robustness of Recovery System... 26 3.2.2 Recovery System Breakdown and Justification... 29 3.2.3 Safety and Failure Analysis... 29 Failure... 30 Causes... 30 Effect... 30 Risk Mitigation... 30
3.3 Mission Performance Predictions... 30 3.3.1 Mission Performance Criteria... 30 3.3.2 Flight Profile Simulations... 31 3.3.3 Thoroughness and Validity of Analysis... 32 3.3.4 Stability Margin... 33 3.3.5 Kinetic Energy... 34 3.3.6 Altitude and Drift... 35 3.4 Verification... 35 3.5 Safety and Environment... 39 3.5.1 Safety and Mission Assurance... 39 Failure Modes... Error! Bookmark not defined. 3.5.2 Hazards... 42 3.5.3 Environmental Concerns... 44 3.6 Payload Integration... 44 4.0 Experiment Criteria... 48 4.1 Experiment Concept... 48 4.2 Science Value... 48 4.2.1 Science Experiment Objectives... 48 4.2.2 Mission Success Criteria... 48 4.2.3 Logic, Approach, and Investigation... 48 Dynamic Targeting... 50 4.2.4 Meaningfulness of Test... 51 Restrained Controller... 51 4.2.5 Relevance of Expected Data... 57 4.2.6 Experimental Process Procedures... 57 4.3 Experimental Design... 60 4.3.1 Design and Construction of Experiment... 60 4.3.2 Repeatability of Measurement... 60 4.3.3 Flight Performance Predictions... 60 4.3.4 Workmanship... 60 4.3.5 Test and Verification... 60 4.4 Verification... 60 4.4.1 Requirement Satisfaction and Verification... 60 4.5 Safety and Environment... 62
5.0 Launch Operations Procedures... 63 Recovery Preparations... 64 Motor Preparation... 64 Igniter Installation... 64 Setup on Launcher... 65 Troubleshooting... 65 Postflight Inspection... 65 6.0 Activity Plan... 66 6.1 Budget Plan... 66 6.2 Timeline... 70 6.3 Educational Engagement... 72 7.0 Conclusion... 75 8.0 Appendix A: Minimum Altitude Calculations (Valves Closed)... 76
1. Summary 1.1 Team Name and Location The University of Michigan Rocket Engineering Association (MREA) from Ann Arbor, Michigan proposes Project Wolverine. 1.2 Administrative Staff MREA s administrative staff member is Iain D. Boyd. The team mentor is Matt Schottler and he holds a Tripoli Level 2 Certification. 1.3 Launch Vehicle Summary The flysheet for Project Wolverine can be found on our website (http://www.umich.edu/~mrea/). 1.3.1 Vehicle Size Outer diameter: 5.52 Inner diameter: 5.36 Total length: 114.00 1.3.2 Motor Choice The motor type for the vehicle is a Cesaroni L-1395 motor (75mm diameter) with 4,895 N-s of total impulse. This determination was made based on Rocksim simulations, hand calculations, and multiple test launches. 1.3.3 Rail Size Project Wolverine will use 2 buttons and launch from a 120x1x1 in 3 rail. 1.3.4 Recovery System To ensure a safe recovery, Project Wolverine will separate twice during the descent. At apogee Project Wolverine will deploy a drogue parachute (36in) which will slow the descent to about 65 ft/s. Then, depending on launch conditions, the main parachute (168in) will be deployed between 1000ft and 500ft to decelerate the launch vehicle to about 17 ft/s for touchdown. Both ejection charges will be ignited by the flight computer. 1.4 Science Experiment The projected experiment for MREA s Project Wolverine is two butterfly valves (separated 180 degrees) that vary pressure drag. During the ascent, the butterfly valves will be actuated based on real time data from GPS and flight computers to vary the pressure drag on the rocket. The goal of this experiment is to put a high impulse motor in the rocket that ensures the rocket will reach at least one mile at apogee, yet have the rocket only reach one mile at apogee due to the increased drag from the butterfly valves.
2. Changes Made Since Proposal 2.1 Changes Made to Launch Vehicle Since CDR the rocket weight has increased to about 38 pounds. This is due to an increase in the static margin from 1.5 to 1.8 (concrete ballast was added to the rocket). The length of the rocket was also increased to 114 inches in order to add the ballast. The parachute size has also increased to 14 ft to meet the kinetic energy requirements for the competition. No other changes have been made. 2.2 Changes Made to Science Experiment No changes have been made to the science experiment. 2.3 Changes Made to Activity Plan We completed a full-scale launch on March 18, and are planning on completing two more launches on April 1 and 14 so that we can perfect our program for valve actuation before the USLI competition.
3. Vehicle Criteria 3.1 Design and Construction The most radical design component of the launch vehicle are two side cans body tubes that extend from the tail of the rocket 59 forward that each contain a butterfly valve 8 aft of their tip. By altering the position of the butterfly valve during flight, a microprocessor can increase or decrease airflow through the tubes, thus changing the drag force on the entire vehicle during flight to target an altitude of exactly 5280 feet. Both the design and construction of the launch vehicle center around three distinct sections that come together to form a cohesive, functional unit. These sections are as follows: 1) the motor section 2) the Avionics Bay (AvBay) section 3) the nose section Here is a more detailed description of each section: NOTE: In all drawings, the following key is used: SHADED SECTIONS: coupling mechanisms (tubes, metal rings) DOTTED LINES: outlines of components inside outer tube DOTTED SECTIONS: solid components inside outer tube SOLID DOTS: sheer pins CROSSED CIRCLES: tapped metal screws 1) Motor section
As the most aft section of the rocket, the motor section is 35 inches long. It contains the motor mount (housing a Cesaroni 30 Pro75 motor casing) and all four fiberglass fins (Public Missile D-07 fin set). The fins are placed 90 degrees apart from each other, with two of them extending through a side can in order to provide stability to the cans and allow for the addition of aerodynamic fillets all the way up and down the sides of both cans. These fillets are really just one-fourth of a can tube epoxied to both the airframe and the side can to encourage laminar flow along the surface of the vehicle and reduce drag. All fins are attached with epoxy directly to the motor mount along their entire root chord. The cans extend all the way to the very aft of the vehicle to ensure that the flow exiting the cans does not interfere with the ability of the fins to correct the orientation of the rocket during flight. To ensure a safe and stable launch, the vehicle employs an AeroPack 75mm motor retention system. This two-piece system consists of a threaded inner ring rigidly attached to the motor mount via epoxied steel screws and an outer ring that screws on to the first piece to securely hold in the motor assembly. Additionally, both launch lugs (more specifically, buttons) are attached to this motor section by a fine-threaded screw into a tapped hole in the airframe and epoxy. Placing both buttons this far aft on the rocket greatly increases the probability that the rocket is as stable as possible upon exiting the launch rail, since both buttons stay on the rail for the maximum amount of time possible. Buttons were chosen over aluminum sliders because of their higher degree of play on the rail (decreasing friction upon rail exit) and easier attachment to the airframe.
To safely recover the launch vehicle, the motor section also houses a drogue parachute above the motor assembly. This drogue parachute deploys at apogee in order to decrease the descent velocity of the vehicle and ensure the effective eventual deployment of the main parachute at 500 feet AGL. The motor section is connected to the avionics bay with two sheer pins along this Apogee Cut. 2) Avionics Bay (AvBay) The avionics bay is the middle section of the vehicle and houses the entire payload and every piece of electronic equipment utilized by the rocket. This includes both the payload avionics, which control the actuation of the valves, and the flight avionics, which govern the ignition of the e-charges used to separate sections at apogee and 500 feet AGL. Although it is the smallest section of the vehicle with a length of 24, it is arguably the most important portion, since it contains the drag mechanism used to alter drag on the rocket during ascent. There is an Access Cut 16 forward of the bottom of this section along the actuation plane of the valves, allowing for easy insertion and removal of the valves. The side cans end at the top of this section. To address an aerodynamic as well as a safety concern during launch, the valves are located 8 below the top of the side cans instead of at the very top of the cans. These top 8 inches of tube act as a flow straightener, making sure that the flow across the valves is as parallel to path of flight as possible. Parallel flow minimizes damaging disturbing forces that might result from angled flow interacting with an actuated valve.
The AvBay is connected via shock cord to both the drogue and the main parachute to ensure a safe recovery. The vehicle is rigidly connected across the Access Cut during flight with eight fine-threaded screws that go through both the airframe and an aluminum coupler ring. This section is also attached to the nose section in a similar manner. 3) Nose section The nose section is the longest section of the rocket with 42 of airframe and a 13 nose cone for a total of 55. There are no cans attached to this section. It attaches to the AvBay along the AvBay Removal Cut with four screws that go both through the airframe and a coupling tube. The vehicle does not separate along this cut during flight; it is only present to allow access and extraction of the electronics board from the AvBay. Moving forward, the main parachute (a SkyAngle Cert-3 XL) occupies most of the space aft of the Main Chute Cut, located 34 below the tip of the nose cone and secured with sheer pins. The rocket separates along this cut at 500 feet AGL to deploy the main chute. Even further forward of the Main Chute Cut is a mass ballast system made up of up to 0-10 ~.5 thick, 5.5 diameter concrete rings. These rings are held together by two bookends made out of fiberglass and a nut that threads down the rod extending through the whole system. These rings are fit snug against the inside of the nose cone and restrained from falling by a 2.5 body tube that contacts both the bottom of the rings and a bulkhead 11 aft of the bottom of the nose cone. The main design component in the nose section that contributes to a safe launch is the mass ballast system, located in the most forward compartment of the vehicle. This system allows for the easy adjustment of the rocket s Center of Gravity (CG) by adding or subtracting a significant amount of weight (~7 lbs with all 10 rings). The extremely forward location of the system maximizes the distance that the CG will move in response to a specific weight gain. Being able to adjust the CG in this manner makes it possible to ensure that the vehicle has a healthy stability margin on the pad and will not become unstable upon launch.
To successfully and safely recover the vehicle upon descent, the main parachute is located in this nose section. Because the chute so close to the AvBay section, the length of the wires to the e-charges used to separate the sections is minimal, reducing the chance that an e-charge breaks or fails during flight. 3.1.1 Structural Elements The MREA design has incorporated structural elements to enable multiple launches and recoveries without structural overhaul. The rocket airframe and side cans are constructed of Blue Tube, which is more resilient than traditional phenolic tubing to cracks upon landing. The recovery system incorporates both a drogue chute and main chute, both of which are securely fastened to the rocket fuselage and held by tubular nylon chord. The nylon chord ties into the chutes via 0.75 inch welded rings that are sewn into the fabric. On the fuselage side, metal eyelets were screwed into the bulkheads and epoxied to the inner diameter of the Blue Tube body. Additionally, the separation sections of the fuselage are supported by wooden bulkheads to provide additional rigidity after separation.
3.1.2 Electrical Elements The avionics are powered by three 9 volt batteries, one for the Raven2 flight computer and two for the Stratologger, drag computer, and servo. The battery retention system consists of three 9 volt battery holders positioned horizontally on the back of the vertically mounted fiberglass board with zip ties holding the batteries on. This system has flight heritage with this rocket and has never come loose or had any disconnections of any kind. The Raven2, Stratologger, and drag circuit board are retained by metal standoffs connected to the vertically mounted fiberglass board. The wiring connections are soldered to the key switch terminals and connected to the three computers via mounted terminal connections. The wiring is then twisted and zip tied together to keep it more organized. See Figure 1.
Figure 1: Recovery Avionics Each of these circuits are independent of each other with their own dedicated manually operated key switches. This gives our avionics two independent and redundant recovery systems. The Raven2 battery is connected to a positive power bus that supplies the voltage to the Raven2, the primary apogee e-charge, and the primary main chute e-charge. The negative wires from the e-charges feed into the Raven2 where their activation is controlled from, then are routed out of the Raven2 via the ground wire to the manually operated key switch, then to the negative terminal of the 9 volt battery. See Figure 2. The Stratologger batteries are connected in parallel to accommodate the fluctuation in loading due to the servo. The positive terminals are connected to the Stratologger and the input to the voltage regulator, applying 9 VDC to each. The Stratologger has positive and negative connections for the backup apogee and backup main chute e-charges and Figure 2: Drag System Avionics
sends an input signal to the arduino containing real time altimeter data. The voltage regulator circuit takes the 9 VDC and steps it down to 6 VDC to power the servo. The output is also connected to a series of 2 diodes, which each drop 0.5 volts, to drop the voltage to 5 VDC to power the arduino. The arduino then powers the accelerometer, the SD card interface, and sends the pulse width modulated signal to the servo. These devices are all connected to the same ground which is routed to the negative terminals of the 9 volt batteries via another manually operated key switch. 3.1.4 Flight Reliability Confidence Our overall mission goal of in-flight calculation and modification of drag coefficient in an effort to successfully control the vertical ascent of a rocket in order to attain a goal altitude requires a high level of confidence in all mission critical components. In addition to characterizing the performance and role of each component in a standalone setting, a battery of system integration tests were performed to assure that data fidelity was maintained across system interfaces, and that commands were executed by each subsystem effectively. During each step in the build process, functionality tests were run on components to allow for quick neutralization of encountered issues, as well as to reduce the burden on full system integrated tests. Hardware Flight Reliability Confidence The Arduino Nano 3.0 microprocessor, ADXL345 accelerometer, SL100 altimeter, and microsd breakout board were all initially tested when received to assure that they were performing to within manufacture specifications. In addition to standalone testing, the system was put together and all integration testing occurred on battery power with the entire system interfacing to assure any real system noise would be accounted for automatically during testing. During this testing it was discovered that the 20 HZ serial data input from the SL100 to the Arduino interrupted the ability to load new software onto the Arduino. To alleviate this issue, and to assure no components were damaged during software uploads a diode was added to the circuitry to inhibit the Arduino from acting as a power source during software updates. No other major issues were discovered, and system noise appeared to be minimal as calibration tests of each component in a standalone setting produced similar results to component testing in the fully integrated setting. Software Flight Reliability Confidence In order to assure that software was robust and could perform the necessary calculations in a timely manner, analysis was performed using noisy subscale test flight data, as well as test data from previous rocket launches performed by MREA. To simulate processer time required for reading data appropriate delays were added to the code where data input would need to occur. The results indicated that the Arduino microprocessor was able to easily perform all the calculations necessary to generate 20 command pulses per second to our servo as well as write all collected data to our microsd card. This testing also built confidence that noisy data was handled well by the software, and no faults would cause an interrupt of the code. Gear Box
During each step of the gear box build process, parts were modified to assure that they fit with reasonable tolerance and that alignment was accurate. To asses mechanical failure points of a given flap deflection command, the servo was attached to the gearbox, and one flap was rigidly held in place to assure it could not move. The system did not slip indicating that the power provided to the servo is the limiting factor on our system, and not a mechanical point of failure. This gave us confidence that even if one of the valves experienced much higher pressure than we expected, we would not have a mechanical failure of our system. Recovery System Mock flight simulations assured that the Raven flight computer and SL100 altimeter were producing significantly large current surges to our ejection charges when required without overburdening our power system. Static ejection charge testing was also performed to assure that stage separation occurred without over pressurization of our rocket body. Based on the assumption all of the 4F black powder was completely turned to a perfect gas, we determined 3 grams is required for apogee separation, and 4.5 grams for main chute deployment. 3.1.5 Test Data and Analysis As a result of the successful test launch in February, in which the full-scale vehicle was launched without any payload components, recent testing focused mainly on the payload and its ability to increase/decrease drag on the entire vehicle. Functional testing of the constructed drag mechanism was carried out by writing a program to run on the microprocessor that continuously swept the valves back and forth through 90 degrees at various speeds, which ranged from 10 degrees/sec to 60 degrees/sec. Using an old-fashioned protractor, it was observed that a full 90 degrees (± 2 deg) was only achieved under speeds slower than or equal to 30 degrees/sec. At faster speeds, the valves did not successfully reach their extreme positions or consistently respond correctly to a specific command from the microprocessor. A more quantitative test of the drag mechanism component was carried out with compressed air and a fully constructed AvBay section. As the microprocessor commanded the valves to sweep through 90 degrees, as in the former test, compressed air was shot down the side cans to better simulate the kinds of forces that the valves would see in flight. The valves were still able to turn up to the maximum psi that the air compressor (normally used for cleaning purposes) could create around 30 psi. Although this pressure was not as high as the pressures that the valves would actually experience during launch, the test was still an important step in fully verifying the ability of the valves to turn under pressure. Members of MREA have succeeded in reserving space in the University of Michigan s 4 by 6 wind tunnel on March 28 th and 29 th to further test this component. An exception to the payload testing was a test to determine the effectiveness of the mass ballast system in the nose of the vehicle in moving the CG of the rocket. This test was conducted by hanging the entire vehicle in its ready-to-launch configuration from a
string-pulley system in the MREA workspace. With no additional mass added to the vehicle, the CG was located ~78 aft of the nose cone, only 3 above the Center of Pressure. With 3.5 lbs added to the mass ballast system (half of its capacity), the CG moved forward approximately 7 inches to a point 71 aft of the tip of the nose cone. Combined with a successful test launch (stable flight, altitude 100 feet over a mile), this test verified that the mass ballast system could increase our stability margin to a healthy number without compromising an unacceptable amount of altitude. 3.1.6 Workmanship and Mission Success The interaction of each subsystem on the vehicle will be critical to the success of the mission. Since we are trying to actuate the drag during flight, we need the gear box and servos to interact efficiently and accurately with the control algorithm. If the butterfly valves do not rotate to the correct angle, the drag coefficient will be incorrect and the elevation of apogee will be altered. It is also crucial to align the fins symmetrically. We need the rocket to fly as vertical as possible to ensure the rocket has the potential to reach one mile at apogee. If we undershoot the one mile goal, our science experiment will not be tested during the flight. It is also critical to fly as vertical as possible so that the butterfly valves can receive the most mass flow, and therefore create more pressure drag on the rocket. 3.1.7 Safety and Failure Analysis The figure below shows the risk matrix for the MREA rocket. The matrix plots the most likely foreseeable failure events on likelihood versus impact. The following tables (Table 1-8) below details possible failure events for various aspects of the project, including boost phase of flight, the control subsystem, environmental concerns, launch operations, personnel hazards, scheduling and planning risks, the recovery subsystem, and the structural integrity of the vehicle. Additionally, these tables include mitigation plans, meant to decrease likelihood and impact of the possible risks. Risk ID Risk Item Mitigation Plan Likelihood Impact BOOST- 01 BOOST- 02 Rocket Engine Failure Motor Becomes Dislodged During Flight Extra igniters will be on hand to replace a failed engine. Motor will be securely attached to rocket via a retention system. Table 1: Failure Modes for Launch Phase of Flight Risk Index 1 2 2 2 2 4 Risk ID Risk Item Mitigation Plan Likelihood Impact Risk Index
CONTROL- 03 CONTROL- 04 CONTROL- 05 CONTROL- 06 CONTROL- 07 Control System Deployment Failure Microcontroller in-flight failure Actuator Servos or Gearbox Jams Only a Single Flap Deploys Vehicle becomes Unstable During Flight The flaps will be aligned in the upright position on the pad so that they will act as fins if computer fails. Extensive testing of the flight computer system will be performed, including static tests, wind tunnel tests, and flight tests. The servo system and gearbox will be verified extensively before launch under static loads, wind tunnel tests, and test flights. A correction mechanism will be programmed into the control algorithm to return single flap to the upright position if detected. Extensive testing on controls will also be performed. Verification of stability will be performed through CFD analysis, wind tunnel tests, and test launches. Table 2: Failure Modes for Controls Subsystem 3 3 9 3 4 12 2 3 6 2 5 10 2 5 10 Risk ID Risk Item Mitigation Plan Likelihood Impact ENVIRON- 08 Extreme Windy Conditions, Heavy Rain, or Snow Launch will be scrubbed and rescheduled. Risk Index 3 3 9
ENVIRON- 09 ENVIRON- 10 ENVIRON- 11 ENVIRON- 12 High Winds Parachute deployment altitude will be lowered to decrease drift distance. In the case of test launch, smaller motor may be used to reduce overall altitude. 4 2 8 The rockets exterior, to avoid wetting of the ignition strips, covers Mild Rain the ignition system. The exterior structure of the 3 1 3 rocket is durable enough to withstand light rain. GPS recovery system will help locate rocket. Launch may be Cloudy Day delayed or a smaller 3 1 3 motor may be chosen to reduce chance of loss of launch vehicle. The angular orientation of the rocket will be Mild Wind adjusted according to 5 1 5 launch day wind conditions. Table 3: Failure Modes for Environmental Concerns Risk ID Risk Item Mitigation Plan Likelihood Impact OPS-13 OPS-14 Vehicle Damaged During Transport Ejection Charges Ignite Before Launch The vehicle will be transported in a separate compartment to ensure it will not be damaged by spare parts. Ejection charges will not be armed and igniters will not be installed until vehicle is on launch pad to prevent a premature launch. A switch on the exterior of the rocket will arm the flight computers after the rocket is on the pad. Risk Index 1 5 5 1 4 4
OPS-15 Spectator is Harmed by Launch Vehicle Announcements will be made to warn spectators that a vehicle is being launched and spectators will be positioned a safe distance from the launch pad. Table 4: Failure Modes for Launch Operations 1 5 5 Risk ID Risk Item Mitigation Plan Likelihood Impact PERS-16 PERS-17 Epoxy toxic fumes, skin irritation, eye irritation APCP Rocket Propellant skin irritation, inadvertent ignition, burns to skin Team members will work in well-ventilated areas, wear face masks, gloves, and goggles at all times during construction. Gloves will be worn at all times to prevent skin irritation. Propellant will be kept away from ignition sources, such as matches, igniters, heat sources, and stored in proper Type 3 or Type 4 magazines to prevent inadvertent ignition. After motor burn, the team will wait 15 minutes before disassembling the motor, while wearing insulated gloves to prevent burns to skin. Risk Index 2 3 6 1 5 5
PERS-18 PERS-19 PERS-20 Cuts and lacerations, burns to skin, eye irritation due to operation of machinery Ejection Charge Test skin irritation, eye irritation, inadvertent ignition, burns to skin Vehicle Test flying debris, eye irritation, inadvertent ignition, burns to skin Wilson Center training is required before using machinery so that proper instruction is taught to everyone. Goggles will be worn at all times to prevent eye irritation. The Wilson Center has more rules that will be followed as well. Personnel involved will remain a safe distance away from ejection charges and the test will be done in an open area. All NAR regulations will be obeyed to provide safety for anyone involved. Table 5: Failure Modes for Personnel Hazards 2 4 8 1 5 5 2 5 10 Risk ID Risk Item Mitigation Plan Likelihood Impact PLAN-21 PLAN-22 PLAN-23 Test Launch Cancellation Late Delivery of Materials Major components need to be redesigned or reconstructed after test flight Multiple dates are scheduled to account for scrubbed launches. Extra launches can be added in cooperation with Jackson Model Rocketry Club. Lead time is left in the schedule to account for delays in materials acquisition. Extra time is left in schedule to account for changes in design. Risk Index 3 3 9 2 1 2 3 4 12
PLAN-24 PLAN-25 Winter break and spring break cause slip in schedule Unforeseen costs result in higher budgetary needs Planned schedule requires minimal time commitment from team members over break. Will apply for extra funding from Michigan Student Assembly and University of Michigan Engineering Council (both accept rolling funding requests). Possibility of asking some team members to cover some of cost of travel. 2 2 4 2 4 8 Table 6: Failure Modes for Scheduling and Planning Concerns Risk ID Risk Item Mitigation Plan Likelihood Impact RECOV- 26 RECOV- 27 RECOV- 28 RECOV- 29 RECOV- 30 RECOV- 31 RECOV- 32 Electric Match Failure Recovery System Flight Computer Failure Electric Match Failure Pressure Senor Inaccuracy Recovery Location System Failure Delayed Time Charge Failure Wire Disconnect There are redundant pyro-charges for the main parachute. Testing of the flight computer. There are redundant pyro-charges for the main parachute. There are redundant sensors on the flight computer. The GPS-based recovery system will be tested. There are redundant charges and two electric matches. All wiring in the rocket will be carefully placed and staked. Risk Index 2 5 10 2 4 8 2 5 10 2 5 10 3 2 6 3 5 15 1 4 4
RECOV- 33 Shock Chord Fails to Hold Parachute Together A ½-inch thick cord fastened securely to the rocket and parachute will be used. Safe charge amounts will be calculated and tested. Table 7: Failure Modes for Recovery Subsystem 1 3 3 Risk ID Risk Item Mitigation Plan Likelihood Impact STRUCT- 34 STRUCT- 35 STRUCT- 36 STRUCT- 37 Parachute Damage or Stuck in Airframe Bond between Valve Tubes and Airframe Fails Rocket Explodes during launch Fins Separate During Launch Careful review of the parachute packing and rocket sizing. Structural integrity will be verified through FEA and static load testing. The rocket body is constructed with a highstrength polymer frame (blue tube). Test launches and static load tests will verify structural integrity. Fins will be attached to the engine mount by going through the body of the rocket. The fins will be attached at two points with fillets on the body tube. Risk Index 1 3 3 2 4 8 1 5 5 1 3 3 Table 8: Failure Modes for Structural Integrity of Launch Vehicle
The results from these tables are combined and mapped in the risk matrix shown below in Table 9. This matrix shows detrimental failures in red, recoverable failures in yellow, and failures with a minimal effect in green. 5 ENVIRON- 12 LIKELIHOOD 4 3 ENVIRON- 09 RECOV-30 ENVIRON- 10 ENVIRON- 11 CONTROL- 03 ENVIRON- 08 PLAN-21 CONTROL- 04 PLAN-23 RECOV-31 2 PLAN-22 PLAN-24 BOOST-02 CONTROL- 05 PERS-18 PERS-20 PLAN-25 CONTROL- 06 CONTROL- 07 RECOV-26
1 PERS-16 RECOV-27 RECOV-28 RECOV-32 RECOV-29 OPS-13 RECOV-33 OPS-14 OPS-15 BOOST-01 STRUCT-34 STRUCT-34 PERS-17 STRUCT-37 PERS-19 STRUCT-36 1 2 3 4 5 IMPACT Table 9: Overall Project Events Failure Matrix While MREA does not foresee any detrimental risks, it does have a number of risks that reside in the medium criticality region. MREA has discussed the extent of these risks and believes that current risk mitigation plans will be sufficient to ensure mission success and safety of all personnel involved. However, MREA will continue to monitor these risks and develop more effective risk management strategies throughout the design process. 3.1.8 Full-scale Launch Test Results The full-scale test launch was successful, proving the design of the rocket was stable, the motor was correct for our experiment, and the rocket was still in launch condition after landing. Figure 3 below shows the rocket reached a maximum altitude of 5,375 ft and velocity of 690 ft/s during the flight. The results of the full-scale launch will be analyzed in more depth in Section 3.3.3. Figure 3: Full-Scale Test Launch Data from Raven Flight Computer
3.1.9 Mass Report Part Diameter [in] Length [in] Unit Mass [g] Quantity Net Mass [g] BODY TUBES Tube 1 [body] 5.5 18 461.3 1 461 Tube 2 [body] 5.5 21 538.1 1 538 Tube 3 [body] 5.5 6 153.8 1 153 Tube 4 [body] 5.5 10 256.3 1 256 Tube 5 [body] 5.5 34.8 891.8 1 892 Tube 11 [Coupler] 5.5 12 301.4 3 904 AVIONICS BAY Tube 6 [Av Bay] 4 18 271.9 1 272 Centering Rings [Av Bay] 5.5o,4.0i 0.25 30 6 180 Av Bay n/a n/a 1000 1 1000 Can flap axle 0.5 10.5 105 1 105 Av Bay Parts [rods/bulkheads] n/a n/a 328.1 1 330 Av Bay Aluminum Parts n/a n/a 1140 1 1140 MOTOR [dry] Centering Rings [motor mount] 5.5o,3.0i 0.25 40.1 4 160 Motor Retention System 5.5o,3.0i 1.3 145 1 145 Motor Thrust Plate 5.50,3.0i 0.25 100 1 100 Tube 7 [Motor Mount] 3 24 228.7 1 230 CANS Tube 8 [Top Can Section] 2.5 8 81.7 3 245 Tube 9 [Mid Can Section] 2.5 10 102.1 3 306 Tube 10 [Low Can Section] 2.5 34.8 355.3 3 1066 MISC Nose Cone 5.5 13" + 4" Sleeve 360.0 1 360 Shock Cord n/a [package] 190 1 190 Fins [Missile Fin-D-07 ] SA per fin: 80.4 314 4 1254 Gear Housing / Flaps n/a n/a 1500 1 1500 Ballast n/a n/a 1300 1 1200 MOTOR Motor Case 3 30 2500 1 2500
Sub Total [dry] Mass: 15837 Total [dry] Mass [5%wt Epoxy]: 16261 Total [dry] Mass [10%wt Epoxy]: 17036 3.2 Recovery Subsystem 3.2.1 Robustness of Recovery System 3.2.1.1 Structural Elements The structural elements of the recovery system consist of the bulkheads, eyebolts, and shock cords. The bulkheads are made of 3/4 inch wood with an eyebolt mounted though the middle of them. There is one placed in the top body tube below the nosecone and two placed on opposite sides of the avionics bay enclosing the electronics and protecting them from corrosive e-charge exhaust. The top body tube bulkhead is fixed in place by epoxy while the avionics bulkheads are cinched together via 2 threaded rods and can be removed for avionics bay access. Shock cords connect the upper two bulkheads together via eyebolts, while the lower shock cord (apogee separation) is connected to the lower avionics bay bulkhead and the motor mount tube. This system has flight heritage with our team. It is the same system our team used last year with the exception of thicker bulkheads being used this year to accommodate the heaver loads from the larger mass. In our 2 test flights this year, the recovery systems structural elements have performed as expected with no signs of structural failure. 3.2.1.2 Electrical Elements The electrical elements of the recovery system consist of the e-charges, the e-charge wires, the e-charge terminals, the flight computers (Raven2 and Stratologger), and 9 Volt batteries. Clear canister ejection charges from Apogee are being used. Ground tests were performed using these e-charges to ensure the calculated amount of 4F black powder was correct and to increase our knowledge working with them. The e- charges are connected to a terminal board mounted on the outside of each avionics bay bulkhead. This allows for easy installation and removal of e-charged without having to open the avionics bay, and helps keep corrosive 4F black powder exhaust from entering the avionics bay. The terminal boards are routed through the bulkheads into the avionics bay where they are connected to the appropriate flight computer e-charge connections. The flight computers are powered by independent 9 volt batteries to give us 100% redundancy in our recovery system. 3.2.1.3 Redundancy The recovery system is 100% redundant to prevent a fault or failure in one system from affecting the other system so that the recovery system will still operate. The recovery system is based on two independent flight computers (Raven2 and Stratologger) each controlling 2 e-charges (one for apogee and one for the main chute). They are each powered by separate 9 volt batteries and have their own manually operated key
switches to turn them on. To further ensure redundancy, only the Stratologger altimeter readings are fed to the drag computer while the Raven2 remains independent of any connections with the drag computer or the Stratologger. 3.2.1.4 Parachute Sizes and Descent Rates The drogue chute size is 36 inches. This drops our initial decent rate from apogee from 105 ft/sec to 65 ft/sec. The velocity drop reduces the kinetic energy of the system and prevents the main chute from zippering the body tubes when it deploys. The main chute size is 14 feet. This chute is designed to give our rocket a decent rate of 15.56 ft/sec down from 27.0 ft/sec from the first test launch. The reduced descent velocity from the drogue chute reduces the kinetic energy of the system to prevent a body tube zipper upon main chute deployment. 3.2.1.5 Drawings and Schematics
Figure 5: Recovery Scheme Figure 4: Recovery Electronics The recovery scheme is shown below in Figure 4 and the electronics used to do this are found in
Figure 5. 3.2.1.6 Rocket- locating Transmitters The locating transmitter used is a Big Red Bee BRB900. The BRB900 transmits on a 900 MHz frequency using 100 mw with a range of approximately 6 miles. The GPS receiver will connected to a laptop computer via USB to monitor the location of the rocket in real time. The GPS will be positioned in the nosecone of the rocket to prevent any interference with the recovery avionics or drag system avionics. 3.2.2 Recovery System Breakdown and Justification The main chute is designed for rockets with a mass of 35-40 lbs. This gives the rocket a descent rate of 15.56 ft/sec. The main chute is attached to the shock cord which pulls it out of the upper body tube upon separation. This deployment process has been tested via ground tests with e-charges. The test consisted of the main chute separation being assembled and set up in flight conditions with the shock cord attached, main chute, and shear pins in the body tubes. The rocket was then positioned on two rolling chairs and strapped down to each with the separation between them. A four gram e-charge was detonated which separated the rocket. The main chute was successfully pulled out of the upper body tube and the results were recorded with possible improvement ideas to make the deployment better. The recovery system electronics are tested by performing simulation flights on the computers. An ammeter is connected to the terminals of the e-charges. The simulation flight is engaged and the computer simulates increasing altitude. Once the simulation reaches apogee, a pulse is measured on the ammeter of approximately 5 amps. As the simulation descends to 500 feet, a second pulse is measured by the ammeter of approximately 5 amps. The 5 amp pulses are what ignite the e-charges at the proper altitude and completes the test of the recovery system avionics. 3.2.3 Safety and Failure Analysis The recovery system had a failure during the last test launch. Once the rocket descended to 500 feet the e-charges fired, which separated the rocket, but the main chute failed to deploy. Upon investigation of this failure, it was determined that the failure was due to three factors: Improper parachute folding/packing, shear pins were too long and held the main chute in the body tube, and the main chute was attached too close to the top body tube bulkhead causing the shock cord to become taught before the parachute cord bundle was fully deployed and the main chute fully out of the body tube. These failures have been tested and corrected with the following actions: Parachute will be packed per written procedure, shear pins will be cut to allow zero overhang into the body tubes, and the shock cord will be lengthened and the main chute attached at the distance that is greater than the length of the parachute cords. These changes will ensure the main chute deploys when the e-charge fires.
Failure Causes Effect Risk Mitigation E-charge doesn t fire 1. Electrical fault 2. Battery disconnection 3. Flight computer programmed improperly Rocket crash lands Use redundant recovery system (flight computers, e- charges, switches, batteries) Main Chute fails to deploy 1. Improper parachute packing 2. Shear pins too long 3. Parachute tied too close to the top of the shock cord Rocket crash lands 1. Use procedure when packing the parachute 2. Cut the shear pins flush with the body tube 3. Attach parachute far enough down the shock cord to allow the shock cord to pull the chute out of the body tube Main Chute cords fouled Improper parachute packing Rocket crash lands 1. Use procedure when packing the parachute Parachute deploys too early 1. Computer programmed incorrectly 2. Shear pins not installed 3. Pressure holes not drilled 1. Main chute deploys during accent 2. Main chute deploys too early during descent 1. Perform pre-flight testing on flight computers 2. Ensure shear pins are installed 3. Ensure all pressure holes are in place 3.3 Mission Performance Predictions 3.3.1 Mission Performance Criteria The following characteristics must be nominally satisfied for a launch to be successful:
1.The drag mechanism must remain in its neutral configuration (butterfly valves completely open) during engine burn to ensure a safe exit from the launch rail and entrance into a stable coast phase after burnout. 2. The control system must iteratively read in altimeter data and actuate the digital servo accordingly to adjust the drag coefficient of the vehicle. 3. The vehicle must reach an altitude of 5280 feet ± 2%. It is preferable that the rocket reaches an altitude slightly less than a mile rather than an altitude over a mile. 4. The vehicle must separate along the Apogee Cut at apogee. Drogue deployment must be successful; the rocket will begin to tumble in two connected sections. 5. The vehicle must separate along the Main Chute Cut at 500 feet, immediately deploying the main parachute. 6.After main chute separation, the three connected sections of the vehicle must fall to the earth without damaging any components. None of the sections may have a kinetic energy over 75 ft-lbs. 3.3.2 Flight Profile Simulations 14 12 10 Accelera'on (g) 8 6 4 2 0 0 0.5 1 1.5 2 2.5 3 3.5 Time (sec) Figure 6: Thrust Curve from Full-scale Launch
Using data gathered from the Raven flight computer during the test launch on March 18 th, the following motor thrust curve was found (Figure 6). The weight of the Aft section is the dry weight with no propellant inside the motor mount. Using the total rocket weight along with the data for the Cesaroni L1395 motor, we estimate the apogee altitude to be 5271 ft. This is within 2% of the altitude obtained on March 18 th of 5375 ft. 3.3.3 Thoroughness and Validity of Analysis Multipoint simulation verification ensured that our flight profile and simulations were operating under similar regimes. Two independent flight profile simulations, one written in Matlab and another using a Simulink discrete time integration method to model controller behavior provided results that were closely correlated with each other and with Rocksim. These results for our second test launch on March 18 are summarized below: Sim. 1 Sim. 2 Rocksim Raven Flight Data Burnout Velocity [m/s] 204 201 200 210 Burnout Altitude [m] 382 367 368 209 Apogee (no controller) [m] 1680 1648 1662 1638 Apogee Time [s] 18.3 17.6 17.4 17.9 One significant point of divergence that is apparent when comparing flight data to our projections is the burnout altitude is over predicted by all of our simulations by over 100 meters. This deviation from our nominal flight path can be attributed to the fact that the motor burn occurred at an angle which was significantly different from the assumption that the rocket would be launched perfectly normal to the ground plane. The picture of the rocket shortly after leaving the launch pad to your right shows that there is a large angle, being made with the vertical which is due in part to launch rail setting, and in part to weather cocking effects on our rocket. The stability margin of our rocket was approximately 2 at launch tending to generate a large angle stable flight condition. Overall we are not worried about this deviation from our predictions as the only real effect it will have is to delay the onset of controller activity by at most a fraction of a second. Our rocket also still managed to achieve over 5280 ft. of elevation with this configuration assuring us that we would not automatically undershoot from a larger than expected flight path angle. Our first full scale test launch was performed on Saturday, January 21 in Three Oaks, MI at 1:00 pm. In addition to verifying our 2 axisymmetric design under real flight conditions, we were able to determine a nominal coefficient of drag from our rocket
body. While the fidelity of the launch acceleometer was approx 0.1g we are still able to gain a good foundation for what behavior we can expect in future flight, and exactly what the dynamics of the rocket are. Our second test launch was performed on March 18 in Manchester Michigan at 4:00 pm. The resulting coast drag coefficent was indicated to be 0.82 with a graph that closely resembles the one below from our first launch. This greatly increased our confidence in the fidelity of the simulations we were running, and how the rocket body would react under different enviornmental conditions Figure 7: Drag Coefficient from Full-scale Launch 3.3.4 Stability Margin The CP of the rocket as determined by Rocksim is located approximately 81 inches from the nose of the rocket. The CG location was found via static ground testing of the test launch vehicle. After hanging the rocket by a string wrapped around the airframe at different points until it did not definitively tip to either side, it was determined that the CG lay approximately 79 inches aft of the tip of the nose. Due to the non-axisymmetric nature of our rocket, we determined that a more appropriate stability margin would be
closer to 2, and roughly 1.3 kg of ballast was added to the nose of the rocket to accomplish this. Once this was in place the CG was again experimentally determined and found to be at approx. 72 inches from the nose of our rocket. CP CG 81 71 Stability Margin = = = 1.8 Diameter 5.5 3.3.5 Kinetic Energy Nose Cone section mass: = 4.62 kg Bottom Section mass: = 7.42 kg Avionics Section mass: = 4.33 kg Total Rocket mass (with no propellant) = 16.37 kg Max Boost Velocity = 210 m/s Terminal Velocity (/w Drogue) = 20 m/s Main Chute Velocity = 7 m/s KE!""#$ (!"#$%) = 1 2 16.37 kg 210 m s! = 361 KJ KE!"#$"%& (!"#$%) = 1 2 16.37 kg 20 m s! = 3.27 KJ KE!"#$"%& (!"#$%&!"##"$!"#$%&') = 1 2 8.95 kg 20 m s! = 1.79 KJ KE!"#$"%& (!"#$%&!"#!"#$%&') = 1 2 7.42 kg 20 m s! = 1.48 KJ
KE!"#$"%& (!"#$!"##"$!"#$%&') = 1 2 7.42 kg 7 m s! = 0.18 KJ KE!"#$"%& (!"#$!"#$%#&'!"#$%&') = 1 2 4.33 kg 7 m s! = 0.11 KJ KE!"#$"%& (!"#$!"#$!"#$!"#$%&') = 1 2 4.62 kg 7 m s! = 0.11 KJ The kinetic energy during the descent before the main parachute is deployed has been significantly lowered since our first test launch despite the increase in mass. This is due to a lower terminal velocity as a result of a drogue parachute being introduced. This was done because of body tube zippering that occurred on our January 21 test launch in Three Oaks, MI. We also are using a main parachute that has approx. 90 sq. feet of surface area to assure that our kinetic energy upon landing is minimized and to negate the 5kg increased weight since our first test launch. 3.3.6 Altitude and Drift In order to maintain a conservative estimate regarding drift distance, we simply consider the rocket traveling at wind velocity and look at the fall time from apogee. We assume our rocket will fall for 70 seconds based on our expected terminal fall velocity and a starting altitude of 5300 ft. d!"#$% = v!"#$ t!"## Wind Speed (mph) Drift (ft.) 5 513 10 1027 15 1540 20 2053 3.4 Verification 1. The launch vehicle shall carry a science or engineering payload: a. Option 1: The engineering or science payload may be of the team s discretion. The science payload for MREA s Project Wolverine is two butterfly valves (separated 180 degrees) that vary pressure drag. During the ascent, the butterfly valves will be actuated based on real time data from GPS and flight computers to vary the pressure drag on the rocket. The goal of this experiment is to put a high impulse motor in the rocket that ensures the rocket will reach at least one mile at apogee, yet have the rocket only reach one mile at apogee due to the increased drag from the butterfly valves. 2. The launch vehicle shall deliver the science or engineering payload to, but not exceeding, an altitude of 5,280 feet above ground level (AGL).
The motor chosen for flight is a Cesaroni L-1395 with 4,895 N-s of total impulse. In our full-scale test launch we reached an altitude of 5,375 ft (below the DQ altitude of 5,600 ft and within our threshold for the experiment). 3. The vehicle shall carry one Perfect Flight MAWD or ALT15 altimeter for recording of the official altitude used in the competition scoring. We will be using the Perfect Flight MAWD as our competition altimeter. 4. The recovery system electronics shall have the following characteristics: a. The recovery system shall be designed to be armed on the pad. b. The recovery system electronics shall be completely independent of the payload electronics. c. The recovery system shall contain redundant altimeters. The term altimeters includes both simple altimeters and more sophisticated flight computers. d. Each altimeter shall be armed by a dedicated arming switch. e. Each altimeter shall have a dedicated battery. f. Each arming switch shall be accessible from the exterior of the rocket airframe. g. Each arming switch shall be capable of being locked in the ON position for launch. h. Each arming switch shall be a maximum of six (6) feet above the base of the launch vehicle. The checklist for launch operation procedures details the order that the electronics shall be armed and operated. There are two exterior arming switches on the rocket which can be reached without a ladder, and the electronics run off separate batteries. There are also redundancies for each recovery stage. 5. The recovery system electronics shall be shielded from all onboard transmitting devices, to avoid inadvertent excitation of the recovery system by the transmitting device(s). The GPS and recovery electronics are housed in separate parts of the rocket. 6. The launch vehicle and science or engineering payload shall remain subsonic from launch until landing. The maximum velocity reached during flight is around 700 ft/s. 7. The launch vehicle and science or engineering payload shall be designed to be recoverable and reusable. Reusable is defined as being able to be launched again on the same day without repairs or modifications. We have launched the same rocket twice before, and plan to have four launches of the competition rocket before USLI. 8. The launch vehicle shall stage the deployment of its recovery devices, where a drogue parachute is deployed at apogee and a main parachute is deployed at a much lower altitude. Tumble recovery from apogee to main parachute deployment is permissible, provided that the kinetic energy is reasonable.
A drogue will be deployed at apogee slowing the rocket to about 65 ft/s and a main will be deployed around 500 AGL slowing the rocket to about 17 ft/s. 9. Removable shear pins shall be used for both the main parachute compartment and the drogue parachute compartment. Shear pins have been installed at all separation events. 10. The launch vehicle shall have a maximum of four (4) independent or tethered sections. a. At landing, each independent or tethered sections of the launch vehicle shall have a maximum kinetic energy of 75 ft-lbf. b. All independent or tethered sections of the launch vehicle shall be designed to recover with 2,500 feet of the launch pad, assuming a 15 mph wind. There are only three independent tethered sections on the rocket and the predicted drift is 1,540 ft in a 15 mph wind. Each independent section has a maximum kinetic energy of less than 75 ft-lbf. 11. The launch vehicle shall be capable of being prepared for flight at the launch site within 2 hours, from the time the waiver opens. Once the waiver opens the rocket motor needs to be assembled and then the electronics shall be checked. Then e-charges will be created and once the final rocket is put together the shear pins will be put into place. This process usually takes 75 minutes. 12. The launch vehicle shall be capable of remaining in launch-ready configuration at the pad for a minimum of 1 hour without losing the functionality of any onboard component. The batteries onboard have been tested to remain functional for over 6 hours. 13. The launch vehicle shall be launched from a standard firing system (provided by the Range) using a standard 10 - second countdown. We will be following all NASA protocols at the launch. 14. The launch vehicle shall require no external circuitry or special ground support equipment to initiate the launch (other than what is provided by the Range). The only external circuitry needed is the ignition charge. 15. Data from the science or engineering payload shall be collected, analyzed, and reported by the team following the scientific method. We will be following the scientific method and presenting our results in PLAR. 16. An electronic tracking device shall be installed in each independent section of the launch vehicle and shall transmit the position of that independent section to a ground receiver. Audible beepers may be used in conjunction with an electronic, transmitting device, but shall not replace the transmitting tracking device. We will be using the Big Red Bee GPS to track our rocket throughout the flight.
17. The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR), Tripoli Rocketry Association (TRA) and/or the Canadian Association of Rocketry (CAR). The rocket motor is an L-1395 from Cesaroni (4,895 N-s total impulse). 18. The total impulse provided by the launch vehicle shall not exceed 5,120 Newtonseconds (L-class). This total impulse constraint is applicable to any combination of one or more motors. The rocket motor is an L-1395 from Cesaroni (4,895 N-s total impulse). 19. All teams shall successfully launch and recover their full scale rocket prior to FRR in its final flight configuration. Currently the competition rocket has been flown twice in the final configuration. Both launches were conducted on L sized motors, and the second launch used the L-1395 motor we are using in the competition. Two more tests launches are planned before Alabama. 20. The following items are prohibited from use in the launch vehicle: a. Flashbulbs. The recovery system must use commercially available low-current electric matches. b. Forward canards. c. Forward firing motors. d. Rear ejection parachute designs. e. Motors which expel titanium sponges (Sparky, Skidmark, MetalStorm, etc.). f. Hybrid motors. None of the above items are used in our launch vehicle. 21. Each team shall use a launch and safety checklist. The final checklist shall be included in the FRR report and used during the flight hardware and safety inspection and launch day. We will be using the launch operations procedures checklist in this report. 22. Students on the team shall do 100% of the work on the project, including design, construction, written reports, presentations, and flight preparation with the exception of assembling the motors and handling black powder charges. Every member of the team is currently a student, and we have done all the work ourselves. 23. The rocketry mentor supporting the team shall have been certified by NAR or TRA for the motor impulse of the launch vehicle, and the rocketeer shall have flown and successfully recovered (using electronic, staged recovery) a minimum of 15 flights in this or a higher impulse class, prior to PDR. Our team mentor is a Tripoli level 2 cert and has participated in the minimum 15 flights required.
24. The maximum amount teams may spend on the rocket and payload is $5000 total. The cost is for the competition rocket as it sits on the pad, including all purchased components and materials and the fair market value of all donated components and materials. The following items may be omitted from the total cost of the vehicle: a. Shipping costs. b. Ground Support Equipment. c. Team labor. Our team has spent $2,800 on all components of the rocket, including test launches and spare materials. 3.5 Safety and Environment 3.5.1 Safety and Mission Assurance MREA s acting safety officer is Kayla Wizinsky. The safety officer is responsible for being active and alert throughout all fabrication, testing, and launches. The following failure modes are repeated from Section 3.1.7 for convenience. Risk ID Risk Item Mitigation Plan Likelihood Impact BOOST- 01 BOOST- 02 Rocket Engine Failure Motor Becomes Dislodged During Flight Extra igniters will be on hand to replace a failed engine. Motor will be securely attached to rocket via a retention system. Table 10: Failure Modes for Launch Phase of Flight Risk Index 1 2 2 2 2 4 Risk ID Risk Item Mitigation Plan Likelihood Impact CONTROL- 03 CONTROL- 04 CONTROL- 05 Control System Deployment Failure Microcontroller in-flight failure Actuator Servos or Gearbox Jams The flaps will be aligned in the upright position on the pad so that they will act as fins if computer fails. Extensive testing of the flight computer system will be performed, including static tests, wind tunnel tests, and flight tests. The servo system and gearbox will be verified extensively before launch under Risk Index 3 3 9 3 4 12 2 3 6
static loads, wind tunnel tests, and test flights. CONTROL- 06 CONTROL- 07 Only a Single Flap Deploys Vehicle becomes Unstable During Flight A correction mechanism will be programmed into the control algorithm to return single flap to the upright position if detected. Extensive testing on controls will also be performed. Verification of stability will be performed through CFD analysis, wind tunnel tests, and test launches. Table 11: Failure Modes for Controls Subsystem 2 5 10 2 5 10 Risk ID Risk Item Mitigation Plan Likelihood Impact OPS-13 OPS-14 OPS-15 Vehicle Damaged During Transport Ejection Charges Ignite Before Launch Spectator is Harmed by Launch Vehicle The vehicle will be transported in a separate compartment to ensure it will not be damaged by spare parts. Ejection charges will not be armed and igniters will not be installed until vehicle is on launch pad to prevent a premature launch. A switch on the exterior of the rocket will arm the flight computers after the rocket is on the pad. Announcements will be made to warn spectators that a vehicle is being launched and spectators will be positioned a safe distance from the launch pad. Table 12: Failure Modes for Launch Operations Risk Index 1 5 5 1 4 4 1 5 5
Risk ID Risk Item Mitigation Plan Likelihood Impact RECOV- 26 RECOV- 27 RECOV- 28 RECOV- 29 RECOV- 30 RECOV- 31 RECOV- 32 RECOV- 33 Electric Match Failure Recovery System Flight Computer Failure Electric Match Failure Pressure Senor Inaccuracy Recovery Location System Failure Delayed Time Charge Failure Wire Disconnect Shock Chord Fails to Hold Parachute Together There are redundant pyro-charges for the main parachute. Testing of the flight computer. There are redundant pyro-charges for the main parachute. There are redundant sensors on the flight computer. The GPS-based recovery system will be tested. There are redundant charges and two electric matches. All wiring in the rocket will be carefully placed and staked. A ½-inch thick cord fastened securely to the rocket and parachute will be used. Safe charge amounts will be calculated and tested. Table 13: Failure Modes for Recovery Subsystem Risk Index 2 5 10 2 4 8 2 5 10 2 5 10 3 2 6 3 5 15 1 4 4 1 3 3
Risk ID Risk Item Mitigation Plan Likelihood Impact STRUCT- 34 STRUCT- 35 STRUCT- 36 STRUCT- 37 Parachute Damage or Stuck in Airframe Bond between Valve Tubes and Airframe Fails Rocket Explodes during launch Fins Separate During Launch Careful review of the parachute packing and rocket sizing. Structural integrity will be verified through FEA and static load testing. The rocket body is constructed with a highstrength polymer frame (blue tube). Test launches and static load tests will verify structural integrity. Fins will be attached to the engine mount by going through the body of the rocket. The fins will be attached at two points with fillets on the body tube. Risk Index 1 3 3 2 4 8 1 5 5 1 3 3 Table 14: Failure Modes for Structural Integrity of Launch Vehicle 3.5.2 Hazards Multiple procedures and protocols are in place to ensure a safe working environment. The work space is kept in a clean condition to prevent injuries and increase productivity. All potentially hazardous materials are securely kept with their MSDS in an easily locatable binder that is shared by all student groups in the workspace. All students must attend safety lectures where injury and accident prevention is discussed in-depth and to use any machining tools the student must have Wilson Center Training. In addition lab supervisors are present to enforce the rules and protocols. Risk ID Risk Item Mitigation Plan Likelihood Impact PERS-16 Epoxy toxic fumes, skin irritation, eye irritation Team members will work in well-ventilated areas, wear face masks, gloves, and goggles at all times during construction. Risk Index 2 3 6
PERS-17 PERS-18 PERS-19 PERS-20 APCP Rocket Propellant skin irritation, inadvertent ignition, burns to skin Cuts and lacerations, burns to skin, eye irritation due to operation of machinery Ejection Charge Test skin irritation, eye irritation, inadvertent ignition, burns to skin Vehicle Test flying debris, eye irritation, inadvertent ignition, burns to skin Gloves will be worn at all times to prevent skin irritation. Propellant will be kept away from ignition sources, such as matches, igniters, heat sources, and stored in proper Type 3 or Type 4 magazines to prevent inadvertent ignition. After motor burn, the team will wait 15 minutes before disassembling the motor, while wearing insulated gloves to prevent burns to skin. Wilson Center training is required before using machinery so that proper instruction is taught to everyone. Goggles will be worn at all times to prevent eye irritation. The Wilson Center has more rules that will be followed as well. Personnel involved will remain a safe distance away from ejection charges and the test will be done in an open area. All NAR regulations will be obeyed to provide safety for anyone involved. Table 15: Failure Modes for Personnel Hazards 1 5 5 2 4 8 1 5 5 2 5 10
3.5.3 Environmental Concerns Risk ID Risk Item Mitigation Plan Likelihood Impact Risk Index ENVIRON- 08 ENVIRON- 09 ENVIRON- 10 ENVIRON- 11 ENVIRON- 12 Extreme Windy Conditions, Heavy Rain, or Snow High Winds Mild Rain Cloudy Day Mild Wind Launch will be scrubbed and rescheduled. Parachute deployment altitude will be lowered to decrease drift distance. In the case of test launch, smaller motor may be used to reduce overall altitude. The rockets exterior, to avoid wetting of the ignition strips, covers the ignition system. The exterior structure of the rocket is durable enough to withstand light rain. GPS recovery system will help locate rocket. Launch may be delayed or a smaller motor may be chosen to reduce chance of loss of launch vehicle. The angular orientation of the rocket will be adjusted according to launch day wind conditions. Table 16: Failure Modes for Environmental Concerns 3 3 9 4 2 8 3 1 3 3 1 3 5 1 5 3.6 Payload Integration The entire electronics and payload system was integrated into the launch vehicle in a way, such that minimal displacement was achieved under vibration and acceleration loads. This minimal displacement was achieved by integrating all electronic and payload systems into common coupler housing. The electronics, gearbox unit, and flight computers are attached to various structural plates that are attached to threaded stringers that attach to the coupler longitudinally. The electronic payload coupler stricture was epoxied to centering rings that were then epoxied to the inside of the rocket body. The outer coupler housing the electronics and payload can be seen in
Figure 8. The inner section of the payload housing system can be seen next to the corresponding outer picture in Figure 9. Note the threaded stringers that extrude longitudinally. The green box in the CAD represents the electronic unit where the grey box denotes the mechanical gear box unit. The mechanical integrity and functionality of this integration system was verified in our successful USLI rocket of 2010-2011. This integration scheme was also validated in our most recent full scale test launch. Figure 8: Payload Outer Coupler Housing Figure 9: Payload Inner Housing Gearbox Design and Integration The gearbox design and integration is a critical component of our payloads functionality. The function of the gearbox unit is to take as input rotational motion from the servo and translate that longitudinal rotational motion to lateral rotational motion about the two shafts. The two shafts, that are extruded out laterally from the rocket body are JB welded to circular aluminum plates that act as our drag flaps. The dimensions of the gearbox were chosen such that minimal clearance was Figure 11: CAD Gearbox Figure 10: Actual Gearbox
allowed to ensure a structurally sound system while at the same time allowing play for the shafts and gears to rotate. The CAD drawing and actual fabricated gear box can be seen in Figures 10 and 11 above. The gear box unit is the most mechanically sensitive component of the entire rocket system, therefore a large amount of effort and thought was put into its integrity and integration. This integration scheme can be seen in Figure 12 below which shows the gearbox and servo units attached to the threaded stringers. Note how the servo is connected to the arduino controller system for servo calibration and testing prior to our test launch. After testing was complete the gearbox servo system was integrated into the electronic payload coupler structure and then into the rocket body. Figure 12: Gearbox Integration The bottom of the servo is adhesively attached to the bottom bulkhead of the payload electronics coupler. It is also attached to the threaded stringers via nuts. The shafts extrude laterally from the gearbox unit out through an aluminum housing ring and into the flaps that are constrained within the outer can couplers. This housing scheme can be seen in Figure 13. This integration scheme was validated during our most recent full scale test launch. After our launch we extracted the
payload electronics system and did electronic continuity and mechanical clearance and integration checks to validate that our system was still operational. Figure 13: Gearbox Housing
4. Experiment Criteria 4.1 Experiment Concept Our experiment is creative and original in the sense that it provides model rockets a way to attain a goal altitude with more confidence and less prior knowledge of flight conditions, motor inconsistencies, and thermal behavior. The ability to reject apogee amplifying an apogee hindering flight disturbances while having a minimal effect on the rockets moment of inertia or rotational rate allows a much more stable predictable flight. This experimentation also introduces a means to calculate vehicle state in flight using inertial measurement to update each other, and a dynamic target to employ a gain scheduling effect. Both of these concepts could aid in missile guidance error correction given a certain amount of a priori knowledge about the flight path in a way that may be more effective than direct gain scheduling. While a seeker-guidance system is not being employed here in 6 DOF, a simple 1 DOF implementation may prove as a proof of concept for a more robust model. 4.2 Science Value 4.2.1 Science Experiment Objectives Our payload will employ a Proportional Integral Derivative (PID) controller which will command a butterfly valve drag system designed to induce pressure drag as a means of regulating vehicle altitude. The most significant objective of our controller is to achieve exactly one mile at apogee, with a secondary goal of remaining stable during the flight. In order to insure both apogee-amplifying and apogee-depreciating perturbations are recoverable the controller will drive the rocket to follow the mean energy path. The controller should be robust enough to recover altitude goals over various launch environment conditions expected during operation in testing and in competition. In order to compensate for a number of launch day conditions, the controller should not rely on any pre-calculated values (i.e. air density, drag coefficients), instead calculating these values during the flight whenever possible. 4.2.2 Mission Success Criteria The control system must demonstrate sufficient control to overcome reasonable atmospheric and/or manufacturing perturbations encountered during flight and recover within 2% of the goal altitude. In addition, the rocket must remain stable throughout all controller commands. 4.2.3 Logic, Approach, and Investigation In order to recognize payload mission goals, a discrete time control system model was developed, and refined through design iteration until measureable outcomes were consistently and successfully simulated with high system fidelity. Below is a figure of the system model as designed in Simulink, with a brief overview as to each mechanisms purpose in achieving mission goals. A more robust discussion of each mechanism can also be found below.
Dynamic Apogee Rectifying Targeting DART Control System Dynamic Target: Used to aid in assuring the mean energy path solution is followed Restrained Controller: Proportional Integral Derivative (PID) derived controller with physical limits Physics Plant: Simulation of vehicle-environment interaction given controller commands Instrument Uncertainty: Propagation of instrument uncertainty into system values Alt. Projection: Projection of rocket apogee altitude with same physics plant model Variables Referred to in Discussion of DART Controller Altitude Above Ground Level (AGL) x AGL altitude when control system is activated x! Target apogee altitude (1 mile) x!"# Command altitude x!"# Altitude Above Sea Level (ASL) x!""#$% Initial Undershoot D! Velocity with respect to altitude x Velocity with respect to altitude when control system is activated x! Acceleration with respect to altitude x Butterfly valve angle of attack (0 corresponds to min drag, configuration) φ Angle of attack α Reynolds number with respect to natural diameter Re!"#! Total vehicle mass m Acceleration of gravity g Natural configuration coefficient of drag C!" Natural configuration cross-sectional area seen by incoming fluid (air) A! Induced coefficient of drag C!" Induced drag configuration cross-sectional area seen by incoming fluid (air) A! Density of fluid (air) using the 1976 standard atmosphere model ρ Temperature at Sea Level T! Pressure at Sea Level p! Gas Constant of Air R
Temperature Lapse Rate (1976 U.S. Standard Atmosphere) k Dynamic Targeting The dynamic target block is used as part of the system which assures that the controller targets the best energy path solution to achieve one mile altitude while retaining the ability to reject disturbances. In this system it is particularly relevant to realize that our vehicle has depreciating control ability as apogee is approached. This is due to the loss of velocity near apogee, resulting in a loss of dynamic pressure, and therefore a decrease in the maximum pressure drag the butterfly valves can produce. To compensate for this effect, we introduce a target based off of the following formula:
x!"# x, x = D! x!"# x!! sin 2π x x! x! x!"# x! x [0, x!"# ] x!"# x [x!"#, ) This nontrivial part of this equation represents a sine wave target that oscillates about a one mile command with the magnitude of oscillation decreasing as the ratio of velocity squared to initial velocity squared. The command is based off a single period of oscillation starting off with an undershoot command during early high-velocity flight due to our vehicle having the most energy to drag at that point. This behavior tends to keep our vehicle near the mean energy path solution, allowing robust disturbance rejection during the entire flight when coupled with a well-tuned PID controller. 4.2.4 Meaningfulness of Test Restrained Controller The restrained controller is used to implement modified PID control while respecting the physical limitations of our system and our environment. The PID controller itself is a
typical implementation as seen below using the forward Euler method in our discrete time domain. PID Sub- Block The command from the PID controller is fed directly into the dynamic saturation block to ensure that the controller is not attempting to actuate drag that the vehicle cannot physically perform. The lower limit for this saturation block is the effective nominal drag of our vehicle and remains constant throughout our flight. The upper limit for this saturation block is the maximum drag our system can induce at any given velocity. This command is then fed into the derivative constraint block. d/dt Sub- Block This feedback driven block assures that the controller is not trying to command that the butterfly valves cannot reach. The servo has a certain maximum rotational speed, so the controller is limited to how much it can change the command value in a particular time step. If the difference is too large, it commands the largest possible value it can in the functional direction in which the system was attempting to actuate. This block assures that our controller does not damage system components, or give commands that the controller does not have enough time to perform. This command is then fed into the perturbation block. Instrument Perturbations Sub- Block
To more accurately model a real physical system, Gaussian white noise is introduced into the drag command at a magnitude of ±5% of the commanded value. This uncertainty not only accounts for inaccuracies in realizing commanded values of drag, but also accounts for aerodynamic flutter possibly present in early phases of our flight. The adjusted drag command output is then considered against one final check. If the system detects that the vehicle is falling apogee separation is assumed to have occurred and drag is modified to represent that for simulation purposes. Physics Plant The physics plant is responsible for simulating our flight path given an effect drag coefficient, and initial conditions. Our physical model is built as follows: Physical Model Assumptions: - α 0 - Re!"!! sufficent for quadradic drag - m, g, C!", A! const - C!"## =!!! C!"A! + C!" A! - A! = A! C!" ) - C!" = C!" (x!, φ)
Governing Equation of Motion!! : D! + D! + mg = mx x = ρ x + x!""#$% x! C 2m!" A! + C!" A! x = ρ x + x!""#$% x! C!"## + g + g Here we recover a second-order nonlinear ordinary differential equation that governs the physics of our rockets altitude. This differential is evaluated numerically, and comprises the physics of our simulation model. Density is evaluated using the following sub-block, based on the 1976 U.S. Standard Atmosphere model: Standard Atmosphere (1976) Sub- Block The above block is based on the following formula where altitude, x, is measured in meters above sea level. ρ x = R T! kx T! p!!!" T! kx Further iteration may include gravity turning effects and some relaxation of the small angle of attack assumption to develop a more robust system model. Which assumptions to relax will be a product of field testing.
Instrument Uncertainty Sub Block To more accurately model our system, uncertainty in measurements and how those propagate throughout the system must be considered. In this block, we consider Gaussian white noise on the uncertainty level of our altimeter. This uncertainty is amplified at high velocity, due to the dynamics of our system resulting in a large change in altitude between every requested altitude value. However, as our vehicle slows down and we receive more altitude measurements with respect to our velocity, we can be more certain that the instrument is behaving as we would expect. Apogee Calculation As our vehicle ascends, we need to have a method to dynamically calculate our drag, and project apogee based on that calculation. Our drag must be calculated on the fly since we cannot predict variations of air density and slight miss-alignments of the butterfly valves which may vary from day to day. The following block uses only information that will be available to our controller, and calculates its current drag.
Induced Drag Coefficient to Effective Drag Coefficient Sub- Block Revising our physics plant, we will use our governing equation of motion (repeated here for convenience) to help describe this mapping of physical properties to effective drag coefficient. x = ρ x + x!""#$% x! C!"## + g Since we are now interested in an analytic answer to the differential equation governing vehicle altitude that can be executed with reasonable computational effort, we will reduce the second order differential s complexity by letting density be constant. Specifically we quantify average density as the mean density between one mile altitude and our current position. This allows us to express our physics equations as x g C!"## = x! ρ x!"# + ρ x 2 Altitude Projection Sub- Block
We now turn our attention to solving the constant-density governing equation of motion. The equation is written (noting that we have now also absorbed density into coefficient of drag): x = x! C!"## + g We know our current speed and velocity from instrumentation, hence using those two values as boundary conditions, we can constraint the governing equation and attain an analytic solution for position. x t = x! + ln g cos t C!"## g + x! C!"## sin t C!"## g! g 2C!"## C!"## x!! + g x(t) = x! C!"## x! + C!"## g cot t C!"## g x t = 0 t!"#$%% = cot!! x! g C!"## g C!"## g x t!"#$%% = x! + ln x!! C!"## g 2C!"## + 1 Note that in these equations the subscript c refers to the current state of the rocket at the time of apogee calculation. This calculation closes our control system feedback loop. 4.2.5 Relevance of Expected Data After the rocket has flown, the number that truly matters is the apogee altitude. Since we are reading the altitude off the competition altimeter, there should be no appreciable error. This is due to the fact that we are trying to achieve an altitude where we are not deducted any points by the competition, not an actual mile in altitude. Since the competition altimeter is used for the official scoring, it is inconsequential what the error is between the measured and actual altitude. 4.2.6 Experimental Process Procedures Durring flight, the control algorithem described above will run on an arduino microprocessor. Below is a skeletal architecture of the basic control loop as implemented in the C++ Arduino architecture. The Arduino inherently is set up to loop after initialization, and much of the variable communication is set up through a custom
controller class interface which has the ability to generate PID commands. Every iteration of control the vehicle conditions as reported by the hardware is checked against constants to assure that the vehicle is in the regime in which we wish to actuate control. After this a PID command is generated, checked against constraints, and fed as a PWM signal to the servo to generate a command. This loop will be repeated at our minimum driving frequency which will be governed primarily by the pulse timing required to keep the servo active. This 40Hz signal will require that redundant commands are sent to our servo, due to the 20Hz sampling rate of our altimeter. Hence, in addition to proving peripheral data, accelerometer measurements may be used for fine adjustment between altimeter samples to help guide altitude control.
4.3 Experimental Design 4.3.1 Design and Construction of Experiment The design and construction of the experiment were described in Section 3.1. 4.3.2 Repeatability of Measurement The precision and repeatability were discussed in Section 3.1.4. 4.3.3 Flight Performance Predictions The flight performance predictions were covered in Section 3.3.1. 4.3.4 Workmanship Although the bulk of construction and design for this payload is not sensitive to manufacturing/design flaws, a select few areas will require a high degree of attention in order to function properly. The counter-rotating differential is one such area. The gears will be purchased, however the gearbox and bearing supports must be constructed. Since these parts will contain rotating gears, it is imperative that they are made with a high degree of precision to prevent gear mashing or skipping. If the gear box locks up the rocket will not be able to alter its drag. Quality craftsmanship will also be required in the bearing support, since a miss-alignment between the bearing location and the location where the axels leave the gearbox may cause unnecessary forces on the gears. In addition to manufacturing requirements, great workmanship is necessary in both the electronics, and the programming on board the chips. The solder work on the PCB board must be aerospace quality, due to the high g-forces encountered during launch and landing. The controller code must be ready for a number of different failures and malfunctions, and accurately guide the rocket to its proper apogee with any number of perturbations. 4.3.5 Test and Verification A launch was scheduled to test the control system on Sunday, March 18 th, however the microprocessor failed shortly before launch. We believe this failure occurred due to mishandling when removing the electronics board from the avionics bay. Steps are being taken to insure the electronics board will not be damaged by the wall of the avionics bay next flight. Unfortunately, this made the test flight run without a control algorithm, with the valves locked in their fully open position. 4.4 Verification 4.4.1 Requirement Satisfaction and Verification 1) The payload must be able to both increase and decrease drag on the entire vehicle during flight. This requirement is satisfied by the high level design of the drag mechanism, which employs two counter-rotating butterfly valves located in opposite can tubes on either
side of the airframe. These valves are controlled by a single digital servo that can rotate through a maximum of 90 degrees away from either side of its neutral position. Since these valves can both decrease and increase airflow through the cans during flight, they can directly induce or reduce stagnation pressure above the valve, thus varying the drag force on the entire vehicle. Verification of this requirement is ongoing. During static testing on the ground, the servo has been successfully programmed to constantly sweep through 90 degrees, verifying the ability of the payload to vary airflow through the cans. Because of the failure of the payload to actuate the valves during the last test flight, as of now there is no experimental test data to give quantitative information relating valve position to drag induced on the vehicle. Data obtained from a compressible fluid dynamics model of this system, included in the Critical Design Review report, qualitatively concluded that the drag induced by a closed valve was significantly higher than the drag induced by an open valve, verifying the ability of the payload to increase and decrease drag. 2) The payload must be able to induce enough drag after motor burnout to pull the vehicle down to exactly a mile in altitude. This requirement is satisfied by the design of the drag mechanism, specifically the diameter of the cans and valves. In order to ensure that the payload is able to knock off at least 300 feet off an unaltered altitude, the diameter of the can/valves is about 2.5. Verification of this requirement was performed by estimating the minimum altitude achievable by the rocket, which corresponds to the valves being closed (perpendicular to flight path) for the maximum amount of time possible. Using data from our full-scale test launch, this minimum altitude is about 5040 feet, which is significantly lower than our target altitude of 5280 feet AGL. Please see Appendix A for a detailed look at these calculations. The data used to calculate this minimum altitude was taken from a flight that reached 5375 feet at apogee; even if the rocket is slimmed down a little bit to give a bigger cushion above a mile, the minimum altitude will clearly stay significantly below a mile. 3) The payload must not induce any instability in the vehicle during flight. This requirement is satisfied by the overall design of the drag mechanism, specifically the counter-rotating motion of the valves. Utilizing a 1 input / 2 output gearbox in the payload (using 3 separate gears), the payload rotates both valves in opposite directions by turning a single input shaft. The purpose of counter-rotating valves is to minimize any forces that might act on an axis other than the z-axis during flight. If the valves rotated the same direction, then upon actuation there would clearly be a force in the x or y direction that would try to tip the rocket (increase angle of attack). Verification of this requirement has been partially completed. A compressible flow fluid dynamics model of the vehicle showed that the net force on the rocket in the xy-plane is essentially zero under the full range of valve configurations. In more detail, although airflow interacting with a partially closed valve does create a significant force in the xyplane (call it the y direction), the airflow that is redirected horizontally as a result of this
interaction creates a nearly equal force against the can wall in exactly the opposite direction (the y direction in this example). These forces effectively cancel each other out for a net force of zero in the xy-plane. This neutralizes any concerns of the vehicle spinning on its z-axis during valve actuation. The most recent full-scale test launch verified that the valves do not create any significant instability in the open position, although it is still yet to be experimentally determined what effect actuation of the valves during flight may have on stability. 4.5 Safety and Environment MREA s acting safety officer is Kayla Wizinsky. As mentioned in Section 3.1.7 she is responsible for the safety involving the vehicle. Since we do not have a payload and are doing an experiment instead, Kayla is the officer. All hazards that could go wrong with the experiment are mentioned in Section 3.1.7 under the vehicle safety. Since there is no payload and this is an experiment there are no additional risks. The risk chart can be found in Section 3.1.7.
5.0 Launch Operations Procedures Project Wolverine will be launched from a one inch squared rail with length 72 inches. The rocket will be attached to the rail by two buttons with a 0.5 inch diameter. The blue tubes will be epoxied around the buttons to ensure strength and integrity during the launch. The following procedures (Table 17) will be followed during all launches of Project Wolverine. Operations Pre-Launch Operations Launch Pad Operations Recovery Preparation The parachute will be folded to prevent overlapping of any cords or material and dog barf will be placed between the parachute and the ejection charge N/A Motor Preparation Igniter Installation Setup on Launcher Launch Procedure to prevent burning. The proper motor delay will be calculated and the black powder charge will be shaved down for this timing. The motor will then be inserted into a casing and placed in the motor mount; the motor is held in the mount by screwing washers and a nut onto a rod which is connected to the motor mount. The igniters will be left packaged until the rocket is taken out to the launch pad. N/A The ejection charges will be wired onto the electronics and then dog barf will be packed into the rocket. Next the The igniter will be installed in the motor on the launch pad. The igniters will be the last component installed on the rocket to lessen the likelihood of a premature launch. The alligator clips will be touched to prevent static charge from igniting the rocket and then the clips will be placed far apart to prevent shorting from occurring. The launch rail will be pulled down until the rail buttons on the rocket can be slid onto the rail. The rail will then be vertically moved into position and aimed in the proper flight trajectory. After the rocket is placed on the launch pad the flight computer and altimeter will
Troubleshooting Post Flight Inspection parachute will be folded and inserted into the rocket and all of the rocket will be assembled. Finally, shear pins will be installed at the two separation points and then a final safety check will be completed. The flight computer will run through a simulation prior to flight to ensure signals are being sent at the proper altitudes. This will be done by connecting the different leads on the flight computer to a multimeter. The rocket will then be searched to ensure all components are attached properly. N/A Table 17: Launch Ops be armed. Then the ejection charges will be installed and a final safety check will be completed. The flight performance will be analyzed to account for mission success or failure. This will be done by a physical inspection, plugging the flight computer into the computer to get values, and listening to the beeps which signify altitude on the altimeter. The rocket will be inspected as if it were going to be launched a second time. If the rocket is deemed ready to launch again then the flight will be considered successful. Recovery Preparations The recovery system pre-flight checks are performed before every launch. The Raven2 flight computer is hooked up to a mini-usb connection, appropriate terminals are shorted together to simulate an e-charge, and a simulated launch is performed. The e-charge current for apogee and main chute e-charges is measured by a multimeter to ensure the Raven2 is programmed properly. The e-charge size for each separation is calculated using a matlab code based on the Perfect Gas Law. The e-charge canisters are then filled with the appropriate amounts of 4F black powder and wadding to ensure the black powder is in constant contact with the ignition wire. Motor Preparation Our rocket motor uses a reload system, requiring the solid motor to be assembled by hand. Each motor contains assembly instructions specifically laying out this procedure and that will be the basis for motor assembly. After the motor is assembled into the motor casing, the casing is then installed into the rocketʼs motor mount, and the aero pack retention system cap is screwed on completing the Motor preparation. Igniter Installation The motor igniter is not placed inside the motor until after the rocket is on the launch pad with the avionics off. The igniter is inserted into the motor until the top of the igniter
meets the top of the motor. A rubber stopper is inserted into the lower end of the motor to hold the igniter into place. Lastly each lead of the igniter is hooked up to the alligator clips leading to the launch control console. Setup on Launcher The rocket is loaded onto the launcher by inserting the launch lugs into the launch rail. The rail is then rotated into the appropriate launch position. Next, the motor igniter is installed into the motor and connected to the launch control console. Lastly, the avionics switches are turned on and the setup is complete. Troubleshooting Troubleshooting avionics problems will be done by voltage, resistance, and continuity checks on the recovery system. Spare batteries and any spare avionics components will be on hand to be used as replacement parts if needed. Motor troubleshooting from a failed motor ignition will be done in compliance with the range safety officer. Postflight Inspection After the flight, the rocket will be first inspected for external physical damage, such as motor retention system, fins, or fuselage. The parachutes and shock cords are then inspected for damage from the e-charges or a violent deployment. The avionics bay is then inspected for burnt wiring, shorts, or open circuits. The flight altitude is then recorded and the post-flight inspection is complete.
6.0 Activity Plan 6.1 Budget Plan MREA has purchased all the hardware necessary to construct a vehicle for two test flights and expects to be able to reuse all of the hardware for any subsequent flight tests and the competition flight without needing to make any major purchases with the exception of engines. MREA expects to spend approximately $6,000 on the entire USLI project, Actual flight hardware has cost MREA about $2,800, which puts MREA well under the competition rule of $5,000 of flight hardware on the launch pad. Table 18 and Figure 14 details all of the expenditures MREA has incurred and expects to incur throughout the remainder of the USLI project. MREA has taken travel and the outreach project into account in all budget calculations. Subsystem/Category Total Cost (USD) Structural Components 722.58 Controls Structural Components 289.43 Avionics Components 573.47 Propulsion Components 1216.91 Travel Expenses 2201.75 Outreach Project 585.61 Shipping Costs 387.13 TOTAL 5976.88 Table 18: Project Cost Breakdown by Subsystem and Category
Outreach Project 10% Shipping Costs 6% Structural Components 12% Controls Structural Components 5% Avionics Components 10% Travel Expenses 37% Propulsion Components 20% Figure 14: Visual Representation of Project Costs In order to fund these efforts, MREA has recently received grants from two student government organizations at the University of Michigan: the Central Student Government and the University of Michigan Engineering Council. These grants total $1,700 and raise MREA s overall funding to approximately $7,200. MREA expects to have plenty of cushion to make additional purchases if any problems arise. Table 19 details MREA s funding sources. Income Source Category Amount Starting Balance Starting 1956.67 UM Aerospace Engineering Dept. Grant Grants 2500.00 Raytheon Company Grant Grants 1000.00 UM Central Student Government Grant Grants 700.00 University of Michigan Engineering Council Grant Grants 1000.00 TOTAL 7156.67 Table 19: Project Income Sources X.2 Project Budget Details Included in the following tables are details on the individual components that have been purchased for the Structural, Controls Structural, Avionics, and Propulsion subsystems. Structural Components Supplier Unit Cost (USD) Quantity Total Cost (USD) Upper Body Tube Apogee Rocketry 56.95 1 56.95
Outer Fuselage/Side Cans Apogee Rocketry 26.95 5 134.75 Can Coupler Apogee Rocketry 9.25 4 37.00 Upper/Lower Body Tubes Apogee Rocketry 56.95 2 113.90 Misc. Lab Supplies Ace Hardware 93.05 N/A 93.05 Blue Tube Coupler 2.56" Apogee Rocketry 9.25 2 18.50 Blue Tube Coupler 5.5" Apogee Rocketry 18.95 1 18.95 Main Parachute LOC Precision 189.00 1 189.00 Plywood Home Depot $34.97 N/A 34.97 Home Depot Supplies Home Depot 25.51 N/A 25.51 Radioshack Supplies Circuit City 31.17 N/A 31.17 TOTAL 722.58 Table 20: Structural Subsystem Component Costs Controls Structural Components Supplier Unit Cost (USD) Quantity Total Cost (USD) Flat Washer McMaster-Carr 3.37 1 3.37 Lock Washer McMaster-Carr 4.03 1 4.03 Screw McMaster-Carr 9.80 1 9.80 Rods McMaster-Carr 6.21 1 6.21 Hex Locknuts McMaster-Carr 5.01 1 5.01 Hex Nuts McMaster-Carr 4.49 1 4.49 Drag Flap Thrust Plate (Ring) McMaster-Carr 51.30 1 51.30 Avionics Bay Apogee Rocketry 38.95 1 38.95 HS-5645MG Digital Torque ServoCity 46.99 1 46.99 Voltage Regulator Resistor Mouser Electronics 0.19 12 2.28 Diode Mouser Electronics 0.34 6 2.04 Gears McMaster Carr 36.48 3 109.44 Eyebolts McMaster-Carr 2.76 2 5.52 TOTAL 289.43 Table 21: Controls Structural Subsystem Component Cost Avionics Components Supplier Unit Cost (USD) Quantity Total Cost (USD) Drag Flap Thrust Plate (Ring) McMaster-Carr 51.30 1 51.30 732 Ohm Resistor Mouser Elec. 0.06 12 0.72 240 Ohm Resistor Mouser Elec. 0.13 12 1.56 300 Ohm Resistor Mouser Elec. 0.13 12 1.56 Capacitor Mouser Elec. 0.31 12 3.72 Capacitor Mouser Elec. 0.98 12 11.76 Arduino Connectors Mouser Elec. 2.06 4 8.24 Accelerometer Regulator Mouser Elec. 0.60 6 3.60 Arduino Voltage Regulator Mouser Elec. 1.70 6 10.20 Connectors Mouser Elec. 1.05 6 6.30 Accelerometer Sparkfun Elec. 28.95 1 28.95 LED Indicator Mouser Elec. 1.39 2 2.78 LED Indicator Mouser Elec. 0.59 2 1.18 LED Mount Mouser Elec. 0.26 4 1.04 Fixed Terminal block Mouser Elec. 0.41 4 1.64 Public Missile Fin-D-07 Public Missles 19.15 4 76.60 Key Switches McMaster-Carr 12.95 2 25.90 E-Charge Canisters Apogee Rocketry 10.00 5 50.00 E-Charge Canisters Apogee Rocketry 9.99 5 49.95 Perfect Flight Altimeter Perfect Flight 67.96 1 67.96 Voltage Regulator Resistor Mouser Electronics 0.19 12 2.28 Diode Mouser Electronics 0.34 6 2.04
Micro-SD Card Board Adrafruit Industries 15.00 1 15.00 Steel Bore Miter Gear McMaster-Carr 18.58 3 55.74 3"x6"x3" Aluminum Block McMaster-Carr 44.12 1 44.12 Servo Hub Servo City 9.99 1 9.99 Coupler Hub Servo City 7.99 1 7.99 Screws Servo City 0.35 4 1.40 Drag Computer Mouser Elec. 29.95 1 29.95 TOTAL 573.47 Table 22: Avionics Subsystem Component Costs Propulsion Components Supplier Unit Cost (USD) Quantity Total Cost (USD) Motor Mount Tube Public Missiles 16.50 1 16.50 Motor Mount Centering Rings LOC Precision 7.00 4 28.00 Motor Retention Apogee Rocketry 52.00 1 52.00 Motor Thrust Plate McMaster-Carr 12.47 1 12.47 5 Grain Motor Retention Giant Leap Rocketry 203.45 1 203.45 Pro 75 Case Spacer Giant Leap Rocketry 24.95 1 24.95 Flight Test Motor #1 Loki 160.00 1 160.00 Flight Test Motor #2 Cesaroni 239.54 1 239.54 Flight Test Motor #3 Cesaroni 240.00 1 240.00 USLI Launch Motor Cesaroni 240.00 1 240.00 TOTAL 1216.91 Table 23: Propulsion Subsystem Component Costs Outreach Project Components Supplier Unit Cost (USD) Quantity Total Cost (USD) Nose Cone LOC Precision 14.65 5 73.25 Payload Body LOC Precision 6.25 5 31.25 Tube Coupler LOC Precision 2.85 5 14.25 Main Body LOC Precision 6.25 5 31.25 Parachute 36" LOC Precision 20.95 5 104.75 Motor Mount LOC Precision 4.75 5 23.75 Centering ring(2) LOC Precision 5.80 5 29.00 Bulkhead assembly LOC Precision 2.85 5 14.25 Rail Buttons LOC Precision 5.25 5 26.25 Shock Chord LOC Precision 6.25 3 18.75 G-10 Fiberglass Giant Leap Rocketry 17.99 5 89.95 F-Class Motors Apogee 42.97 3 128.91 TOTAL 585.61 Table 24: Outreach Project Costs Travel Expense Unit Cost (USD) Quantity Total Cost (USD) Car Rental 250.00 2 vehicles 500.00 Gas 1 0.25 1950 miles 487.50 Hotel 2 242.85 5 nights 1214.25 TOTAL 2201.75 Table 25: Travel Cost Estimations 1 Gas estimates are calculated based on $4.50 per gallon and 20 miles per gallon fuel efficiency. 2 Hotel has been booked and cost estimates are made based on on booking prices.
6.2 Timeline The below timeline is the schedule that MREA has been using to meet all of its project goals. It includes all tasks completed since CDR, as well as all remaining tasks. Thus far, MREA has been successful at meeting its project goals. Most importantly, MREA has already completed two full-scale subsystems tests (without the control system active) and has proven launch vehicle safety as well as the ability to fly a successful flight. MREA has at least one more test flight scheduled (with the capability to do up to two more), in order to tweak and refine the controls system. At this point, MREA does not foresee any problem executing a successful launch in Huntsville on April 21. January 2012 4 22 Critical Design Review Development 7 15 Preliminary Subsystem Construction and Ground Tests 21 Test Flight #1 (No active control) 23 CDR Completed 23 Feb. 10 Construction and Preliminary Integration February 2012 1 10 CDR Presentation 1 15 Systems Ground Tests 16 29 Construction of Final Vehicle Subsystems March 2012 1 15 Final Vehicle Assembly and Integration 1 26 Flight Readiness Review Development 18 Test Flight #2 (No active control) 18 25 Test Flight Analysis 26 FRR Completed 26 31 Final Flight Test Preparation April 2012 1 Flight Test #3 (Full active control) 3 FRR Presentation 2 14 Final Systems Adjustments 14 Flight Test #4 (if needed; full active control) 18 Travel to Huntsville 19 20 Flight Hardware and Safety Check 21 Launch Day 22 Travel Home 23 30 Final Exams May 2011: 1 6 Post-Launch Assessment Review Development 1 6 Launch Evaluation and Data Analysis 7 PLAR Completed 18 Announcement of USLI Competition Results
The below Gantt chart in Figure 15 gives a visual representation of the tasks that MREA has completed since CDR (in blue) and the tasks that MREA has remaining for the USLI competition. Test Flight #1 CDR Completed Preliminary System Integration CDR Presentation Final Subsystems Construction Final Assembly & Integration FRR Development Systems Ground Tests Test Flight #2 Test Flight Analysis FRR Completed Flight Test Preparation Test Flight #3 FRR Presentation Final Systems Adjustments Test Flight #4 (TBD) Final USLI Preparations Travel to Huntsville Flight Hardware & Safety Check Launch Day Travel Home Final Exams PLAR Development Launch Data Evaluation PLAR Completed Announcement of USLI Results 1/13 2/2 2/22 3/13 4/2 4/22 5/12 Figure 15: Gantt Chart of Tasks Since CDR (Completed in Blue, Remaining in Red)
6.3 Educational Engagement The Michigan Rocket Engineering Association has completed its educational engagement requirement as of Friday February 17, 2012. The project took its first successful run this February as the members of MREA visited Pinckney Community High School (PCHS) for a three week seminar. MREA aimed to teach a PCHS physics class about the basics of rocketry through a small class competition organized by MREA members. Members visited Mr. Joel Craig s physics class everyday for three weeks. The seminar s timeline and competition rules can be found below. Timeline Jan 30-31: Lesson on thrust, drag, and stability Feb 1-3: Rocksim tutorials and work days. Final design selection. Feb 6-15: Build Days Feb 16: Payload Integration Feb 17: Launch Day Rules Each team was given set materials that they could alter. The only things that were alterable were the lengths of the body tubes, the lengths of the payloads and the geometry of the fins. Changing these aspects in Rocksim gave them a chance to test launch their rocket within the program trying to reach as close as 750 feet as possible in altitude. The final goals regard a payload section that the students were required to build that would safely house a raw egg during the flight. The team that got the closest to 750 feet and kept their egg safe was declared the winner of the competition. As an introduction to the project, members taught students about basic rocketry concepts like thrust, drag, and stability, as well as a basic tutorial of Rocksim. The students were given a template rocket in which they were only allowed to change the lengths of their body tubes and the shape of their fins. Given this template and a the set of competition rules, the students got into groups and designed rockets that would get them as close to their goals as possible. MREA members reviewed each design to make sure each was stable and safe for flight. Once approved, members cut all of the materials to specifications given by the students designs. MREA chose not to make the students, who may not all have had woodshop training, responsible for cutting any of the rocket parts for safety reasons. MREA members spent two weeks helping the students build their rockets, and on Friday February 17, PCHS had its first rocket launch on its football field. The science classes at PCHS were all invited to attend the launch and it was expected that about 200 students watched the event from on the field and from inside classrooms.
MREA has established a good relationship with PCHS and has been invited back to teach the seminar again next year. MREA also hopes to mentor PCHS s first TARC team next year as our outreach project grows. The goal is to establish more of an interest in rocketry at PCHS and hopefully around the school district of Pinckney. "It's a lot of pressure when you're trying to launch these things in front of a lot of people," Craig said. "It was a great turnout, and you can see how much the students really enjoy doing these kinds of projects." - Joel Craig (Physics Teacher at PCHS) Article by Daily Press & Argus reporter Frank Konkel Figure 16: MREA members (from left to right) inspect the connections on the launch pad. Photo by ALAN WARD / DAILY PRESS & ARGUS Figure 17: "Team America retrieves their rocket after launching. Photo by ALAN WARD / DAILY PRESS & ARGUS
Figure 18: Team "Flash" shows off their rocket before launching. Photo by ALAN WARD / DAILY PRESS & ARGUS Figure 19: MREA members at the Launch on Feb 17. Photo by ALAN WARD / DAILY PRESS & ARGUS For news clips and pictures please visit: http://www.livingstondaily.com/article/20120221/news01/202210312/pinckney- Michigan-partnership-makes-learning-blast-video-photo-gallery-?odyssey=tab%7Ctopnews%7Ctext%7CFrontpage