2008-2009 University Student Launch Initiative

FLIGHT READINESS REVIEW (FRR) Panther II Heavy:

Measurement and CFD Prediction of Liquid Slosh Behavior During Model Rocket Flight

Submitted by: The Florida Institute of Technology

Melbourne, Florida, 32901

March 18, 2009

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Executive Summary Panther II Heavy is the Florida Institute of Technology’s rocket in NASA’s USLI competition. The Florida Institute of Technology is located in Melbourne, Florida. The Mentors for the project are Dr. Daniel R. Kirk, Assistant Professor – Department of Mechanical and Aerospace Engineering. Other participating faculty include Dr. Hector Gutierrez, Associate Professor – Department of Mechanical and Aerospace Engineering, and Mr. H. Greg Peebles III, P.E., Director – University Safety Office. Four inch fiberglass tubing was selected as our airframe material due to its low cost, light weight, and high strength. Our two preceding rockets, Panther I and Panther II, both used a four inch fiberglass airframe which allows us to pull from previous experience. Also, using a four inch diameter rocket still allows plenty of room for our scientific payload while keeping the size of the motor required to achieve our desired altitude an L930. We chose the L930 motor from Loki Research for a variety of reasons. We have worked with Loki Research on developing motors in the past, so we are confident of their work. The L930 motor is fairly inexpensive, allowing for us to test fire several motors before our test flights and have extra motors for the contest. The motor was selected once the preliminary vehicle weight estimation was completed. The Loki Research L930 was selected because of its simulated capability of lifting the rocket safely to the target altitude of one mile, while maintaining relatively low acceleration throughout the burn. The recovery system was designed to accommodate an initial high speed deployment because of payload needs, and a secondary deployment to bring the rocket to an appropriate landing velocity. The RocketMan Enterprises Pro-EXP 4ft. parachute will be used for the first stage of recovery as it is specifically designed for subsonic, high speed deployment. A RocketMan Enterprises 12ft. standard parachute will be used as the second stage, main deployment that will decelerate the rocket to a safe descent. The payload experiment for the Panther II Heavy rocket is the observation of slosh during high- and low-gravity flight regimes. This task will be completed by recording infrared video of a small tank partially filled with water during the flight. The images from the initial acceleration as well as the free fall period at apogee will be compared to a CFD simulation of the rocket using the 6 DOF data collected during the rocket flight.

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Changes Made Since CDR The FRR is the final summary of the completed or near-completed components of Panther II Heavy. All testing of all components has been completed or will be in the next two weeks. A successful launch with a dummy payload occurred on March 14, 2009. All changes from the CDR are listed below: Airframe:

• Fins reduced from 3/8” thickness to 1/8” • Drogue and main parachute airframe locations switched

Motor:

Several test fires of motor using Florida Techs 6DOF Thrust Stand Recovery:

• Testing: o Complete Avionics Testing o Completed all recovery calculations

Time in free fall Velocity/Altitude of deployments Deployment shock of drogue Shear pin calculations Black Powder amounts Shock cord length for each section

o Ejection charge test Live Black Powder charges All components of recovery system installed for test

o Flight Test All aspects of system tested completely in a full scale flight Recovery system was successful Initial flight data analysis indicates a verification of calculations

Payload:

Dummy payload completed o Equivalent center of mass, length o Simulated slosh tank used

Payload layout design completed o Assembly instructions completed

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Table of Contents

Executive Summary ...................................................................................................... 2

Airframe Selection ......................................................................................................... 5

Motor Selection ............................................................................................................. 7

Recovery Subsystem .................................................................................................... 8

Payload Integration ..................................................................................................... 12

USLI FRR Hazard Analysis ......................................................................................... 13

Payload Criteria ........................................................................................................... 18 Selection, Design, and Verification of Payload Experiment ....................................... 18 Payload Concept Features and Definition ................................................................. 21 Science Value ............................................................................................................ 23 Assembly ................................................................................................................... 25 Failure Modes of Payload .......................................................................................... 29

Safety & Mission Assurance ...................................................................................... 30 Regulatory Compliance ............................................................................................. 30

Rocket Simulation ....................................................................................................... 32

Project Plan ................................................................................................................. 40

Appendix A: Background on Fluid Slosh Research at Florida Tech ...................... 41

Appendix B: Senior Design Safety Plan Requirement ............................................ 58

Appendix C: MSDS Sheets ........................................................................................ 60

Appendix D: Payload Specifications ........................................................................ 77

Appendix E: Payload Calculations ........................................................................... 82

Appendix F: Rocket Performance Calculations....................................................... 86

Appendix G: Launch Test Directive .......................................................................... 92

Appendix H: Parts List ............................................................................................... 99

Appendix I: References ........................................................................................... 101

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Airframe Selection

The airframe tubing selected is Hawk Mountain Enterprises 3.91"x4.03" fiberglass tubing. The high strength and low weight as well as ease of attaching other parts to this tubing makes it an ideal airframe. One main advantage of this tubing it is ability to resist buckling. A four inch diameter rocket that is almost 13 feet tall has a very large likelihood of buckling during the boosting stage of flight. Euler's equation is used to calculate the critical loading of the rocket during ascent. The K value was chosen to be two as this is the value which most resembles the loading of a rocket during ascent, that is one end fixed and the other end free to move. The K value ranges from one to two depending on the loading conditions of a column under compression.

Assuming a static load of 25 pounds, an acceleration of 10g, and a maximum

velocity of 550ft/s (375mph), total load on the rocket would be 280.8lbs. This is assuming a massless rocket with all of the weight at the nose and accounting for the dynamic pressure.

The rocket is completed. Since we decided to go with much smaller fins due to an

error in our thrust stand’s off axis sensors, we will reconstruct the booster using thinner 0.125 inch thick G10 for the fins and the same 0.187 inch thick G10 for the centering rings. To keep our rocket stable, we calculated the center of pressure (CP) using the Barrowman Equations.

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Figure 1: Detailed Drawing of Rocket

Figure 1 shows a complete schematic of the airframe for our rocket. The payload is in the most forward position in the airframe because it will be the heaviest part and will help keep the CP/CG relation in check. After consulting with Ky Michaelson of Rocketman Enterprises, we decided on using a two stage recovery system. The main parachute section will be held in place during the first stage of recovery. The electronics bay in the recovery section of the airframe will be held in place using an appropriate number and size of radial bolts. The actual electronics bay will be constructed of Hawk Mountain Enterprises coupler tubing and a two inch spacer of body tubing. This tubing will then be slid into the airframe and bolted in place creating two chambers for the drogue and main parachute. The large bulkhead at the base of the payload section will be made of one half inch thick 6061-T6 aluminum. This bulkhead carries the largest load in the rocket. During ascent, it will have the weight of the payload pressing down on it, and during recovery, the weight of the payload will be pulling it apart. Using the appropriate number of radial ¼-20 machine bolts, it will be attached in the same way as the electronics bay. The main reason for the ½” thickness is to allow room for the drilling and tapping of bolts.

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Motor Selection

With the completion of the preliminary vehicle design, a maximum liftoff weight of 30 lbs. was able to be computed. With this critical rocket parameter determined, the motor was selected. The most appropriate motor that was capable of lifting the vehicle and payload to one mile was determined to be the Loki Research L930. This motor was selected for its longer burn time, initial and average thrust, and its total impulse. The longer burn time of four seconds was a desirable trait as it reduces the drag during the boost. The drag is reduced because the maximum velocity is not reached until later in the flight, and drag is directly proportional to velocity. Therefore, the lesser the amount of time at high velocity, the less the total impulse drag has on the rocket. The initial thrust of the motor reaches 250 lbs which accelerates the vehicle fast enough for it to reach a stable velocity by the time the rocket reaches end of the launch rail. The average thrust of 930N provides a steady acceleration under the 15g payload limit throughout the entire boost. The total impulse of 3587Ns will guarantee the rocket will reach the target altitude of one mile, under the expected conditions.

The motors have been ordered to allow excess time for their arrival. A total of nine motors were ordered for the entire project. Three of these motors will be tested on the 6-Degree of Freedom (DOF) test stand to gather statistically significant data for the simulation. Four motors have been designated as use in test flights, with two being with the dummy payload, and two with the real payload. The dummy payload will be used for two test flights so that the real payload will not be damaged in the event that an anomaly occurs during a test. One motor is to be used at the actual competition, leaving one motor as a backup. Before and after each motor burn, all dimensions and parameters of the motors will be determined to guarantee the greatest possible consistency between flights. Loki Research was chosen as the motor manufacturer because of their guarantee to consistency between all of the motors they build. In addition to nine propellant grains being ordered, three complete motor cases were ordered for increased efficiency between test flights as multiple motors can be assembled or cooling at once, and as backups. Six nozzles were ordered so that nozzle decay could be analyzed, and new nozzles would be available for several flights if nozzle decay is determined to be a major issue. The nozzles need to be exactly the same size and shape for each flight to have the greatest consistency in the motor burns. Because an active guidance system is not being utilized on the vehicle, the motor performance must be extremely consistent between flights so that adjustments can be effectively made to the rocket for it to reach the target altitude.

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Recovery Subsystem

The scientific experiment onboard the rocket requires the maximum allowable microgravity time during flight to be achieved. To accommodate this need, the rocket will be allowed to freefall for approximately seven seconds where the weight of the rocket will be at a minimum. This minimum will be used to simulate a microgravity condition. Based on initial simulations, the rocket will be traveling about 200ft/s at the predetermined end of the freefall period. Because allowing a rocket to freefall deviates from the typical flight profile of model and amateur rockets, a recovery system designed to handle these changes will be employed.

In order to achieve the desired recovery profile, several commercially available flight computers will be used. Flight computers are devices designed to apply current to designated outputs to fire charges at certain events during rocket flight. For the first recovery stage, each flight computer had to be able to detect inertial apogee and fire a deployment charge at a predetermined time after that point. This meant that the flight computers had to be accelerometer based for event detection and programmable for the desired time delay. The second recovery stage required the flight computers to fire the main deployment charge at an altitude of 800ft. The flight computers selected to fulfill these tasks were the G-Wiz HCX and the Ozark Aerospace ARTS 2. Both flight computers have basic data recording capabilities that will be utilized during the initial test flights, and have multiple outputs that are programmable to fire at any time during the flight. Two different manufactures were chosen to add redundancy to the system. This eliminates issues caused by design flaws because it is unlikely both units will have a same or similar flaw, therefore adding a backup system. The deployment time and altitude of both flight computers will be offset slightly to ensure the charges from each do not occur at the same time, over pressurizing the vehicle

The layout of the recovery system had to deviate from conventional two stage recovery designs in hobby and amateur rocketry. Conventional designs typically deploy a drogue parachute from the center of the vehicle, and a main parachute immediately under the nose cone. However, to accommodate the altered flight profile the drogue parachute will be deployed from the center of the vehicle immediately under the payload compartment. This allows for the drogue to be as close to the payload as possible since the payload is the highest concentrated weight in the vehicle. This helps to maximize the effectiveness of the drogue and prevents the need for shear pins or rivets to secure a recovery section break between it and the payload. The main parachute could not be placed in the conventional location at the forward end of the rocket because the payload needed to be easily accessible at that location. Instead, the main parachute will be located in a compartment aft of the drogue and avionics compartments.

The recovery system design incorporates the use of two independent stages of parachute deployments. The first recovery stage will consist of a small drogue parachute designed to slow the rocket from a ballistic decent to a controlled decent between fifty and seventy-five feet per second. The parachute selected for this task is a RocketMan Pro-EXP four foot diameter designed for high speed deployment. Ky Michaelson, an accomplished amateur rocketeer and owner of RocketMan, was contacted for his advice concerning the selection of the correct parachute for this unusual scenario. He recommended the use of the Pro-EXP series parachutes as they

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are designed for high speed, sub-sonic deployment. A four foot diameter parachute of this series was chosen to provide the desired rate of decent. For the second stage of recovery, a main parachute will be deployed slowing the rocket to 15ft/s. A 12ft. Standard RocketMan parachute was selected for this purpose as no abnormal decent characteristics will take place during second recovery stage. RocketMan parachutes were chosen because of their known reliability, strength for the high speed deployment of stage one, and their resistance to tangling and deployment failure due to their limited number of shroud lines. To reduce deployment and inflation shock of the parachutes, Giant Leap Sliders will be used. The Sliders function to slow the inflation of the parachute by actively reducing the reefing.

The parachutes will be deployed with the use of pyrotechnic charges fired with the use of electric matches. Electric matches are small devices that have two lead wires connected to the output terminals of the flight computer. A bridge wire connects the two lead wires. The bridge wire is often nichrome or tungsten which is designed to rapidly heat up when current passes through it. The heat generated is used to ignite a pyrogen that is covering the bridge wire, which in turn ignites the ejection charge. There will be redundant ejection charges for each parachute. This helps prevent ejection failure in the case of electric match failure, or incomplete combustion in one of the charges. Each deployment stage will contain two mortars housing charges. Each charge will be wired in parallel to each flight computer so that if a flight computer fails, the backup flight computer can still fire both the primary and backup charges.

To protect the main parachute during ejection charge firing, a sabot will be constructed to completely surround the parachute. Once ejected, the sabot will release the parachute allowing a complete deployment. The drogue parachute will use a simpler, more conventional design by using a Kevlar heat shield which is lighter and requires less space than a sabot. A sabot is not necessary for the drogue parachute due to the parachutes small size.

Shock cord attaching all sections of the vehicle during its recovery is necessary for the success of each flight. The minimum allowable strength of the shock cord was determined by predicting the maximum weight of the rocket and all of its components and finding the necessary strength if the deployment created an instantaneous acceleration of at least 100g‘s. This yielded the minimum allowing strength to be 3500lbs. Kevlar shock cord from Giant Leap Rocketry will be used during both recovery stages because of its resistance to heat and for its strength. The size that will be used will be one half inch tubular Kevlar with strength of over 5000 lbs. A total of 100 ft. of shock cord will be used to allow for adequate time for each section of the rocket to decelerate during the deployment of each parachute. This follows the rule of thumb to have at least six times the length of the rocket in shock cord. The shock cord will be attached to the vehicle with U-bolts at the forward and aft bulkheads of the drogue and main parachute compartments.

To minimize the space allocated for parachutes in the rocket and to increase their deployment reliability, a Florida Tech professor with significant skydiving experience was consulted. The professor demonstrated what he recommended the best ways to pack the parachutes were within the rocket. This, in addition to deployment testing, will be used to finalize the packing method of the parachutes. Once

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constructed, the sabot will be tested with the main parachute installed to ensure a clean release of the parachute.

Testing of the complete recovery system was accomplished upon the completion of the vehicle. This included ejection charge tests, avionics tests, and eventually a complete flight test.

The G-Wiz HCX and Ozark Aerospace ARTS 2 flight computers were tested for basic functionality with the computer interface. This required the installation of all necessary software and drivers, and the hardware interface for each unit. The flight computers were tested with their sensor and output test functions in their respective software packages. The accelerometer and barometer sensors were found to be responsive and the outputs were tested with simulated electric matches. These were constructed using low-wattage light bulbs wired to the flight computer outputs which provided a safe alternative to live charges.

Successful completion of the recovery system ground and flight tests relied on the ability to characterize all flight events in advance. To accomplish this, the first necessary calculation determined the deployment time after apogee. This calculation was a basic model of free-fall incorporating drag and gravitational forces acting on the rocket during its initial descent for low-speed, subsonic flight. The data produced by this spreadsheet was analyzed to determine a suitable deployment time that provided the scientific payload with enough low gravity time for data collection, as well as the being within the structural limits of the vehicle and parachutes. The selected time for freefall was 7 seconds based on these calculations. The approximate altitude of deployment is therefore 4500ft, allowing ample time for a complete drogue parachute deployment and reaching of a safe terminal velocity.

With a drogue deployment time, velocity, and altitude selected, additional recovery parameters could be determined. A critical parameter to for the recovery system to functional properly is the need to know the shock the rocket undergoes when the drogue and main parachutes are deployed. This is especially necessary for the design of the structure of the vehicle and the necessary shear pin size for the main parachute. It was determined that the maximum acceleration on the rocket during the drogue deployment would be 16 gees. All other recovery system calculations depended on this value in addition to a factor of safety.

The shock cord length for each section was determined by distributing the 100ft available between the main and drogue compartments. The majority was used in the drogue because of its high speed deployment requirement. The main parachute compartment received enough to allow for a complete deployment of the sabot, shock cord, and parachute. In both sections, the components tied to the shock cord were secured in specific locations to prevent impact between those sections while under parachute descent.

Next, the minimum shear pin sizes were determined for both the drogue and main compartments. The shear pins are used to prevent premature parachute deployments as they must reach a specified shear strength before the parachute may leave the rocket. The shear pins for the drogue were calculated based on standard high power rocketry shear pin calculations. In addition to these calculations, the drag profile of all components of the vehicle was analyzed at the expected deployment velocity. This was to ensure that the booster would not be trying to separate from the main and

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payload sections due to additional fin and base drag, over the forward sections profile drag. It was determined that the profile drag was much greater than the combined fin and base drag, causing the net force to hold the compartment closed and decreasing the necessary shear pin strength. A total of four 4-40 type nylon machine screws were needed for the drogue shear pins.

The main parachute number and size of shear pins was determined based on the drogue deployment accelerations. Because of the large negative acceleration expected to coincide with drogue deployment, much more significant shear pins were needed to prevent the large momentum of the payload section from prematurely opening the main compartment. The shear pins were therefore sized to meet the expected loading on the vehicle at the moment of drogue deployment and inflation. It was determined that six #6 nylon machine screws would be used to meet this objective.

Now that the parachute compartment and shear pin sizes were known, the amount of black powder for the ejection charges could be determined. The amount of black powder needed was calculated by determining the required design pressure inside each chamber that would successfully shear the shear pins and deploy the recovery components. The charge for the drogue deployment was calculated to require 2 grams of black power, while the main charge was determined to require 9 grams black power.

These calculations were used as the starting point for ground testing of the deployment systems. Too little charge will result in an incomplete ejection, possibly jeopardizing the safety of the flight. Too great of a charge could cause damage to the vehicle structure, shock cord, or parachutes. Therefore, it is imperative to complete ground testing in a controlled environment to prevent ejection failures. Testing of the ejection charges is accomplished using the actual electric matches and charges to be used in flight, but with a proven launch controller for safety purposes over using the automated flight computers for such tests.

The drogue test was begun with 2 grams of black powder as calculated and was placed in its respective compartment attached to the avionics bay. To seal the end opposite of the avionics, a dummy bulkhead was built to simulate the actual compartments above the drogue. It was designed so that it could be weighted as necessary to simulate all forces acting against the ejection charge. The first test with 2 grams of black powder demonstrated that it was not enough as an incomplete ejection took place. The shear pins failed as designed, however, the shock cord and parachute did not completely deploy. Additional powder was slowly added to subsequent tests until a full deployment was reached. It was found that 4 grams of black powder were necessary for this.

The main parachute was also tested in a similar fashion as the drogue. However, it was packed into its sabot to simulate it complete flight configuration. It was found that the calculated 9 grams of black powder were optimal for causing the shear pins to fail and completely deploying the main parachute and sabot assembly.

With the completion of the recovery system’s ground testing, the components were integrated into the rest of the vehicle assembly. The avionics were secured in the avionics bay. Screw switches were mounted to the body wall of the avionics bay for external access to the power of the flight computers. Arming shunts were installed to prevent the accidental firing of the charges by the flight computers until the shunts are removed moments before launch. Each flight computer, including the Perfectflite

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MAWD used for competition altitude recording, has its own power supply and power screw switch to prevent interference, and for added redundancy. Both flight computers used for parachute deployments have independent arming shunts.

To complete testing of the recovery system, a test flight was conducted. This was done to test the preparation procedures of the system, as well as its actual in flight performance to ensure it will successfully withstand the unusual nature of parachute deployments required. The recovery system was prepared and installed in the vehicle in advance of traveling to the launch site. This allowed it to be prepared in a clean, controlled environment increasing reliability and consistence between flights. The ejection charges were the only component not to be installed in the vehicle in advance for safety reasons.

At the launch site, once all other components of the vehicle were prepared, the recovery system was inspected to ensure it retained its correct installation during transportation to the launch. The shock cords were then attached to the ends of the compartments to the U-bolts on the bulkheads and the eye bolt on the motor. Ejection charges were then carefully installed as to not damage them or cause them to accidentally fire. All of the compartments were connected once the ejection charges were ready and the shear pins were placed in their respective locations.

During the recovery portion of the launch, the rocket successfully reached apogee then freefell for the programmed seven seconds where the first altimeter fired the first drogue charge. The drogue deployed completely while the second, redundant charge fired for the drogue. The main parachute followed a similar proceeding as the primary altimeter fired the first main charge at 900ft. deploying the parachute, and the backup charge firing at 800ft.

Based on preliminary analysis of the avionics flight data, all aspects of the recovery system worked as expected. A maximum deceleration of just of 16 gees was recorded at the drogue deployment, indicating that the calculations were valid. However, both the drogue and main parachute descent rates were slightly faster than expected, but were well within the safety limit of the vehicle and landing speeds. The data provided by the competition approved Perfectflite altimeter demonstrates the successful flight test of the recovery system and vehicle. This can be seen in Appendix F: Rocket Performance Calculations. It can be seen to very closely follow the Rocksim predictions of the flight profile.

Payload Integration The payload is being designed for easy integration into the rocket in such a way that

the design of the rocket body and the payload section does not affect the other. The payload section will be slid into the body tube and secured to the U-Bolt extruding from the large bulkhead at the base of the payload section. The payload will have no other connections or interaction with the rocket and will be a completely separate entity. All necessary external connections to electronics within the payload section will be located on the top bulkhead under the nosecone for easy access. The payload is being designed to fit snuggly within the rocket tube and therefore does not need to be attached in any other manner other than the U-Bolt.

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USLI FRR Hazard Analysis

As the project progresses this analysis will gain significant depth and will be a living document. Currently identified hazards are: Exposure to hazardous materials, exposure to deflagration by-products, customary machining/fabrication dangers, and rocket launch flight line dangers.

At the end of the Safety & Mission Assurance Proposal will be a compilation of all presently identified MSDSs for known hazardous materials that are applicable. This list will grow as the project moves forward. A preliminary list of MSDS sheets may be seen in Appendix C: MSDS Sheet. As each activity within this project is undertaken a complete set of procedures and relevant modification to the overall safety plan will be kept. This will include a description of all relevant personal protective equipment for whatever activity is being described. Mission Assurance Test Plan

A variety of test regimens will be employed to assure a successful flight and payload function at HARA in April. Review of the proposed project Gantt chart shows that the Vehicle will undergo flight testing first with a dummy payload and then with the actual payload under various conditions to provide data to refine the UCAT software predictive ability for the April launch. The payload itself will also be tested on the ground using shaker tables, etc. to simulate flight conditions before flight to evaluate its resistance to vibration and g-force and make modifications before flight testing. Flight testing of the payload will allow us to modify any unresolved technical issues and develop baseline data before the HARA flight. In the PDR, CDR, and FRR these test regimens will be fleshed out in greater detail.

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Launch Procedure

For the day of launch we will be providing our own 20ft tall launch rail which shall be erected upon a trailer for stability of the rail. As shown in the following figure, the launch rail satisfies the height requirement for the Panther II Heavy rocket. This requirement was derived from the need for a slow launching rocket to maintain dynamic stability before leaving the end of the rail. A higher speed launch would prove detrimental to the avionics and flight recorders for the slosh experiment.

Figure 2: The launch rail shall be fitted upon the pad shown on this trailer. (right) Figure 3: This 20 ft tall launch rail shall be used for the USLI competition. (left)

Outside of logistics set by MSFC for launch day, the Team will be keeping to strict guidelines, some of which mentioned in the Test Directive, and others mentioned here:

• Checklist for essential launch equipment including rocket components must be constructed

• Prior to procedures taken in rocket assembly, communication of said actions will be made to professional personnel present.

• Launch countdown will be loud and clear for the safety of all present.

The Team will carefully and meticulously test all electronics, wiring, fittings, U-bolts, and additional payload components. Once that is complete the rocket will be assembled and all shear pins and other outer bolts will be tightened. Finally the motor assembly will be fitted to the end and locked in place. While these procedures for rocket setup are being completed, other members will be adjusting the launch rail and preparing for the

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rocket to be attached. The Team will then carry the rocket to the rail and slide said rocket into place. The rail will be lifted and checked for alignment and finally the igniter will be inserted and secured. The wires for this igniter will then be run back to the Launch Director. The Launch Director will approve and implement any necessary deviations to this concept of operations. The following checklist will suffice for launch operation procedures:

Make a thorough run-through of Item checklist so that all parts are available. Tie avionics bay with yellow fabric to block black powder charge Charges attached

o Blue attached to 1 of 2 terminals o Red attached to 1 of 2 terminals o Check for wire security o Twisted in parallel o Total of two (2) charges attached

Kevlar rope rolled and avionics bay inserted Remove rocket stop from rail Secure fit avionics bay into rocket fuselage Insert motor Thread shock cord & attach to motor Secure motor bay (shear pins) Walk rocket to rail and secure Make 2nd stage of rail Raise rail and lock in place with large screw joints Igniter inserted Blue tape wrapped on end to secure igniter Continuity check (avionics) Retreat to ignition station and wait to launch Once rocket successfully returns to the ground send out “runners” Retrieve rocket in each separate section making sure not to add to any potential

damage Return back to Panther II tent for post-flight analysis

Launch personnel are protected to the greatest extent possible through pre-planned

procedures which are carefully reviewed and approved. These procedures control the entire sequence of events before, during and after launch. The controlling procedure is the launch countdown, which determines who does what activity and when this activity will be performed. This procedure is carefully monitored to ensure that there are no deviations from this document.

Personnel and equipment safety is provided by restricting these assets from areas where danger exists through implementation of the Launch Danger Area (LDA), described in this plan. Spectators will either be sheltered in the blockhouse or remain in designated areas outside the Launch Danger Area, greater than 300 ft from the launch rail.

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Failure Modes The failure modes of the airframe currently being examined are the possibility of

buckling under ascent, fin flutter, and shock cord damage during recovery. A complete set of buckling calculations along with distributed loads will be completed. Currently, we have done a preliminary buckling calculation on the overall length of the rocket. Fin flutter is another common and serious problem with high powered rockets. The fluttering can steer the rocket off course or cause the fins to come completely detached. We will use FinSim in conjunction with RockSim to calculate the flutter velocity of the fins. Knowing the maximum velocity of our rocket will allow us to adjust the shape of the fins so there is no chance they will flutter.

One common failure mode in large rockets occurs during the recovery stage of flight. We consulted Ky Michaelson of Rocketman Enterprises to ensure the parachute could withstand the forces we needed. Another step to keep the rocket from separating from the parachutes during recovery was to attach an active reefing device. This will open the main parachute slowly keeping the rocket from tearing away from the parachute. Another failure would be the shock cord tearing through the side of the rocket during either of the two recovery stages. The solution to this is currently being investigated. We are also running redundant altimeters of different makes. This will greatly reduce the risk of some sort of a hardware or software flaw from one manufacturer hindering the safe recovery of the rocket.

One last major failure mode is a CATO of the engine. By using the LOKI hardware that several of the team members have experience with, along with several static fires and test flights planned, we will have the assembly of these particular motors down pat. Personnel Hazards

Philip Meyer and Greg Peebles will be the Safety Officers for the team. Due to dusty

and fibrous nature of fiberglass, all personnel machining and fabricating parts made from fiberglass will be respirator trained and certified by the University Safety Officer. The fabrication of the airframe and any other parts made from fiberglass or another composite will be done in the schools composite laboratory in a well ventilated area. All non composite parts that will need to be fabricated will be done in the schools machine shop by certified students. All hazardous materials will be dealt with as the need arises by the University Safety Officer or the University Director of Laboratories. Any pyrotechnic charges will be handled and assembled by the teams BATF licensed personnel.

Following the NAR safety code, a perimeter of 100ft will be cleared of debris and all personnel (team members or otherwise) will be required to stand a minimum 300ft from the rocket. If a misfire occurs, the ignition system will be shut down and a 60 second hold will be started before anyone can approach the rocket. In order for the rocket to be launched, a set of criteria must be met; winds less than 20 miles per hour, sky must be clear of aircraft, the trajectory must not take the rocket into clouds or over people, the launch rail must be within 20 degrees of vertical, and the blast area from the motor must be cleared of dry grass or other easily flammable materials.

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Environmental Concerns

Some concerns we have are the fire danger of the area we will be testing, safe disposal of waste material and spent motor parts. Using the Keetch Byram Drought Index, we will not launch if the drought index for our launch site is at or above 600. All of our waste material will be disposed of using the method suggested by the University Safety Officer. We will be provided with waste disposal containers for used epoxy and fiberglass, oily rags, and other non-landfill wastes and hazardous materials.

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Payload Criteria Selection, Design, and Verification of Payload Experiment

The payload for the Panther II Heavy rocket is designed around observing slosh of a liquid within a rocket in microgravity. To obtain this goal, the payload section will contain a tank half filled with water, a camera to view the tank under all flight maneuvers, a video recorder for playback of the video, and a 6 Degree of Freedom (DOF) data recording system for the use of comparing the observed data to a Computational Fluid Dynamics) CFD simulation of the flight. Figure 4 below shows the layout of the payload section as designed in Pro Engineer. To make work easier, the design will separate the payload into two distinct systems: the video and slosh tank section and the 6 DOF section.

Figure 4: Science payload assembly diagram

To capture the video coming from the camera, we opted to go for a simple mini

VCR. This VCR uses a micro SD (Scan disk) card to store the video in real time. We will be using an inferred interface to control the recording and saving of video data during flight. This will also be controlled by the Programmable Interface Controller) PIC micro controller so that it is precisely timed with the 6 DOF sensors to capture useful video data. The entire system will be triggered by a trip wire attached to the side of the rocket

In order to compare the actual launch of the rocket to the simulated launch in UCAT, all six degrees of freedom need to be measured and recorded during flight. To accomplish this task, Panther II Heavy will be using a series of accelerometers and gyros precisely positioned to measure the axial and rotational accelerations

6 DOF data recording system

VCR

Slosh tank

6 DOF data recording system

Connection to lower part of rocket

Camera

Two 12 Volt NiMH Batteries

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respectively. To improve reliability and redundancy, Panther II Heavy will contain two complete and independent 6-DOF systems that are fully capable of measuring and recording the accelerations of a rigid body. Should both systems log the data correctly, the two sets of data can be combined to characterize the complete motion of the rocket even if it does not stay completely rigid. Since the initial conditions of position and velocity are known, integration of these accelerations will provide an accurate trajectory plot as well as a velocity plot with respect to time which then can be input for a CFD simulation of the slosh experiment.

Furthermore, the accelerometer chosen to be incorporated into the circuit boards is a ±18 g dual-axis accelerometer pictured below in Figure 5. The accelerometers will each be responsible for measuring the acceleration in two directions, so their positions on the board must be such that all three axial directions are represented. The gyro chosen to be included in the electronics board is a ±75 degree/sec gyro picture below in Figure 6. The gyros only measure the rotation acceleration in one axis, so three gyros will have to be positioned around the board to record all relevant data. These accelerometers and gyros are analog devices, which means that the reading is accurate to an infinite number of digits and is available at any time scale. In order to guarantee data throughout the entire flight, the rocket is being designed to stay under 18 g’s of acceleration and 75 degrees/sec of rotation. This means that the initial thrust from the rocket motor and the ejection charge of the recovery system will be designed specifically not to exceed these parameters.

Figure 5: ±18 g dual-axis accelerometer (left) (Sparkfun) Figure 6: ±75 degree/sec gyro (right) (Sparkfun)

Other important components on the circuit board are the PIC and the EEPROMs (Electrically Erasable Programmable Read Only Memory). The PIC is the brain of the electronics section as it is the responsible for handling all communications between each component on the board. The PIC will collect the data from each accelerometer and gyro, convert the analog signal into a digital format, then send each reading to the EEPROMs for saving. The PIC is also responsible for sending the collected data to a computer after each flight through a wired connection. However fragile the PIC is, it should not be seeing any loading prior to, during, or after the flight, and therefore we are not concerned with the PIC breaking. The EEPROMs, as their name suggests are the memory on the board. Four EEPROMs are being used on each board to ensure enough

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data storage space. The EEPROM is ideal for our application because it saves data quickly (on the order of a couple milliseconds), and should power be lost during flight, the data is safe from being erased because the EEPROM does not need power to keep previously saved data.

In order for the board to survive the g forces that are expected during flight the board must be completed with the utmost care. The parts that are being used are connected to the board using solder holes. If the hole isn’t completely filled and the components aren’t held in place properly they could break off during flight causing the data to be lost for the flight. The board has to be sturdy enough to survive multiple flights to get the correct data so the components will be checked after each flight to ensure the accuracy of the data.

In order to test all of the components and make sure that they are working properly the sensor boards will be tested on a shaker table. The shaker table is a machine that will “shake” the board in a variety of directions, rotations, and speeds to get a wide sample range to verify the measurements over the operating range of the sensors. The sensors data will then be compared to a known standard to ensure that the data received from the sensors is as accurate as possible. After this the sensors will be calibrated against the known standard so the analog data that is retrieved from the EEPROM’s can be converted into a known g force and rotational speed measurements. After this is done we will be placing the boards into a test rocket to determine if they will be able to survive the flight and accurately record the data. The comparison for the actual flight will be with a ROCKSIM simulation that the computer will generate given the rocket specifications and flight day weather conditions.

Currently the first version of the board has been completed and testing the system will be beginning once the final components are received from the manufacturer. After careful consideration it was decided to use a different PIC then the one that we originally planned on using. The new PIC (18LF542) is exactly the same as the one that was planned on being used with the exception that it can operate at three volts instead of five volts. This will also allow us to keep the sensor data sheet information intact as the sensors are calibrated to use three volts in their operation instead of five. The initial reason for running five volts through the board was to power the VCR through the board but because of the restriction on time we are simply going to add a voltage step-up to the VCR to take the three volts running through the board and increase it to five volts so the VCR can operate. If we were to use five volts as originally planned we would have to create our own data sheets to verify the g forces and rotational speed are accurate. As the board is tested we are going to make sure that all the components work properly and troubleshoot any problems that may arise. If there is anything that is seriously wrong a revision to the schematic will be made and new boards will have to be ordered and the testing procedures will start again.

Since the boards have not been tested yet there is no way to show how accurate or reliable the boards will be. In Appendix D there are data sheets for the sensors that can show how accurate the sensors are and what their operational limits are.

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Figure 7: Circuit diagram (left) compared to completed board (right) Payload Concept Features and Definition

The slosh dynamics in cryogenic fuel tanks under variable acceleration profiles, including microgravity, is a pressing problem that severely affects the reliability of spacecraft launching. The science payload that will be flown on the Florida Tech USLI rocket will be dedicated to making measurements of liquid sloshing during the rocket’s flight. This science payload addresses a need within the rocket propulsion community for an improved understanding of how propellant slosh is coupled with vehicle dynamics.

Slosh is a pressing problem for spacecraft stability and control. As fuel is consumed, the fuel tank becomes increasingly empty, and the remaining fuel is excited by the motion of the spacecraft. Since the reaction forces and moments caused by fuel slosh can degrade the pointing accuracy of the system, it is important to predict the effect of slosh on the spacecraft attitude control system, [45, 45]. Control of vehicle position is a demanding task during rendezvous and docking, and virtually all attitude control scenarios are affected by slosh. Researchers at the NASA Marshal Space Flight Center (MSFC) have indicated that slosh motion with the Space Shuttle replacement Orion Crew Exploration Vehicle (CEV) is a critical concern slosh during ‘delicate’ docking maneuvers with the International Space Station (ISS).

The Launch Services Program (LSP) at the NASA Kennedy Space Center has identified the modeling of liquid propellant slosh in the upper-stages of launch vehicles as a critical action item to ensure successful mission planning and execution. For example, in preparation for orbital insertion of a payload, the upper-stage of a rocket typically undergoes a series of maneuvers which may lead to sloshing motion of the propellants. An example of a Boeing Delta II upper stage is shown in Figure 8 and Figure 9.

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Figure 8: Schematic layout of a Delta II upper-stage. The propellant tank shown on the left (larger tank) contains liquid hydrogen and the propellant tank to the immediate right of it contains liquid oxygen.

Figure 9: XSS-10 view of Delta II rocket: An Air Force Research Laboratory XSS-10 micro-satellite uses its onboard camera system to view the second stage of the Boeing Delta II rocket during mission operations Jan. 30, [17].

Liquid propellant reaching the relief and orbital control vents within the tanks may

result in a significant increase in expelled mass causing a dynamic instability which may lead to mission failure, [4]. An example of the highly dynamic, chaotic motion of a liquid propellant sloshing within a tank during a spacecraft reorientation maneuver is shown in Figure 10.

Figure 10: Fluid sloshing in propellant tanks under variable acceleration maneuvers, [4,16]. Slosh modeling and experimental results cannot be found in the open literature for the scales and conditions relevant to current upper-stage mission profiles. NASA KSC, Sierra Lobo, Inc., and the Mechanical and Aerospace Engineering Department at Florida Tech are working together to develop a highly instrumented experiment to make the first detailed measurements of fluid slosh dynamics under relevant low-gravity conditions. These data will then be correlated against computational fluid dynamics (CFD) models to develop an enhanced predictive tool to assess propellant sloshing on mission performance.

Another important scenario in which slosh can be critical is the increase in nutation (wobble) in spinning spacecraft - an instability that arises from the dumbbell formed by the center of mass of the liquid and the center of mass of the dry spacecraft when the rotation axis corresponds to the minimum moment of inertia. Even a minute amount of liquid (e.g. 1.2 kg in a 452-kg spacecraft) can induce catastrophe. An additional consideration in launch vehicle design is potential overlap of the slosh resonant frequencies with the control system and structural dynamics frequencies. To date, current numerical models which examine the dynamics of fluid-structure coupling during

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slosh events have not been benchmarked against variable acceleration and experimental data, and a direct comparison of measured slosh with numerical predictions would be of significant interest to the engineering and scientific community. In another example, the NEAR spacecraft interrupted its insertion burn when fuel reaction was larger than anticipated. The result was that this unpredicted dynamic behavior prevented NEAR from orbiting Eros and resulted in a significant mission delay. Science Value

A new framework is proposed for the experimental and computational study of slosh under dynamic conditions. It is the goal of the Florida Tech USLI team to provide a data set which researchers can use to benchmark CFD codes. This data set will be gathered on the 1-mile altitude flight solid rocket used for the USLI competition. The payload will contain a small water-filled tank and the slosh motion of the water will be monitored during flight – from the high acceleration lift-off through deceleration and zero-gravity through apogee. The video data of the slosh tank will be correlated with full six-axis acceleration measurements taken during the rocket’s flight. An overview of how the accelerations measured during the rockets flight are utilized in CFD codes and how the results of those codes are compared with images acquired during the experiment is shown in Figure 41.

Figure 11. Proposed framework for experimental benchmark of CFD codes. The translation and rotation components of the 6-DOF acceleration vector are used to predict a slosh event.

Many more details regarding slosh research, and in particular slosh research at

Florida Tech, can be found in Appendix A. The next sections of this work review previous slosh research completed at Florida Tech, as well as details on the experimental payload that will be used in the USLI experiment. Scientific Research Objectives for Panther II Heavy The specific research objectives are summarized below:

1. Development of an experimental methodology to characterize slosh dynamics by flying a high-instrumented experimental package to measure liquid slosh events during the Panther II model rocket flight at the 2008-2009 USLI competition.

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2. Use the data recovered from the Panther II rocket (6-DOF acceleration history and images of liquid slosh) to further benchmark the predictive capability of modern CFD tools for slosh applications in variable acceleration environments – including dynamic mesh tools.

3. Use the 6-DOF acceleration history of the rocket to further benchmark the NASA KSC UCAT rocket simulation program.

The payload will be approximately five pounds and designed to fit inside of a sixteen

inch long by four inch diameter tube that will securely fit into the rockets payload bay. An approximate design can be seen in Figure 12. This modular approach will allow the scientific payload to be quickly swapped out for the dummy payload, thus greatly reducing the risks during airframe testing. Because the payload will be modular and completely separate, the airframe will be far less complex and much easier to design and fabricate.

Figure 12: Proposed layout for payload within body tube in payload compartment. Body tube not

shown above.

In order to select the electronics system that will be used to make measurements of the acceleration system of the rocket during its flight trajectory, some background is in order. Liquid sloshing is the result of the relative motion between a liquid and its container. A novel framework to characterize fluid sloshing is proposed here: the time history of the rigid body acceleration of the tank relative to an inertial frame is sufficient to uniquely characterize a sloshing event, provided that the initial liquid distribution within the tank is known. The proposed experimental implementation of this approach is

Nosecone

Payload

Aluminum bulkhead (inside coupler tube)

Coupler tube for insertion into rocket

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therefore based on the simultaneous measurement of the 3-axial acceleration of the sloshing tank at two points, plus measurement of the ax, ay, az components of the angular velocity vector. These are sufficient to uniquely determine a slosh event, for a given initial distribution of liquid within the tank.

The experimental characterization of the acceleration of a rigid body can be done in a number of ways. A first approach is based on the use of three 3-axial accelerometers attached to three different points within the tank’s support. The basic kinematic relationship between these measurements, for any two points, is:

/ /( )B A B A B Aa a r rα ω ω= + × + × ×

r r rr r r r (1) Where Bar and Aar are measured by the accelerometers located at points A and B, αr

and ωr are the angular acceleration and velocity of the tank, respectively, and /B Arr is the relative location of B with respect to A. Three sets of tri-axial measurements ( , ,A B Ca a ar r r ) should be sufficient to provide six equations whose unknowns are the components of αr and ωr . However, it has been found that such a system of six equations with six unknowns is sparse and often singular. As a result, when using tri-axial accelerometers, the minimum number required to uniquely determine the rigid body acceleration of an object, is four. This implies that any rocket project attempting to merely use 1 or 2 tri-axial accelerometers to measure the acceleration history of their rocket’s trajectory is not appropriate – either more tri-axial accelerometers must be used, or a combination of tri-axial accelerometers and gyroscopes must be employed.

A simpler approach consists of six acceleration measurements (e.g. from two tri-axial sensors A, B) and three orthogonal angular rate sensors to measure the components of the angular velocityωr . This is the approach that will be used in the USLI competition. A straightforward approach to experimentally characterize slosh events is presented here. It is based on the simultaneous measurement of the rigid body acceleration of the sloshing tank as described above, along with simultaneous capture of the shape and velocity of the sloshing surface based on an array of orthogonal camera synchronized by a hardware trigger.

The challenges of the experiment include: dampening the video camera, the software program to collect data, having the accelerometers measure all vibrations, and testing the sensors before launch. During the flight the rocket might have vibrations that can diminish the resolution of the camera that we are using. The camera has to be isolated from the rest of the rocket to make sure that proper resolution is maintained. If vibrations are too high, there is the possibility that the components inside may shake themselves apart if they are not properly made and fastened. The sensors that we are using can measure up to 18 g’s before saturation, so our rocket cannot exceed this certain acceleration that is to be determined. Assembly

The payload has been designed for ease of assembly and integration into the body of the rocket. The payload itself will be assembled and attached to the large upper bulkhead which will then be frankenbolted to the body tube. The only other need for the

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payload section is three holes near the main electronics board for connections to the triggering device. As can be seen in Appendix E: Payload Calculations, the payload structure is very durable and all components are designed to withstand the entire weight of the payload at 100g acceleration. This is a valid assumption since the test flight experienced a maximum of 60gs during the main parachute ejection charge. This in part is due to the high level of attention to detail for this section; the members were cut on a CNC machine or commercially bought. The CNC is capable of machining the raw material to greater than ±0.001 precision.

The following show a step by step pictorial procedure for the assembly of the payload section:

Step 1: Attach two L-brackets to the bottom of one rectangular board

Step 2: Attach one payload bulkhead with two L-brackets as shown

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Step 3: Connect the 6DOF electronics board using standoffs

Step 4: Place both batteries in position and secure with double sided tape. Secure

onto rectangular board with bent piece of aluminum and double sided tape

Step 5: Attach camera to second payload bulkhead leaving nuts in place for the L-

bracket all thread rod. Using coupler tube as spacers, position slosh tank and secure with large all thread rod

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Step 6: Attach second rectangular board using L-brackets

Step 7: Attach second 6DOF board in same manner as the first. Attach the

uppermost L-brackets

Step 8: Attach VCR in same manner as the batteries

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Step 9: Attach the payload cap to the topmost L-brackets

Step 10: Attach lower all thread rod to upper bulkhead Step 11: Attach all wires to their respective locations and secure with zip ties

Step 12: Slide payload into payload bay and secure the upper bulkhead with four ¼-20 frankenbolts

Failure Modes of Payload

The payload failure modes consist of breaks or cracks in the slosh tank and electronics breaking. The slosh tank will be filled with water and there is a possibility due to improper manufacturing or exceeding force limits the tank could crack causing the water to leak out or lead to catastrophic failure in the tank. The two circuit boards that we are going to be using are made up of sensitive electronics that could potentially break during the flight. If the slosh tank breaks it could cause water to leak onto the boards shorting them out. Another possibility would be the accelerometers and gyroscopes coming loose or breaking off during flight. The VCR that we are using to record the flight data may not engage properly because of the wire coming out of the solder hole or the electronics could break causing the system to fail. With the compartmental system that we are using there is the possibility that the G10, due to the forces experienced during flight, could fracture and break. The most likely place for this to happen would be underneath the slosh tank as it is the heaviest part of the payload section. If a catastrophic failure happened it could crush the components underneath it and change the center of gravity of the rocket causing the rocket to go out of control and possibly crash. One of the reasons for using the two sensor boards is to collect data for the UCAT program even if one of the boards fails. Due to the board not being completed yet or accurate specs of the materials from the manufacturer no analysis of the unit can be completed so the limits of what kind of forces the electronics can handle cannot be completed. From initial analysis of the rocket body and the forces that it will experience we predict that the electronics will be able to function properly throughout the flight. Because there are no toxic or harmful materials in the payload the only associated hazards would be cuts and bruises from mishandling the equipment.

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Safety & Mission Assurance Regulatory Compliance The MSFC USLI Competition safety requirements regarding ATF, DOT, EPA, FAA, OSHA, & TRA/NAR are already in place within the existing Florida Tech Rocketry Program. These will be discussed in their respective areas within this plan.

Sport Rocketry Certifications: Greg Peebles TRA(05964)/NAR(72854) Level 2 Justin LaFountain TRA(11722) Level 1 David Jarkey TRA(11267) Level 1

Sport Rocketry Association Affiliations: Spaceport Rocketry Association (SRA) http://www.spaceportrocketry.org/ Northeast Florida Association of Rocketry (NEFAR) http://www.nefar.net/ Tripoli Mavericks http://www.rocketmavericks.com/

Aviation Regulatory Experience: Regular Correspondence & Communication with FAA Southern Region Office in Atlanta arranging high power rocket launches and other singular rocket waiver activity per FAR Part 101. Communications with FAA AST relating to sounding rocket activities as well as USAF 45th SW.

BATFE Licensure:

Greg Peebles User of Low Explosives 1-FL-009-34-9B-12637 BATFE Approved 50# Low Explosives Storage & Transport

DOT Hazardous Materials Training (per 49 CFR 172.704): Greg Peebles September 28, 2007

Florida Tech General Safety & Insurance Requirements

In addition to the MSFC USLI safety requirements, the FIT Senior Design Safety Plan Requirement analysis will be followed during the course of this competition. This can be seen in Appendix B. Federal Bureau of Alcohol, Tobacco, Firearms, and Explosives (BATFE)

BATFE regulates the following aspects of this project: Purchase, storage, transport, and use of all Explosives. Within this project the following regulated materials will be used: Ammonium Perchlorate Composite Propellant (APCP), Black Powder (BP), Electric Matches (EM), Rocket Motor Igniters (Igniter). These will all be handled per ATF Publication 5400.7 a.k.a “The Orange Book.” Federal Department of Transportation (DOT)

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While BATFE controls the ownership and ability to transport items intrastate, the

DOT controls how those items may be transported. The Low Explosives APCP and BP are subject to the strictest DOT hazardous transportation regulations of the BATFE regulated materials. Transportation of these items will be done to conform with: 49 CFR 173 with special emphasis on Subpart C; 49 CFR 172.411; and 49 CFR 172.504(c). Federal Environmental Protection Agency (EPA)

As this project is performed solid wastes will be generated. Several of these solid wastes are classified as Hazardous Wastes (e.g.: Fiberglass dust, cured epoxy, uncured epoxy contaminated materials are all “Flammable Solids, Organic N.O.S.” solder contaminated sponges & discarded solder coated metal is also a heavy metal contaminated Hazardous Waste) and Universal Wastes (Electric Component discards). Florida Tech is classified as an EPA Large quantity Generator of Hazardous wastes per RCRA, as such all such wastes will be handled per RCRA law in 40 CFR 265. Federal Aviation Administration (FAA)

Rocket flights will all be conducted with appropriate waivers to 14 CFR, Subchapter F, parts 101.21 thru 101.23 and compliance with part 101.25 and any special additional requirements contained within the FAA issued waiver(s) for the launch site(s) SRA – Palm Bay, FL; NEFAR – Bunnel, FL; HARA – Manchester, TN. Federal Occupational Safety and Health Administration (OSHA)

OSHA regulates all aspects of the construction and operation of this project. This falls into 3 areas, the first two of which are interrelated: 1) Use and Training to use personal protective equipment (PPE), 2) Permissible exposure limits to hazardous substances. The MSDS for any given material used will identify what if any exposure issues exist and 29 CFR 1910 will identify what training and use restrictions are required for a given PPE. 3) All general accident avoidance activities. National Association of Rocketry (NAR)

In addition to the applicable Federal standards, this competition is subject to the safety code of the NAR. This can be readily found at http://www.nar.org/NARhpsc.html . In addition to this specific code, this will involve absolute compliance with the Huntsville Area Rocketry Association’s (HARA) Range Safety Officer the day of the competition launch.

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Rocket Simulation The NASA Launch Services Program’s (LSP) Universal Controls Analysis Tool

(UCAT) is used to analyze the flight performance of a series of specific launch vehicles. In order to demonstrate the modeling tool as a unique LSP capability, a series of non-proprietary, generic launch vehicles have been simulated ranging from relatively simple single-stage, solid propellant sounding rocket to a more complex two-stage liquid rocket. Working in close collaboration with the Control Systems Analysis group in the Mission Analysis Branch at the NASA Kennedy Space Center and the 45th Space Wing at Patrick Air Force Base new features and capabilities will be added to the UCAT modeling suite. UCAT is a graphical user interface program created within Simulink 6.3 (© Copyright 1990-2005 MathWorks, Inc.) to create a user-friendly environment for engineers and developers as seen in Figure 13.

Figure 13: UCAT CG Kinematics Block

UCAT is a six-degree of freedom rocket simulation program with various coordinate

frames. These coordinate frames are flat earth, rotating and non-rotating spherical earth, Earth Centered Earth Fixed (ECEF, non-rotating oblate spheroid earth model), and Earth Centered Inertial (ECI, rotating oblate spheroid earth model). UCAT uses the US Standard Atmosphere model 1976 to simulate its atmosphere from 0-1000 km of altitude. Various gravity models are used for simulation including dependence on inertial position relative to the earth’s coordinate frame. UCAT has been linked with other third party software during simulation to increase fidelity of the rocket simulation. With relevant input data UCAT can simulate the effects various aerodynamic forces and moments a rocket encounters. Simulating solids, hybrids, and liquid propellants with high accuracy has been the focus of UCAT’s development since Florida Institute of Technology (FIT) has been involved. UCAT also has the capability of simulating the three forces, and moments a rocket encounters due to its propulsion unit. Most rocket simulation programs simulate only one degree of freedom the rocket encounters for

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thrust. However FIT has an unusual thrust stand that can measure six degrees of freedom on its thrust stand that yields more accurate results when simulating a rocket motor in UCAT.

For the preliminary design phase of the rocket RockSim 8.0 will be used to get an idea of which motor to use for a conceptual rocket. RockSim 8.0 is a commercially available rocket simulation program by Apogee Components, © Copyright 1997-2003 Paul L. Fossey-PKF Systems. This utilizes three degrees of freedom (up/down, up-wind/down-wind, and pitch) simulation. The maximum altitude is 86 km, and it uses only a flat earth model. RockSim 8.0 cannot simulate a rocket with an initial roll rate, launch above ground, canted nozzles, or any initial velocity. RockSim 8.0’s highest degree of accuracy can be obtained in pure vertical flight with no wind. The preliminary simulations can be seen in Appendix E: Payload Calculations Pin Shear: Assuming: 4.5lbs at 100 g’s on the pins The pin at each section is in double shear Diameter of the pins is 0.164 in (we are using an 8-32 bolt) (McMaster Carr) The bolts will be made out of Al6061 T6 Shear strength (SS) of the bolts is 30 000 psi (Matweb)

Lessons learned: The bolts will not be the weakest link in the payload section and are designed to carry a load much greater than they will ever see. Pin Tensile Stress: Assuming: 4.5lbs at 100 g’s on the pins Each pin holds half the total weight of the payload Minor diameter of the pins is 0.1248 in (we are using an 8-32 bolt) (Engineer's Handbook) The bolts will be made out of Al6061 T6 Tensile Strength (TS) of the bolts is 40 000 psi (Matweb)

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Lessons learned: The bolts will not fail in tension even under 100g acceleration of the payload section. Pin Tread Pullout Stress: Assuming: 4.5lbs at 100 g’s on the pins Each pin holds half the total weight of the payload Minor diameter (Kn) of the pins is 0.1248 in (we are using an 8-32 bolt) (Engineer's Handbook) Number of threads per inch (n) is 32 for an 8-32 bolt (Engineer's Handbook) Minimum pitch diameter of external thread (Es) is 0.1399 (Engineer's Handbook) The bolts will be made out of Al6061 T6 Shear Strength (SS) of the bolts is 30 000 psi (Matweb) There will be two nuts on each end of the rod to double the fastener thread engagement

(Le) but for the purpose of worst-case scenario, the Le will be used for the shear area calculation

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Lessons learned: In the worst-case scenario of one of the nuts falling off, the bolt will not fail under a 100g load. Column Buckling: Assuming: 4.5lbs at 100 g’s on the column Clamped at both ends (Le= ½ L) Each column is made of G10/FR4 fiberglass E=2 200 000 psi (Professional Plastics) L=6.5 in w=3.70 in (width of each column)

Substituting in the moment of inertia I and rearranging,

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Lessons learned: The G10/FR4 chosen will be 0.187 in thick, which is the closest commercially available thickness for a factor of safety of 2. Therefore, buckling will not become a problem. G10/FR4 Bearing Stress: Assuming: 4.5lbs at 100 g’s on the column from the bolts The column is made of G10/FR4 fiberglass Thickness of fiberglass=0.187 in Diameter of the bolt is 0.164 in (a #8 bolt) (McMaster Carr) Tensile strength (TS) of the G10/FR4 is 35 000 psi (Professional Plastics)

Lessons learned: Even in the worst case scenario, the bolts will not rip through the fiberglass that is holding the payload. Tube Buckling: Assuming: 30lbs at 10 g’s on the tube Worst Case Scenario of one fixed end and one free end (Le=2L) The tube is made of G10 fiberglass E=2 200 000 psi (Professional Plastics) L=132 in (This is the entire length of the payload tubing) do=4.03 in (This is the outer diameter of the tubing) di=3.91 in (This is the inner diameter of the tubing)

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Lessons learned: Buckling of the payload tube is of major concern if the payload tube were empty. Because the tube will be filled with the electronics, the tube will not really buckle at these loads.

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Appendix F: Rocket Performance Calculations Recently Apogee Components has purchased Splash, a six-degree of freedom

rocket simulation program, while splash does simulate wind effects, rotating oblate spheroid, gravitational effects varying with altitude and latitude, ISO 1978 standard atmosphere up to an altitude of 631 km. The purpose of Apogee Components purchasing Splash is to upgrade RockSim from three to six degrees of freedom. However, according to Sean P. Stapf, Ejection Seat & Rocketry Analyst, USAF 45th Space Wing: “Splash results have been an absolute disaster”, the aerodynamic forces were not acting in the correct direction of the rocket. Paul Fossey of Apogee Components assures that the new correction of errors from Splash were completed, and assimilated into a new version of RockSim (RS-PRO).

UCAT is following the same validation plan as used by the USAF 45th Space Wing. This is done by the use of an unclassified ‘well-worn’ multi-purpose rocket called Hydra. UCAT is completely independent of the popular commercially available rocket simulation programs for validation and comparison. Florida Institute of Technology will be using UCAT for its rocket simulation program due its capability to simulate the competition rocket’s trajectory, and dynamics affecting the rocket’s payload. Application to NASA KSC’s UCAT

In order to perform independent verification and validation (IV&V) of vehicle performance, NASA’s Launch Services Program (LSP) and Florida Tech have developed the full 6-DOF Universal Controls Analysis Tool (UCAT) in order to analyze the flight performance of numerous launch vehicles and spacecraft. In order to improve capabilities, and enable true 6-DOF controls coupled with sloshing motion. UCAT and CFD can be effectively linked together to form a comprehensive simulation tool that will provide researchers with a unique modeling package in which to assess the impact of propellant slosh of rocket dynamics, as shown in Figure 14.

Figure 14: UCAT – CFD interaction model and information passing

To investigate a hybrid UCAT-CFD simulation tool for determining propellant slosh

motion on vehicles dynamics, the following work plan will be undertaken: 1. CFD slosh calculations with a controlled 1-DOF tank motion and relevant

information (forces and torques) exported to UCAT. a. A tank will be modeled in CFD, filled to 1/3rd volume with water (or

cryogenics), and the gravitational field set (g/gearth = 1 – 10-4). Surface tension, wall wetting angle, time-step, grid resolution, etc., are also set.

b. A user defined function (UDF) controlled by Matlab will be developed to input an arbitrary 1-DOF acceleration profile to FLUENT.

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c. CFD performs the unsteady calculation, sends results to UCAT, and receives the next acceleration step. The process is repeated providing a complete history of the slosh motion for a given acceleration-time history.

2. 6-DOF Motion with Slosh-Vehicle Feedback Interaction: An important aspect of this research is to examine how slosh motion influences vehicle performance. The purpose of this task is to couple slosh force with the motion of the tank:

a. An updateable force history is calculated to include internal fluid slosh forces in addition to external rocket thrust forces.

b. After the first time-step, CFD outputs the fluid force on the tank wall. Unlike Task 2 the fluid force now couples with the thrust, either primary thrust from the engine or thrust from the reaction control system (RCS).

c. UCAT updates the vehicle dynamics to include the sloshing force. Mass properties are updated for inclusion into the system dynamics equations.

d. The RCS sends new commands to compensate for slosh, resulting in new forces acting on the tank being sent to FLUENT for the next iteration.

The 6-DOF model is shown in Figure 15.

Figure 15: UCAT – CFD 6-DOF interaction model for each calculation time-step

3. UCAT- CFD for Actual Maneuvers and Slosh Mitigation Strategies: a. Utilize actual trajectories for launch vehicles and the CEV to study the

impact of slosh dynamic on vehicle dynamics and maneuvers. b. Examine UCAT’s potential for slosh mitigation by modifying the tank

geometry or through the inclusion of slosh baffles.

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Project Plan

Shown below in Figure 16 is a Gantt chart with our preliminary schedule for the project. At this stage of planning, the focus is on deadlines provided by the competition, shown by black diamonds, as well as deadlines created by ourselves, shown in red.

Figure 16: Panther II Heavy Gantt Chart

Project Budget

The funds necessary for Panther II to compete in the 2008-2009 USLI contest have

been procured and come from three sources: 1. Support from the Florida Tech Provosts Office, College of Engineering, and

Department of Mechanical and Aerospace Engineering: $5,000 2. A grant from the Florida Space Grant Consortium: $5,000 3. A donation from faculty advisor, Dr. Daniel Kirk, for the development of the slosh

dynamics research payload: $2,500 These funds are sufficient for designing and testing the rocket. The actual vehicle that flies in the USLI contest will cost less than $5,000, but additional funds have been set aside for testing and verification. A basic cost break-down is given below:

• Flight vehicle, parts and fabrication (including motor): $2,000 • Electronics and instrumentation: $2,500 • 10 motors for statistical testing on the Florida Tech thrust stand: $1,000

A detailed parts list can be viewed in Appendix H: Parts List.

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Appendix A: Background on Fluid Slosh Research at Florida Tech Literature Review and Background on Fluid Slosh Research

Analytical solutions based on first principles (the basic fluid and motion equations) have been developed for simple tank geometries, base excitations and boundary conditions based on potential theory to describe the velocity potential of the sloshing liquid, [1, 7, 11, 23, 24, 32, 44], and experimental verification of simple, low-amplitude slosh (without droplet break-up) has been performed, [1]. Simplified models of lateral excitation of a liquid with a free surface have been developed using a spring-mass representation, and the liquid motion can be represented by a second order linear differential equation

Theoretical and experimental work has been done on establishing the damping characteristics of flat, concave, and convex bottom tanks. The complex damping characteristics of large-amplitude slosh motion was of major interest during the U.S. and Russian space programs, and basic experimental work studying this behavior was accomplished, [31, 42], with strong nonlinear effects observed when the tanks were excited near their eigenfrequencies. Furthermore, there is a sensitive the gain at resonance is to the damping coefficient. Historically, damping has been difficult to predict analytically especially for non-linear large amplitude slosh with accuracy sufficient for design verification.

More recently, attempts have been made to use computational fluid dynamics (CFD) models to make slosh predictions during the high acceleration ascent phases of a rocket, although very little work has been done in cases of very-low gravity when the vehicle is in space, [44, 45]. When available, predictions can be used to validate the performance of simpler models, and a direct comparison of measured damping with CFD-predicted damping is of significant interest.

It is generally difficult to predict slosh motion, and the problem becomes even more challenging when the tank has internal structures such as baffles or diaphragms. One approach for predicting slosh motion is to use scaled model testing, such as that done at Southwest Research Institute, [15], the results are largely qualitative and there has not yet been direct data comparison with detailed numerical (CFD) models. Under the effect of a large force field such as Earth's gravity, the liquid center of mass trajectory is easier to predict. When liquid excursions remain small, a linear second-order oscillator provides a useful representation of the slosh dynamics, which can be integrated with the state equations of the complete vehicle system. More complicated plants result from nonlinear slosh models, which have been shown to yield reliable predictions over larger ranges of coordinates, or by including additional oscillators and an increased state dimension [5]. Recently, new measurements tools, instrumentation, and state-of-the-art numerical models have become available which can be utilized to extend our understanding of fluid slosh beyond simple lateral excitations to cases of large magnitude slosh, fluid break-up and droplet formation, and the ability to couple these effects to the dynamics of the spacecraft.

Although predicting the dynamics of slosh has been actively investigated since early stages of the US space program, [7, 9, 35], a satisfactory formulation is yet to be found. Substantial efforts have been made in recent years to improve the modeling, simulation and control of slosh dynamics. Previous work in slosh dynamics has been mostly theoretical, [7, 8, 9], or treats the mass of fuel as a variable inertia only, i.e. the viscosity, surface tension and other important fluid effects are not considered, [37]. Other studies have focused on identifying and analyzing available flight and test data to identify conditions leading to mission failure. The FLEVO project, under the direction of the National Aerospace Laboratory (The Netherlands) has been perhaps the most substantial effort devoted to fill the gap between numerical simulations to predict slosh, [24, 44], and the development of an experimental framework to measure and characterize slosh under microgravity, [45].

The experimental characterization of sloshing fluids is a challenging problem. The most important techniques previously used have been: (i) using load cells to measure the support force holding the tank where the sloshing motion is taking place, as described in several papers from South West Research Institute1, (ii) arrays of capacitance probes, such as those used in the FLEVO project, and (iii) ultrasonic ranging techniques.

Load cells can provide valuable information when trying to determine physical parameters of a lumped-parameter model of the sloshing liquid, such as damping ratio or natural frequency; therefore, since load cells can only measure the forces and torques acting on the tank at the support point, they cannot provide detailed information on the distribution of fluid inside of the tank. Several different fluid configurations can produce the same forces or torques at the support point, i.e., it is not possible to

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uniquely determine a slosh event from a time history of forces and torques measured at the support point since many different fluid configurations may produce the same forces and torques.

Capacitive methods estimate the thickness of the fluid wetting the tank’s wall based on the change in capacitance of an array of sensing elements attached to the tank’s wall. The main shortcomings of this method are the limitation in thickness range that can be measured, and the fact that only a discrete number of points in the tank can be monitored. More resolution is obtained at a high price in cost and complexity - the FLEVO experiment, for instance, included two-hundred seventy such probes.

Ultrasonic ranging methods have been used to accurately determine fill levels in cryogenic fuel tanks. They work well in quasi-static conditions, and when measuring fill levels in a small number of discrete points. They provide the approximate distance from an ultrasonic source to the point in the liquid’s surface where the beam is pointed at a given time. The use of ultrasonic ranging does not seem well suited to dynamically measure a sloshing surface, and has not been described for this purpose to the author’s knowledge.

Previous research on slosh dynamics by the South West Research Institute1 was used as a basis to gain insight on the forces that dominate liquid motion during slosh. Three non-dimensional quantities, the Bond, Weber and Froude numbers, indicate the dominant flow regime. They are calculated from characteristic parameters, such as velocity, acceleration, density, viscosity, and tank radius. These non-dimensional numbers relate capillary, inertial and gravitational forces. The Bond number relates capillary forces to gravitational forces. The Weber number relates capillary forces to inertial forces. Finally the Froude number relates inertial forces to gravitational forces. A full description of these parameters and the corresponding flow regime intervals is available in the SWRI1 report. Current Research at Florida Tech: Experimental Efforts

Liquid sloshing is the result of the relative motion between a liquid and its container. A novel framework to characterize fluid sloshing is proposed here: the time history of the rigid body acceleration of the tank relative to an inertial frame is sufficient to uniquely characterize a sloshing event, provided that the initial liquid distribution within the tank is known. The proposed experimental implementation of this approach is therefore based on the simultaneous measurement of the 3-axial acceleration of the sloshing tank at two points, plus measurement of the ωx, ωy, ωz components of the angular velocity vector. These are sufficient to uniquely determine a slosh event, for a given initial distribution of liquid within the tank.

The experimental characterization of the acceleration of a rigid body can be done in a number of ways. A first approach is based on the use of three 3-axial accelerometers attached to three different points within the tank’s support. The basic kinematic relationship between these measurements, for any two points, is:

/ /( )B A B A B Aa a r rα ω ω= + × + × ×r r rr r r r

(1) Where Bar and Aar are measured by the accelerometers located at points A and B, αr and ωr are the

angular acceleration and velocity of the tank, respectively, and /B Arr is the relative location of B with

respect to A. Three sets of tri-axial measurements ( , ,A B Ca a ar r r) should be sufficient to provide six

equations whose unknowns are the components of αr and ωr . However, it has been found4 that such a system of six equations with six unknowns is sparse and often singular. As a result, when using tri-axial accelerometers, the minimum number required to uniquely determine the rigid body acceleration of an object, is four.

A simpler approach consists of six acceleration measurements (e.g. from two tri-axial sensors A, B) and three orthogonal angular rate sensors to measure the components of the angular velocityωr . This is the approach that will be used in the USLI competition. A straightforward approach to experimentally characterize slosh events is presented here. It is based on the simultaneous measurement of the rigid body acceleration of the sloshing tank as described above, along with simultaneous capture of the shape and velocity of the sloshing surface based on an array of orthogonal CCD cameras synchronized by a hardware trigger. An additional setup uses manufactured stereo vision cameras coupled with customized software to produce a three-dimensional surface. Image flow techniques can also be used to estimate the

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“velocity” or flow of a given image, in a similar sense that 1-D position measurements can be used to estimate the 1-D velocity of a moving target.

A set of scaled liquid tanks has been designed to mimic maneuvers such as those in upper-stage rocket propulsion systems, or liquid fuel and water storage tanks onboard space stations. The tanks are modular and are capable of housing various candidate fluids. Several basic tank shapes have been selected, including spherical and cylindrical (Figure 17).

Figure 17: (left) Tank geometries considered for this study including flat bottoms (A), concave bottom shape (B), convex bottom shape (C). (right) a fabricated tank component

The measurement of the liquid location due to various base motion patterns will be accomplished by a vision-based system, based on a set of three orthogonal cameras, currently being developed at Florida Tech.

A challenging and novel aspect of the proposed research is the development of an adequate instrumentation framework for measurement and characterization of the slosh dynamics.

The experimental characterization of slosh dynamics is a very challenging problem since measurements need to be taken in multiple points inside of the tank. In the FLEVO experiment, platinum wire rings embedded in the tank walls were used to measure capacitance at 270 locations throughout the tank - these sensors provide liquid depth information over the location with a maximum resolution of 30 mm (Vreeburg and Chato, 2000). Miniature thermal dispersion meters were used to measure velocity at 10 locations within the tank. Other slosh studies have been based on drop tests with accelerometers attached to the tank's walls (Hubert, 2004), although the information that can be retrieved from such experiment is limited to the total force exerted by the liquid mass on the tank's walls.

A vision-based system to experimentally characterize slosh dynamics will be developed. At the moment, we are using one camera attached to the tank's frame, as shown in Figure 10. The proposed system uses an array of three cameras mounted orthogonally to a frame attached to the slosh tank. The proposed system will use stereo-vision techniques to estimate the morphology of the liquid’s surface during slosh. Figure 18 shows the slosh dynamics experiment that will be used for ground studies (Earth gravity levels) and that is in the process of being prepared for low-gravity flight conditions. The liquid tank can be excited over a wide range of 1 and 2 degree of freedom motions. Some examples of photos taken of sloshing events are shown in Figure 19.

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Figure 18: Slosh dynamics experiment mounted on a one degree of freedom motion table

Figure 19: Slosh images acquired from 1-DOF lateral excitation motion

A wide range of slosh amplitudes and frequencies has been mapped for various tank geometries, fill levels, and liquid types (different density) under both transient and steady-state acceleration conditions. A selected set of these cases has also been under investigation using computational fluid dynamics and will form the test matrix for the cases conducted on the low-gravity flight experiment (discussed in the subsequent section).

Florida Tech has developed a wall-wetting measurement technique for slosh experiments based on edge detection. Examples are show in Figure 20-Figure 24.

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Step

1

Figure 20: A 30 frame per second image of the sloshing motion within the tank is taken. The camera is fixed to the frame of the tank walls, such that the relative motion between the fluid and the container is captured.

Step

2

Figure 21: The image is converted to binary.

Step

3

Figure 22: A reference background image, with no liquid and with no motion, is subtracted and only the binary image of the liquid within the tank remains.

Step

4

Figure 23: The image processing software operates on the binary image shown in Figure 23 and locates all fluid edges and fills in the area.

Step

5

Figure 24: Data analysis and plot of liquid-gas edge detection. Data analysis provides total surface area being wetted by liquid, percentage of tank internal surface wetted by liquid, maximum liquid height along tank wall, all as a function of time.

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The image processing technique is capable of determining surface wetting and fluid distributions even under large amplitude nonlinear slosh conditions.

The following section outlines the results from the third flight test that Florida Tech slosh dynamics group conducted in September 2008 aboard the Zero-G 727. The third test consisted of two slightly modified assemblies from the second flight test. One will carry two instrumented tanks with three orthogonal cameras. The second will carry two tanks, stereo vision cameras, accelerometers, and gyros. The first flight test of the campaign consisted of parabolic trajectories that were conducted in a Piper Warrior and were used to understand the slosh behavior relative to aircraft flight path. The second flight was on board C-9B aircraft. This test allowed the team to become familiar with flight logistics and to test the sensors, mounting assemblies and cameras. Flight data was good, but the original accelerometer configuration used one set of tri-axial sensors. The full motion profile was not captured. Non-uniform lighting conditions were also noted and procedures for correcting it were put in place. This section will also discuss future plans for motion experiments and techniques for data processing.

In an aircraft-induced reduced gravity environment fluid slosh is caused by the flight path, more specifically, the change in acceleration associated with the inflection point during the transition from climb to nose over. Constant accelerations establish stable surfaces normal to the resultant acceleration vector. It only when changes in acceleration occur that the surface becomes dynamic.

Figure 25: Typical acceleration time histories during zero-g parabollic maneuvers (z-direction) for small and large aircraft testing

A partially filled tank with liquid was bolted to the air frame. The tanks are constructed using plates,

domes, and cylinders of cast acrylic, glued using IPS acrylic adhesive7. Each tank consists of a 9.5” cylindrical section, with a 12” wide, 0.5” thick flange at each end of the cylinder, and the option of a 12” flat plate or 10” dome with 12” flange to cap off each end of the tank. Using two domes to form a 10” sphere with 12” flange is also an option. A series of 0.25” diameter holes around the flanges allow the tank to be securely mounted during the experiment.

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The first instrumented tank uses 14 probes that are generally used for recognizing liquid level. They were placed in an arrangement to gain the most information from the aircraft induced slosh. They operate on recognizing the change of heat transfer due to the presence of the liquid. Indicating that liquid is present. The probes are cone shaped and extend approximately 0.5” into the flow. Exact geometry tanks are also tested to compare the differences in flow due to the obstructions. The second instrumented tank uses one probe attached to cryotracker tape located in the center of the tank.

The static flight experiment seen in fig. 2 used three orthogonal single axis strain gauge DC accelerometers, and three orthogonal 24-bit RGB color CCD cameras. Cameras and sensors were externally triggered for each maneuver. The cameras were set to acquire each frame when triggered by an external TTL source provided by a signal generator. The first trigger starts the analog data acquisition, and successive trigger pulses at equal time steps synchronize image acquisition at a fixed sampling rate on all three cameras. Each set of three simultaneous images was stored to hard disk before the next trigger pulse is sent, which limited the frame rate to 4.2 frames/sec. A future implementation based on FireWire cameras will allow frame rates up to 40 frames/sec, synchronized by the FireWire bus.

Figure 26. Experimental setup for recording sloshing events in microgravity. Pill-shaped tank.

Accelerometer data shows the three phases of a typical “zero-g” parabolic maneuver. Due to sensor

saturation former Piper flight test data was expanded to the same time scale for maneuver characterization. Entry to the parabola takes place from 0 to approx. 40 seconds. From 0-20 seconds the aircraft is in “steady level flight.” In the plot shown, the “zero-g” region is from 45 to 67 seconds, and relies

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on the pilot’s manual control of the aircraft’s acceleration. It is important to note that the presented “zero-g” is in fact constantly fluctuating within three error bands. Transitional areas with large jerks govern the resulting motion of the liquid. The capillary and gravitational forces are dominated by the liquid motion established by the inertial forces inherent to the climbing maneuver. Thus the climbing portion of the parabolic trajectory creates a slosh wave that is dependent on the pilot’s control of maneuver entry and nose over. Small yaw and roll perturbations cause differences among successive data sets. These perturbations are due to flight conditions and pilot control. The 1.8 g climb phase settles the liquid at the bottom of the tank just prior to the slosh wave generated by the aircraft’s pitch deceleration. It is clear that this environment is unlike a drop tower, which minimizes lateral forces. In that case, purely capillary forces dominate as those shown in Zarm’s drop tower tests. Without lateral influence the liquid surface attains a spherical profile. This can be shown in Zarm reports. It is also shown in CFD results obtained using the dynamic mesh approach. CFD results are shown in subsequent sections.

Figure 27. Experimental setup for recording sloshing events. Instrumentation and vision Systems.

NASA’s Zero-g aircraft creates approximately 20 second intervals of “zero-g” in a coordinate system attached to the plane, by flying in parabolas as shown in flight documents6. To create a “zero-g” interval, the aircraft must first climb and then drop along the falling edge of a parabolic trajectory. In fig. 4 one can see that the experiment experiences this aircraft induced inertia wave. In the CFD simulation in fig. 4 the same inertia wave is created by applying a pitching acceleration profile coupled with a “zero-G” region at the apex of the maneuver. Differences occur primarily due to the camera perspective and the fact that the CFD solution is based upon an arbitrary 45 degree rotation just prior to a “zero-g” regime. It was an initial example to prove that the dominate forces that initiate slosh are due to the aircraft pitching motions. That is why it is essential to not only capture the acceleration data but rotational data as well.

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Figure 28. Slosh event during microgravity test, recorded by an array of orthogonal cameras.

Figure 29. Slosh event at various stages of a microgravity test.

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Figure 30. Slosh wave caused by the climbing maneuver during a zero-gravity parabolic trajectory.

It would be ideal to start a “zero-g” experiment with the liquid as close as possible to stationary/parallel

to the bottom of the tank, the tank must be moved in a way that compensates for the plane’s acceleration during climb. Slosh compensation techniques based on real-time acceleration feedback have been described in robotic applications. Ideal compensation of the 2-D acceleration in the vertical plane (pitch) as the aircraft climbs would require control of two axes of motion, namely the pitch of the tank relative to the aircraft’s frame, and a translation along the aircraft’s main axis.

A control system that compensates the pitch acceleration of the plane to minimize the slosh wave created by the plane’s climb maneuver can be built based on a linear actuator that rotates the slosh tank relative to a pivot point fixed on the aircraft’s frame. The instantaneous acceleration of the tank can be calculated from real-time accelerometer measurements as described above, and the pitch compensation system is based on applying the acceleration necessary so that the total acceleration vector is perpendicular to the bottom of the tank at any point in time. A numerical verification of this concept can be achieved by generating a CFD slosh simulation with an arbitrary angular velocity and acceleration applied to the tank in the pitch plane. If the same and opposite angular rates were applied to the gravity vector on the CFD, minimal or no slosh should occur.

The collected data will be the basis for comparisons with state-of-the-art computational fluid dynamics (CFD) models, run with the same boundary conditions and acceleration profile of the rocket. A back-to-back set of experiments and numerical simulations will become available, which can be further used to develop reduced order analytical models to be used in a wide array of low-gravity fluid dynamics predictive tools, such as in the estimation of damping coefficient of the sloshing fluid. To date there are no mathematical models of liquid sloshing motion that have been benchmarked against low-gravity experiments. The collected data and associated numerical tools and models will help develop a fundamental understanding of the dynamics of any liquid in a low-gravity environment. These models can then be coupled with full 6 degree-of-freedom models of spacecraft motion to make realistic prediction of internal propellant slosh behavior on vehicle dynamics. Current Research at Florida Tech: Numerical Modeling

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In parallel to the significant amount of experimental research that is being conducted at Florida Tech, a computational fluid dynamics (CFD) effort is also in place. The goal of this effort is to develop computational tools and models which are capable of simulating the interaction of propellant slosh with the dynamic behavior of a spacecraft. There are several approaches in computational fluid dynamics to simulate liquid sloshing. The most popular approach is to solve the Navier-Stokes equations for moving free surfaces by finite difference or finite volume discretization schemes4-6. Although both interface tracking and capturing methods are used to find free-surface solutions, it is generally accepted that simulation of truly general free-surface motion (as found in slosh) can be best attained by interface capture methods, such as the volume of fluid (VOF) methods in FLUENT 6.3.

Three ways to simulate liquid slosh in FLUENT 6.3 have been identified, momentum attachment, dynamic mesh, and changing acceleration field. Each method has its specific advantages and disadvantages as will be discussed. Three simple case studies are run to directly compare the three tools, and to compare ground tests directly with CFD. In the first example, a 9.5” by 10” tank begins at a constant velocity of 1m/s for 0.5 seconds and decelerates at 1m/s2 to 0 and is allowed to settle for an additional second to compare settling dynamics. The second case is a drop tower test that uses both dynamic mesh and acceleration attachment. This was done to compare results to Zarm’s drop tower tests.29,102,33 The third is a linear translation case that was run to benchmark the dynamic mesh directly against experiment and to mark it as the best choice for continued study.

The three methods are described and compared for flexibility and computational speed. Momentum attachment is based on approximating the swaying motion of the partially filled tank as a non-stationary force using a user defined function (UDF). The tank motion can be generally expressed as:

2 sin( )sway mF a tρ ω ω= (2) where a is the tank displacement amplitude, ρm is the liquid density and ω is the swaying force frequency. After compiled in FLUENT, the UDF uses the DEFINE_SOURCE macro to indicate the unit force in the horizontal momentum equation, and attaches the liquid phase using non-moving tank wall boundary conditions. Recent examples of CFD slosh simulation using similar methods are shown in Hadzic et al.18 , Aliabadi et al.38, Standing et al.41, and Rhee 39. The momentum attachment method described above is especially suitable for investigating force feedback acting on the tank wall due to liquid sloshing.

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Figure 31: Comparison of developing slosh CFD tools used to simulate slosh on a 2D partially filled tank.

A second method, called Dynamic Mesh Model (DMM), is proposed. DMM is capable of simulating

both the liquid slosh as well as estimating the dynamic force exerted by the sloshing fluid on the tank. DMM is used to model flows where the shape of the domain is changing with time due to motion of the domain boundaries. To simulate slosh using DMM, the tank’s motion can be arbitrarily defined by the linear and angular velocities of the tank in time, so that the slosh force feedback can be ignored. The tanks center of mass moves along a predefined trajectory of its, unaffected by the liquid’s motion. In order for the DMM to move along its predefined trajectory the grid is regenerated at each time step, this uses extra computational resources.

In the third method the gravitational operating conditions are changed in steps to mimic the relative accelerations due to linear translation. The vertical acceleration during the simulation remains constant at 9.81m/s2, during the deceleration region the linear acceleration is changed to 1m/s2 and subsequently set to zero at 1.5 seconds The fluid is allowed to settle under its own influence at 9.81m/s2. Like the DMM acceleration patching can simulate both the liquid slosh and the dynamic forces exerted by the fluid on the wall. It also allows for fluid reflections whereas the momentum attachment cannot. The tank is stationary during the simulation so mesh resolution can be increased without concern for errors due to grid regeneration.

Figure 32 shows important qualities of each case the primary regions of interest are between t=1.5-2 seconds At t=1.5 seconds the maximum wave for all cases is established at the end of the deceleration region. One must be careful when looking at results it seems that the 2.5 second images are all identical, but at 2.0 seconds the momentum attachment case is clearly different. In the momentum case the forcing function creates a less than maximum possible slosh wave with the less potential energy relative to the DMM and acceleration patch. In the other two cases fluid motion is controlled by external forces and each cell does not have uniform velocity during deceleration. This creates a situation with more potential energy due to increased wave height. During the settling phase shown at t=2.0 seconds the momentum attachment slosh is returning to the opposite side of the tank before the other cases due to the reduced maximum wave height. The DMM and acceleration cases are still experiencing the transition from maximum potential to kinetic at 2.0 seconds.

The DMM unlike the momentum and acceleration attachments does not need to be preprocessed for input. The entire computational domain experiences realistic dynamics. During linear translation the differences are not seen; however, when rotations are introduced the acceleration field is a position

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dependent gradient and further development of CFD parameters must be completed to include the gradient in computations. In terms of the momentum attachment the velocity of individual fluid cells are forced at each time step. This cause a situation where vertical velocities and reflections are overcome be the defining function. This can be advantageous when calculating the worst case force feedback on the wall, but in our study the full time scale dynamics are relevant and fluctuations of fluid velocities are needed to portray the most accurate slosh. This velocity restriction sets up fluid for incorrect settling regions.

The establishment of computational grid is a critical part of attaining quality results. In the current simulation, O-grid meshing method was introduced to generate the numerical domain grid. This method could not only provide a simple control for creating grid on specific areas but also reduce the unstructured mesh to a minimum amount. The mesh near the tank wall region was refined because this specific region is the boundary where interaction between the tank wall and the liquid phase is crucial. In this region, detailed physics for capillary action must be captured when the grid is coarse the fluid is dominated by the cells that lie strictly along the boundary.

To investigate the affects of grid resolution, dependence studies were conducted on both the ground test and drop test. Two numerical domains with different grid resolution were created for both. The numerical domain for the ground test consisted of 500,000 structured and unstructured mesh elements. With the same grid generation strategy, the mesh elements in the other domain with the coarse mesh reduced the element count to 50,000. The simulation cases were running on a compute node of a 48-node IBM system. The configuration of each compute node has x330 series with 1 PIII processor with 512 MB SDRAM RDIMM2. The main problem is that the server is shared by the university and has high traffic. The computation time of the fine grid resolution case was over 168 hours (7 days) while the coarse mesh case had only 24 hours to get results. Figure 33 shows the comparison of three time steps. They are taken at 1.6 seconds (0.1 second after the tank stopped moving), 3.0 seconds and 9.0 seconds. The coarse mesh fails to capture the defined ripples inherent to the the experiment and fine CFD. It does capture the relative liquid angle and initial wave wake, but the cell size is too large to keep the solution from having a stepped surface.

To demonstrate the influence of mesh resolution, a example of dropping tower test applied a linear acceleration of 9.81m/s2 to simulate a free fall. In this case (Figure 34) the coarse mesh resolution of 1,698 structured cells affected the solution dramatically. After complications with trying the DMM method to compare with the acceleration patch for the drop case, the mesh resolution was increased to 59,650 structured cells. After this adjustment the solution yielded correct results. It took 2 hours to compute the coarse solution and 10 hours to compute the fine mesh. Also instead of tricking the solver by changing the gravity field each time step, the domain experiences the same motion as a real free fall making the DMM approach more reliable when it come to complex motions. As previously stated by controlling the motion at the boundaries the fluid is free to interact with itself. This is the best solution for flexible input, any fully characterized motion profile can be used.

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Figure 35: DMM Drop tower test mesh density improvement. Both simulations are at 4.0 sec

Figure 36: Ground Test Direct Comparison with CFD from top to bottom: 1.6s, 3.0s, and 9.0s

A test was run to compare linear translation under the regular gravity (1G) and CFD with an exact

motion profile. The experimental and numerical simulation results are shown in Figure 19. For this study, a cylinder-like tank numerical domain was selected for computation. The domain has a

diameter of 9.50 inches and a total height of 14.75 inches (from the top surface of the cylindrical part to the bottom of the hemisphere part).

The liquid phase was partially attached into the domain with liquid level of 6.75 inches. By using dynamic mesh method, the whole tank domain accelerated from 0 to 0.24 m/s in 1.5 seconds and then completely stopped and allowed to settle until 10 seconds. In order to compare the numerical results with experimental results, a test of a moving tank was conducted. The tank has the same dimensions and motion profile as which were used in the numerical simulation cases. Results are shown in fig. 8 and are for time steps 1.6, 3.0, and 9.0 seconds. It is important to note key feature that are similar in the time steps. For the first time step it is 0.1 seconds after the tank decelerates to a stop. This is a point of maximum wave height. Wall adhesion can be detected by the raised ring on both the high and low sides of the tank. At 3.0 seconds one can see that the levels and the symmetric ripple formation are in agreement between simulation and experiment. Finally at 9.0 seconds after some time damping the solution is still accurate and shows a dimple formation that has a reflected wave captured in the center. This double sided dimple is a recognizable feature in both the experiment and simulation.

Initial CFD simulations of low gravity slosh in two and three dimensions were used for qualitative comparisons between first-order slosh models such as those presented in Ref. 1, and results from CFD (Figure 37.b). It is important to note that the walls of the SWRI1 tank in 8.a are elliptical and the fluid depth is approximately equal to its static radius at the minimum liquid/gas interface. The tank in 8.b is a cylinder with 9.5” radius and 24” height with a square bottom. In addition, the contact angle of the CFD case is 40 degrees, whereas in the SWRI report1 no prior knowledge of contact angle is known. The main visible difference is in the smoothness of the liquid/gas interface on the left hand side of the frame (Figure 38.a) compared to the small wave forming a bump on the left wall of Figure 39.b. This difference is due to the square bottom of the tank. Identifiable similarities are the constant center, raised sloping right hand side and left dipped sided with a parabolic shaped rising slope. In Figure 40.c and 8.d the steady-state configurations are shown - in low gravity, inertial and capillary forces dominate and extensive wall wetting is expected. In the CFD, the tank started in a micro gravity environment of 0.001go and was allowed to establish the equilibrium contact angle before a linear translation of the tank was introduced. As expected, the bulk of the fluid accumulates along the walls and corners in the steady state.

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Figure 41: Comparison of zero-g slosh, as predicted by first-order model to CFD

Figure 42: CFD Drop Test at 6.0 seconds

The CFD allows the use of an arbitrary acceleration time history for the tank, as it generates a solution for the liquid distribution at each time step.

In the case of the drop tower CFD, the fluid established its contact wall angle which was set to zero. Only vertical accelerations changed, capillary action was the dominant factor for establishing the surface. A mostly spherical surface results (Figure 43) and is in excellent agreement with typical results from drop tower tests7,30,34.

The examples prove that the DMM and the acceleration patch model are the best suited methods for this study. Combined with a VOF multiphase model, they will provide an accurate parallel platform for numerical simulations in future studies.

The numerical models can be used to predict fluid behavior, droplet break-up, surface tension effects, gas-liquid phase changes, etc. An example of volume of fluid simulation is shown in Figure 44. The tank is cylindrical with elliptical end caps and is filled 25% by volume with liquid; the rest being gas. A slosh baffle, is also included in this simulation, and is modeled as a circular ring. Initially the liquid occupies a lower quadrant of the tank, as shown in the initial condition, and then at the start of the simulation the fluid is allowed to flow.

Initi

al C

ondi

tion

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1 se

cond

2 se

cond

s

3 se

cond

s

4 se

cond

s

5 se

cond

s

Slosh motion at Earth’s surface gravity

level, g = 9.8 m/s2. Slosh motion at 1/100th of Earth’s surface

gravity level, g = 0.098 m/s2. Figure 44: Numerical modeling of fluid slosh motion at Earth gravity level (g = 9.8 m/s2) and at

1/100th of Earth gravity level (g = 0.098 m/s2). In the left column, the fluid flows in Earth’s gravitational field. The right column has the same initial condition, but the local acceleration is 1/100th of that of that on the surface of the Earth - each horizontal row in the figure is at the same time, and the differences between flowing liquids depending on the local gravitational acceleration is clearly seen.

Figure 45 shows two additional simulation results where the tank is set into an arbitrary acceleration profile, and the fluid within the volume responds accordingly. The picture on the right shows a slosh simulation where the internal geometry of the tank includes 3 slosh baffles to attempt to damp some of the liquid motion.

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Figure 45: Examples of CFD studies of slosh in a cylindrical container.

The computational simulations will be compared directly against those conditions tested on the

aircraft flights. The exact geometry of the tanks, aircraft acceleration and tank motion profile, and fluid properties will be inputs into the CFD simulations. The simulations will then be conducted under the same acceleration conditions as the flights and the simulations results can be compared directly against those of the numerical model. These comparisons will be performed for both the ground-based testing (Earth gravity levels) and the low-gravity flight tests. These experiments will provide the first opportunity to benchmark numerical codes at low gravity levels.

Figure 46. CFD simulation of a slosh event due to acceleration of the tank relative to an inertial frame.

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Appendix B: Senior Design Safety Plan Requirement 1.1 Safety Plan Report Overview The Senior Design Safety Plan Report (SDSPR) must include the following elements (the report may not be submitted as a bulleted list): 1.1.1 Project General Description (PDG) 1.1.2. Hazard Analysis (HA) 1.1.3. Human Safety Analysis (HSA) 1.1.4. Failure Modes and Effects Analysis (FMEA) 2.0 Project General Description Each SDSPR must include up to 2 pages which describe the project and the suggested methodology to accomplish the work. This should also include any testing required during or leading up to the final work. 3.0 Hazard Analysis (HA) The HA should include the following elements: 3.1 Description of any materials or conditions which will be used/implemented in the course of the work that contain an inherent hazard. This description should also include a detailed analysis of what Hazard the described material or process presents. 3.2 Environmental Impact Analysis. If applicable, Material Safety Data Sheets (MSDS) should be obtained for all materials to be used in any phase of the project. These should be included in the HA and proper storage and disposal of these substances should be discussed. Also a discussion of any regulatory agency (such as EPA or FAA) policy, which has jurisdiction over the materials/procedures you will use. 4.0 Human Safety Analysis (HSA) 4.1 Personal Protection Equipment (PPE). During the course of your HA you will identify types of PPE which are required for your work. This section must include what PPE are to be used under what conditions/situations you will encounter during your work. 4.2 General Work Safety. The Occupational Safety and Health Administration (OSHA) has numerous procedures that have been identified to protect people during most any working circumstance. This section must also include reference to the appropriate OSHA statute that affects your project work and a description of how you will comply. 4.3 University Insurance. If your project will involve travel, involve working with an outside business or school, and/or involve working with anything that could endanger you or the public (i.e. driving a student built vehicle, flying or firing a rocket, etc.) you must contact Wanda Givens wgivens@fit.edu to discuss if the Florida Tech Insurance

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coves such activities. If no, then you must look into obtaining the appropriate additional insurance to proceed. 5.0 Failure Modes and Effects Analysis (FMEA) 5.1 Identification of any reasonable way in which your project or testing could fail or go wrong. Some like to describe this as a set of best guesses on how Murphy’s Law could impact your work. 5.2 Once you have identified potential failure modes you must include what will or could occur as a result of such a failure 5.3 Failure Mitigation. This section should include a description of what steps you are taking to avoid the kinds of failure modes you have identified.

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Appendix C: MSDS Sheets G10FR4

61

62

63

64

SIG5 Minute Epoxy

65

66

67

68

69

70

71

72

73

74

75

76

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Appendix D: Payload Specifications The following table is for the accelerometers parameters. It shows the minimum and maximum allowable parameters. The most important parameters here are the sensor measurement range and the sensitivity. As you can see we will be able to measure up to 18 g’s. The sensitivity will be used to convert the voltages we collect to G-forces measured.

Table 1: Accelerometer Data Sheet (Sparkfun)

The following table is for the sing axis gyros and contains the minimum and maximum allowable values for all the parameters. In the case of the gyros, the most important parameters are the Dynamic Range and the Scale Factor. Using a similar equation as the one above, angular velocity can be calculated.

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Table 2: Gyro Data Sheet (Sparkfun)

Table 3 shows some of the characteristics of the PIC microcontroller we are using. We chose this processor because of its 10-bit Analog-to-Digital Module. As you can see in the table below, our MCU has 8 input channels. Another feature to note is the maximum operating frequency of 40 MHz. This feature allows us to have a sampling rate close to 1 KHz per channel. Our EEPROM uses SPI in order to communicate to our MCU. This is another reason we picked the PIC18F452 since it also has a MSSP serial communication module.

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Table 3: PIC18f452 Data Sheet (Sparkfun)

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CMOS IR Camera Module - 640x480

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SanDisk V-Mate SDVM1-A-A30 Video Memory Card Recorder General Features: Enjoy your video content anywhere Capture and record video content from analog source directly onto flash memory card Schedule recording Playback video content Easy to set up Plug and Play No PC required Compatible with most flash memory media formats. Front Panel Features: Power LED Recording Status LED Memory card slots Rear Panel Features: ON/Off 5V DC plug 3.5mm input for composite video, audio left & right 3.5mm output for composite video, audio left & right USB mini B IR emitter (connection to set-top box or VCR) Supported Memory Cards: Secure Digital (SD) SDHC (Secure Digital High Capacity) MiniSDHC MicroSD MicroSDHC MultiMediaCard (MMC) RS-MMC (Reduced Size MMC) MMCPlus MMCmobile Memory Stick (MS) Memory Stick Pro Memory Stick Duo Memory Stick Pro Duo Specifications: Codecs: MPEG-4 Simple Profile H.263 Formats: MP4, 3GP, 3G2 Resolution: Up to 640 x 480 pixels Frame Rate: 30 fps Record up to 3.5 hours of high quality video per gigabyte Unit Dimensions: 5.1 x 2.6 x 0.8-inches (L x W x H)

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Appendix E: Payload Calculations Pin Shear: Assuming: 4.5lbs at 100 g’s on the pins The pin at each section is in double shear Diameter of the pins is 0.164 in (we are using an 8-32 bolt) (McMaster Carr) The bolts will be made out of Al6061 T6 Shear strength (SS) of the bolts is 30 000 psi (Matweb)

Lessons learned: The bolts will not be the weakest link in the payload section and are designed to carry a load much greater than they will ever see. Pin Tensile Stress: Assuming: 4.5lbs at 100 g’s on the pins Each pin holds half the total weight of the payload Minor diameter of the pins is 0.1248 in (we are using an 8-32 bolt) (Engineer's Handbook) The bolts will be made out of Al6061 T6 Tensile Strength (TS) of the bolts is 40 000 psi (Matweb)

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Lessons learned: The bolts will not fail in tension even under 100g acceleration of the payload section. Pin Tread Pullout Stress: Assuming: 4.5lbs at 100 g’s on the pins Each pin holds half the total weight of the payload Minor diameter (Kn) of the pins is 0.1248 in (we are using an 8-32 bolt) (Engineer's Handbook) Number of threads per inch (n) is 32 for an 8-32 bolt (Engineer's Handbook) Minimum pitch diameter of external thread (Es) is 0.1399 (Engineer's Handbook) The bolts will be made out of Al6061 T6 Shear Strength (SS) of the bolts is 30 000 psi (Matweb) There will be two nuts on each end of the rod to double the fastener thread engagement

(Le) but for the purpose of worst-case scenario, the Le will be used for the shear area calculation

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Lessons learned: In the worst-case scenario of one of the nuts falling off, the bolt will not fail under a 100g load. Column Buckling: Assuming: 4.5lbs at 100 g’s on the column Clamped at both ends (Le= ½ L) Each column is made of G10/FR4 fiberglass E=2 200 000 psi (Professional Plastics) L=6.5 in w=3.70 in (width of each column)

Substituting in the moment of inertia I and rearranging,

Lessons learned: The G10/FR4 chosen will be 0.187 in thick, which is the closest commercially available thickness for a factor of safety of 2. Therefore, buckling will not become a problem. G10/FR4 Bearing Stress: Assuming: 4.5lbs at 100 g’s on the column from the bolts The column is made of G10/FR4 fiberglass Thickness of fiberglass=0.187 in Diameter of the bolt is 0.164 in (a #8 bolt) (McMaster Carr) Tensile strength (TS) of the G10/FR4 is 35 000 psi (Professional Plastics)

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Lessons learned: Even in the worst case scenario, the bolts will not rip through the fiberglass that is holding the payload. Tube Buckling: Assuming: 30lbs at 10 g’s on the tube Worst Case Scenario of one fixed end and one free end (Le=2L) The tube is made of G10 fiberglass E=2 200 000 psi (Professional Plastics) L=132 in (This is the entire length of the payload tubing) do=4.03 in (This is the outer diameter of the tubing) di=3.91 in (This is the inner diameter of the tubing)

Lessons learned: Buckling of the payload tube is of major concern if the payload tube were empty. Because the tube will be filled with the electronics, the tube will not really buckle at these loads.

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Appendix F: Rocket Performance Calculations To meet the criteria of the USLI competition, the vehicle must successfully lift a scientific payload to an altitude of exactly one mile, and return it safely. Initial designs of the vehicle have been made to complete this objective, while leaving the option for design changes to alter vehicle performance. Table 4: Predicted weights of the major vehicle components Item Manufacturer Weight (lb) Nose Cone Hawk Mountain 1.00Payload Body Tube Hawk Mountain 2.10Payload Florida Tech 10.00Forward Coupler Tube Hawk Mountain 0.67Upper Bulkhead Florida Tech 0.75Recovery Body Tube Hawk Mountain 3.00Main Parachute RocketMan 1.67Avionics Coupler Tube Hawk Mountain 0.47Avionics Forward Outer Bulkhead Florida Tech 0.15Avionics Forward Inner Bulkhead Florida Tech 0.07Avionics 1.00Avionics Aft Inner Bulkhead Florida Tech 0.07Avionics Aft Bulkhead Florida Tech 0.15Drogue Parachute RocketMan 0.54Booster Body Tube Hawk Mountain 1.83Booster Coupler Tube Hawk Mountain 0.59Motor Mount Tube Hawk Mountain 1.25Forward Centering Ring Florida Tech 0.06Mid Centering Ring Florida Tech 0.06Aft Centering Ring Florida Tech 0.06Fins Florida Tech 1.01 Total 26.49

Vehicle Parameters:

Figure 47: Basic vehicle parameters including layout, center of gravity, center of pressure, static margin, overall dimensions, and liftoff weight.

CG CP

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Motor thrust data used in initial simulations:

Figure 48: Motor thrust curve used in the preliminary simulations Table 5: Basic motor specifications used in initial simulations Specification Value Diameter 76mm Length 498mm Case Mass 1.724kg Propellant Mass 1.814kg Average Thrust 883.8N Peak Thrust 1123.0N Total Impulse 3533.5Ns Burn Time 4.0s

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Initial Flight Simulations

Figure 49: Simulated acceleration data of the vehicle at the predicted weight of 25.6lb

Figure 50: Simulated velocity data of the vehicle at the predicted weight of 25.6lbs

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Figure 51: Simulated altitude data of the vehicle at the predicted weight of 25.6lbs

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Figure 52: Simulated acceleration data of the vehicle at the maximum weight of 30lbs

Figure 53: Simulated velocity data of the vehicle at the maximum weight of 30lbs

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Figure 54: Simulated altitude data of the vehicle at the maximum weight of 30lbs

Figure 55: Perfectflite Data from test launch

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Appendix G: Launch Test Directive 1. GENERAL The following sections will provide an overview of the procedures and methods involved in preparing Panther II-Heavy for launch April 18th for the University Student Launch Initiative (USLI). 1.1 Mission Objective The primary objective of this mission is to successfully launch Florida Tech’s student built rocket, Panther II-Heavy (furthermore noted as the Team), at the USLI competition April 18th in Huntsville, Alabama. The vehicle configuration consists of an LOKI L930 motor with a complex science payload. The payload will consist of the two sensor boards, the slosh tank, and a VCR to record the slosh tank during flight. The sensor boards include 5 sensors to measure the accelerations and rotations that the rocket will experience during flight. The slosh tank is the science portion of the experiment. This will be recorded during flight to see how the fluid reacts during the flight and then this data will be used to create an accurate CFD model to better understand how fluids act in low gravity. 1.2 Mission Description This mission will include a LOKI L930 motor in a 14-foot length, 4-inch diameter G10 fiberglass tube. The vehicle will be launched vertically from a 20-foot launch rail that the Team will provide at day of launch. The Launch Director will initiate a countdown from the Bragg Farm Launch location and provide ignition voltages to the LOKI L930, which will boost the vehicle to a goal apogee of 5280 feet. 1.3 Launch Preparations For the day of launch the Team will carefully and meticulously test all electronics, wiring, fittings, U-bolts, and additional payload components. Once that is complete the rocket will be assembled and all shear pins and other outer bolts will be tightened. Finally the motor assembly will be fitted to the end and locked in place. While these procedures for rocket setup are being completed, other members will be adjusting the launch rail and preparing for the rocket to be attached. The Team will then carry the rocket to the rail and slide said rocket into place. The rail will be lifted and checked for alignment and finally the igniter will be inserted and secured. The wires for this igniter will then be run back to the Launch Director. The Launch Director will approve and implement any necessary deviations to this concept of operations. Launch personnel are protected to the greatest extent possible through pre-planned procedures which are carefully reviewed and approved. These procedures control the entire sequence of events before, during and after launch. The controlling procedure is the launch countdown, which determines who does

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what activity and when this activity will be performed. This procedure is carefully monitored to ensure that there are no deviations from this document. Personnel and equipment safety is provided by restricting these assets from areas where danger exists through implementation of the Launch Danger Area (LDA), described in this plan. Spectators will either be sheltered in the blockhouse or remain in designated areas outside the Launch Danger Area, greater than 300 ft from the launch rail. 1.4 Launch Criteria To accomplish the mission objectives, launch clearances are required from: a) Marshall Space Flight Center Personnel (MSFC) b) Launch Director (LD) c) Operations Safety Manager (OSM) 1.5 Panther II Heavy System The Panther II-Heavy is a one stage vehicle consisting of a solid-propellant LOKI L930 motor. This class of vehicle will provide the necessary acceleration to ensure payload security as well as achieve the goal altitude of 5280 feet. The components of the vehicle are described in the following sections. 1.6 Motor LOKI L930 This motor was manufactured by LOKI Research (herein described as LOKI L930). The motor will provide a slower acceleration in comparison to motors of its caliber. This slow acceleration is a staple of the LOKI L930 motors for those seeking to test “sensitive” payloads (sensitive referred here as having a limit of only 20 g-forces at any point during flight). 1.7 Launch Rail For the day of launch we will be providing our own 20 foot tall launch rail which shall be erected upon a trailer for stability of the rail. As shown in the provided reports, the launch rail satisfies the height requirement for the Panther II Heavy rocket. This requirement was derived from the need for a slow launching rocket to maintain dynamic stability before leaving the end of the rail. A higher speed launch would prove detrimental to the avionics and flight recorders for the slosh experiment. 1.8 Logistics for Launch Outside of logistics set by MSFC for launch day, the Team will be keeping to strict guidelines, some of which mentioned above (in Launch Preparation), and others mentioned here: -Checklist for essential launch equipment including rocket components must be constructed - Prior to procedures taken in rocket assembly, communication of said actions will be made to professional personnel present. -Launch countdown will be loud and clear for the safety of all present.

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1.9 Launch Operations Team The Team will be responsible for all launch operation procedures. Additional professional personnel will be provided at MSFC request. 1.10 Photography Photographic documentation and video recordings will be provided by personnel within the Team and any additional personnel provided by MSFC. 1.11 Telecommunications If necessary, the telecommunications will be provided by MSFC. 1.12 Access Restrictions The only definitive restrictions pertaining to launch operations will be ground clearance of all students and professionals when any given rocket (especially Panther II) is set for launch. Depending on the motor this may be over 500 feet. 1.13 Launch Safety To maintain launch safety, the Team will make all actions clear to professional personnel present prior to set-up. All personnel will be required to stand at the defined safety zone once Panther II is set for launch. 1.14 Personnel Certifications 1.141 Formal launch team training and certification is required for the Launch Director (LD), Launch Engineer (LE), and Pad Supervisor (PS). Launch Director training encompasses all three positions; therefore certification as LD includes certification as LE and PS. 1.142 Certification as Launch Director consists of a formal training course and supervisory participation in five (5) launches. Certification as Launch Engineer and Pad Supervisor consists of a formal training course. Upon completion of these criteria, the students will receive a certificate verifying their certification status. 1.143 Ordnance/Explosive General Safety Training is required for personnel handling ordnance at Cape Canaveral. For the LD, LE and PS positions, this training is desired, but not mandatory. Bureau of Alcohol, Tobacco, Firearms, and Explosives (BATFE) Low Explosives User Permit (LEUP) and Range Safety Escort is required for non-Cape Ordnance Personnel ordnance handling. 1.15 Launch Team Certifications (certificates on file at FSA): Greg Peebles � Low Explosives User Permit, 2006 � Hazardous Materials Training, DOT, February 23, 1.16 Rocket Nominal Flight Data Figure 56 provides Panther II rocket nominal flight data when using a LOKI L930 motor launched from the 20-foot launch rail.

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Figure 56: Simulated launch with weight of 30 lbs. 1.17 Launch Vehicle Statistics LOKI L930 Weights (lb): Motor (with propellant): 9.259 Launch Weight: 36.9 Propellant: 4.140 Expended Motor: .5.119 Motor Performance: Total Impulse (N-sec): 3600 Action Time (sec): 3.0 Average Thrust (N): 1112.06 Vehicle Dimensions (in): Overall Vehicle Length: 168 Motor Length: 19-5/8 Motor Diameter: 2.99 2. GROUND SAFETY PLAN This Ground Safety Plan identifies the hazardous systems which exist on this vehicle, and defines the safety category for each hazardous system. Depending on the vehicle safety category, personnel restrictions may be imposed during various launch operations.

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2.1 Hazardous Systems The booster motor, with igniter, contains 0.529 pounds of propellant. 2.2 Launch Danger Area This section aims to define a set radial distance which personnel are considered in “harm’s way” once Panther II is set to launch. For now a tentative Launch Danger Area (LDA) will be set for a minimum of 500 feet. Corrections will be made once iterations with MSFC are completed. 2.3 Weather Constraints All weather constraints will be provided by MSFC, closer to launch day. Once determined, this report will adjust accordingly. 2.4 Safety Responsibilities To be added. 2.5 Launch Vehicle Pyrotechnic Information Two 9 gram avionics charges will be inserted for separation of essential compartments to release a Drogue chute followed by a parachute. The blast from these two charges will be minimal and all actions pertaining to these pyrotechnics will be advised by Mr. Greg Peebles. 2.6 Thrust Termination System This vehicle does not contain a thrust termination system. 2.7 Other Hazardous Systems and Requirements This vehicle contains no additional hazardous systems. 2.8 Radio Frequency (RF) Transmitters Launch team communications will be by FRS Radios. Frequency is 467.5625 MHz ±0.25 MHz (Channel 8 FRS radio). Power is 0.5 watts max. 2.9 Radioactive Materials There are no radioactive materials on this vehicle. 2.10 Explosives Classifications The classification levels used to classify explosives are adopted from the U.S. Department of Defense document 6055.9-STD, "Ammunition and Explosive Safety Standards". These classification levels and their respective hazards are as follows: 1.1 Mass Detonating 1.2 Fragment Producing 1.3 Mass Fire 1.4 Moderate Fire: No Blast

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The following list includes all explosives on this vehicle and their explosive classifications: Booster motor, without igniter 1.4C Booster motor, with igniter 1.4C Avionics charges, 1.4A 2.11 Hazardous Circuit Classifications This vehicle contains no hazardous circuits. 2.12 Personnel Restrictions Active essential personnel will be permitted in the area only when necessary and with the Launch Director’s or Pad Supervisor's permission. 2.13 Circuit Approval All circuits utilized for this launch will be commercially bought in order to simplify the certification process. Once the circuits are all known, MSFC will approve the ones that fit to the criteria. 4. NOMINAL LAUNCH SEQUENCE OF EVENTS In order to provide a successful, professional, and safe launch, the Team will be providing a thorough list of procedures starting several minutes before launch, running through to several minutes post-launch.

Make a thorough run-through of Item checklist so that all parts are available.

Avionics bay tied with yellow fabric Charges attached

o Blue attached to 1 of 2 terminals o Red attached to 1 of 2 terminals o Check for wire security o Twisted in parallel o Total of two (2) charges attached

Kevlar rope rolled and avionics bay inserted Remove rocket stop from rail Secure fit avionics bay into rocket fuselage Insert motor Thread shock cord & attach to motor Secure motor bay (shear pins) Walk rocket to rail and secure Make 2nd stage of rail

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Raise rail and lock in place with large screw joints Igniter inserted Blue tape wrapped on end to secure igniter Continuity check (avionics) Retreat to ignition station and wait to launch Once rocket successfully returns to the ground send out “runners” Retrieve rocket in each separate section making sure not to add to any

potential damage Return back to Panther II tent for post-flight analysis

Additional procedures will need to be added once full slosh payload is inserted.

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Appendix H: Parts List Table 6: Airframe components Manufacture Item Quantity Cost Total Cost G-Wiz HCX 1 $235 $235 G-Wiz USB Interface 1 $35 $35 Ozark Aerospace

ARTS 2 1 $180 $180

Hawk Mountain

FNC-4.0-4 to 1

1 $55 $55

Rocketman 12ft Standard Parachute

1 $145 $145

Rocketman 4ft Pro EXP 1 $57 $57 Hawk Mountain

FT-3.00 0.5 (2.5ft) $18.75/ft $46.88

Hawk Mountain

FTEX2-3.91 3 (5ft each) $18.96/ft $284.40

Hawk Mountain

CT-3.91 (28in needed) $18.16 $42.36

Giant Leap Rocketry

Slider 4 1 $11.72 $11.72

Total $1092.36 Table 7: Payload Components

Qty Part # Description Unit Price Total Price

1 PGM-08278 JTAG USB OCD Tiny - Programmer/Debugger for ARM processors $51.95 $51.95

1 BOB-00544 Breakout Board for microSD Transflash $14.95 $14.95

2 PRT-00127 microSD Socket for Transflash $3.95 $7.90

2 COM-08163 Flash Memory - microSD 1GB $9.95 $19.90

4 SEN-00848 Accelerometer Breakout Board - ADXL321 +/-18g 29.95 $119.80

6 SEN-00394 Gyro Breakout Board - ADXRS401 - 75 degree/sec $64.95 $389.70

2 PRT-08483 Polymer Lithium Ion Batteries - 2000mAh JST Connector 16.95 $33.90

1 PRT-08293 LiPoly Fast Charger - 5-12V Input $19.95 $19.95

4 COM-00526 Voltage Regulator - 3.3V $1.95 $7.80

4 COM-00107 Voltage Regulator - 5V $1.25 $5.00

2 COM-00527 Voltage Regulator - Adjustable $1.95 $3.90

1 PRT-08290 5V DC to DC Step Up - VPack PCB $9.95 $9.95

1 WRL-00151 Transceiver nRF2401A with Trace Antenna $21.95 $21.95

1 SEN-07904 CMOS Camera Breakout Board $14.95 $14.95

1 SEN-00637 CMOS Camera Module - 640x480 $19.95 $19.95

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1 LCD-08600 Color LCD - Breakout Board $44.95 $44.95

2 LCD-00569 LCD $19.95 $39.90

2 BOB-08745 3.3 v level shifter $1.95 $4.90

1 LCD-00709 Basic 16x2 Character LCD FSTN $15.95 $15.95

1 BOB-00718 Breakout Board for FT232RL USB to Serial $14.95 $14.95

1 SEN-08773 CMOS IR Camera Module - 640x480 $34.95 $34.95

1 SEN-08739 CMOS Camera Module - 640x480 $31.95 $31.95

1 BOB-08401 Breakout Board for USB miniB $1.95 $1.95

4 COM-08581 Crystal SMD 24MHz $0.95 $3.80

8 COM-00104 Trim pot 10K $0.95 $7.60

1 WRL-08947 RF Link 4800bps Receiver - 315MHz $5.95 $5.95

1 WRL-00151 Transceiver nRF2401A with Trace Antenna $21.95 $21.95

2 BOB-00196 Breakout Board for RF-24G Transceiver $1.95 $3.90

2 PRT-08631 RCA Jack $0.95 $1.90

1 WRL-08945 RF Link Transmitter - 315MHz $3.95 $3.95

4 COM-00102 SPDT Mini Power Switch $1.50 $6.00

4 PRT-00587 USB miniB SMD Connector $1.50 $6.00

4 COM-00097 Mini Push Button Switch

$0.35 $1.40

4 COM-07844 Triple Output LED RGB - SMD $2.25 $9.00 Index & Account:

Order Total: $1002.50

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Appendix I: References 1. Abramson, H. N., \The Dynamic Behavior of Liquids in Moving Containersss," NASA-SP-106,

1966. 2. Armenio, V., 1997, “An Improved MAC Method (SIMAC) for Unsteady High Reynolds Free

Surface Flows,” Int. J. Numer. Methods Fluids, 24, pp. 185-214. 3. Arndt, T. and Dreyery, M.: "Damping Behavior of Sloshing Liquid in Laterally Excited

Cylindrical Propellant Vessels", 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 8 - 11 July 2007.

4. Berglund, M. D., et. al., The Boeing Delta IV Launch Vehicle – Pulse-Settling Approach for Second-Stage Hydrogen Propellant Management., Acta Astronautica, Volume 61, Issues 1-6, June-August 2007, Pages 416-424.

5. Bridges, T. J., 1981, “A Numerical Simulation of Large Amplitude Sloshing,” Pro. Third Int. Conf. on Numerical Ship Hydrodynamics, Basin d’Essais des Carenes, Paris, France, pp. IV-4-1-VI-4-15.

6. Celebi, M. S. and Akyilkdiz, H., 2002, “Nonlinear Modeling of Liquid Sloshing in a Moving Rectangular Tanks,” Ocean Eng., 29, pp. 1527-1553.

7. Dodge, F. T. et al.: "The new dynamic behavior of liquids in moving containers", Technical Report, Southwest Research Institute, San Antonio, TX, 2000.

8. Dodge, F. T.: "Engineering study of flexible baffles for slosh suppression", NASA Technical Report, Sep. 1, 1971.

9. Dodge, F. T.; Garza, L. R.: "Simulated low-gravity sloshing in spherical tanks and cylindrical tanks with inverted ellipsoidal bottoms", NASA Technical Report, Feb. 1, 1968

10. Enright, P. J. and Wong, E. C.: "Propellant slosh models for the Cassini spacecraft", Proc. AIAA/AAS Astrodynamics Conference, August 1994, pp. 186-195.

11. Grayson, G. D, Lopez, A., Chandler, F. O., Hastings, L.J., Tucker, S.P.: "Cryogenic Tank Modeling for the Saturn AS-203 Experiment", 42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 9-12 July 2006.

12. Haberbusch, M. S. et al.: "Flexible temperature sensing probe", US Patent No. 6,431,750. 13. http://science.nasa.gov/headlines/y2001/ast03apr_2.htm 14. http://www.boeing.com/defense-space/space/delta/delta4/d4h_demo/book14.html 15. http://www.swri.org/4org/d18/mechflu/fluiddyn/space.htm 16. http://www.flow3d.com/apps/aero/app_aero_main.html 17. http://www.globalsecurity.org/space/systems/xss.htm 18. I Hadzic, Frank Mallon, and M Peric., 2002 “Numerical Simulation of Sloshing,” Technical

report, Technische University Hamburg-Harburg. 19. Ibrahim, R., A., Liquid Sloshing Dynamics: Theory and Applications. Cambridge University

Press, © 2005. 20. International Launch Services, Inc. and Lockheed Martin Corporation, Atlas Launch System

Mission Planner’s Guide, p. A-20, International Launch Services, McLean, Virginia, 2004. 21. Kassions, A. C., and Prusa, J., 1990, “A Numerical Model for 3-D Viscous Sloshing in Moving

Containers,” Recent Advances and Applications in Computational Fluid Dynamics, O. Baysal, ed., ASME FED, New York, Vol. 103, pp. 75-86.

22. Kim, Y., 2001, “Numerical Simulation of Sloshing Flows with Impact Load,” Appl. Ocean. Res., 23, pp. 53-62.

23. Lamb, H., Hydrodynamics, Cambridge University Press, Cambridge, 7th ed., 1945. 24. Luppes, R., Helder, J. A. and Veldman, A.E.P.: "The Numerical Simulation Of Liquid Sloshing

In Microgravity", European Conference on Computational Fluid Dynamics, Delft, The Netherlands, 2006.

25. M.S. Haberbusch and A. Chmiel, “Ultra Light Weight Flexible Sensing Probe Technology for Cryogenic Propellants,” AIAA 2000-3793, Proceedings of 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, pp. 1-11, AIAA, Huntsville, Alabama, 2000.

26. M.S. Haberbusch and C.B. Bryant, “Liquid Hydrogen Testing of an Ultra-Light Flexible Temperature and Liquid Level Sensing Probe,” Proceedings of 48th International Instrumentation Symposium, San Diego, California, 2002.

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27. M.S. Haberbusch, “Ultra-Light Flexible Temperature and Liquid Level Sensing Probe for Cryogenic Propellant Mass Gauging,” Advances in Cryogenic Engineering, Vol. 47B, American Institute of Physics, pp. 1292-1299, 2002.

28. M.S. Haberbusch, A.J. Culler, T.P. DePhillips, C.T. Nguyen, A.F. Skaff, and A.J. Yeckley, “A Cryogenic Lightweight Mass Gauging System for Expendable and Reusable Launch Vehicles,” AIAA 2002-3598, Proceedings of 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, pp. 1-12, AIAA, Indianapolis, Indiana, 2002.

29. Michaelis, Mark and Michael, Dreyer E. Test-case number 31: Reorientation of a Free Liquid Interface in a Partly Filled Right Circular Cylinder upon Gravity Step Reduction (PE). 28359 Bremen, Germany: ZARM, University of Bremen, 2003.

30. Michaelis, Mark and Michael, Dreyer E. Test-case number 31: Reorientation of a Free Liquid Interface in a Partly Filled Right Circular Cylinder upon Gravity Step Reduction (PE). 28359 Bremen, Germany : ZARM, University of Bremen, 2003.

31. Mikishev, G. N. and Dorozhkin, N. Y., Izv. Akad. Nauk. SSSR Otd., Vol. 4, 1961, pp. 48 { 83. 32. Miles, J. W., On the Sloshing of Liquid in a Cylindrical Tank," Tech. Rep. GM-TR-18, The

Ramo-Woolridge Corp. Guided Missile Research Div., April 1956. 33. Mourik, S. van, Veldman, A.E.P. and Dreyer, M.E. Simulation of Capillary Flow with a

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