NASA Student Launch College and University. Preliminary Design Review

Size: px
Start display at page:

Download "NASA Student Launch College and University. Preliminary Design Review"

Transcription

1 NASA Student Launch College and University Preliminary Design Review Institution: United States Naval Academy Mailing Address: Aerospace Engineering Department United States Naval Academy ATTN: NASA Student Launch Capstone 590 Holloway Road Mail Stop 11B Annapolis, MD Project Title: Navy Rockets Date: 3NOV17

2 NAVY ROCKETS TEAM MISSION The mission of Navy Rockets is to provide an expansion and application of classroom knowledge through a unique project based engineering opportunity. Navy Rockets strives to develop members morally and mentally by imbuing them with the highest ideals of engineering leadership and practice. During this year s Student Launch program, Navy Rockets will deliver a rocket and ground support element that incorporates a payload delivery system that meets all required criteria as defined by NASA and the Student Launch handbook. Overall, Navy Rockets is committed to excellence in practice, delivery, and conduct. NAVY ROCKETS CHARTER The vision of Navy Rockets is to: Supplement academic material in both the aerospace and engineering fields Expand each midshipmen s knowledge and experience to become more proficient and well-rounded members of the engineering community Provide leadership opportunities in a technical environment to better serve midshipmen as future leaders in today s Navy As a team we strive to: Seek out projects that can benefit the aerospace community and reinforce our own educational objectives ii

3 Deliver quality research and products on time, based in sound engineering and business practices, and operate to a level above client expectation As representatives of the armed services we will: Conduct ourselves in a professional manner and bring credit to both the United States Naval Academy and the United States Naval service. We are committed to excellence in practice, delivery, and conduct. iii

4 T a b l e o f C o n t e n t s List of Figures... vii List of Tables... ix 1 Summary of PDR Report Team Summary Launch Vehicle Summary Payload Summary Changes Made Since Proposal Changes Made to Vehicle Criteria Changes Made to Rover Criteria Vehicle Criteria Selection, Design, and Rationale of Launch Vehicle Mission Statement, Requirements, and Success Criteria Vehicle Design Material Selection Rocket Body Propulsion and Motor Choice Aerodynamic Analysis Fin Design Stability Drag Wind Tunnel Testing Avionics Electronics Avionics Bay Structure Avionics Testing iv

5 3.7 Recovery Recovery System Summary Black Powder Charges Component Selection Recovery Redundancies Mission Performance Predictions Flight Profile Stability Kinetic Energy Drift Alternate Calculation Method Safety Personnel Hazard Analysis Failure Modes and Effects Analysis Environmental Concerns Project Management Risks NAR/TRA Procedures Safety Brief Pre-Launch Brief Legal Consideration Material Safety Data Sheets Payload Criteria Payload Objective Payload Systems Design Review System Design Research Leading Payload Design Deployment Design Leading Deployment Design v

6 5.7 Payload and Launch Vehicle Interfacing Project Plan Requirements Verification Timeline Budget and Funding Material Acquisition Plan References APPENDIX A: Milestone Review Flysheet APPENDIX B: Open Rocket Design with Detailed Parts List APPENDIX C: Wind Tunnel Test Plan Introduction Airspeed Variation on Rocket Model APPENDIX D: Recovery System Checklists Appendix E: Drift Calculations Raw Data APPENDIX F: High Power Rocket Safety Code Appendix G: Legal Information Appendix H: MSDS Sheets vi

7 L i s t o f F i g u r e s Figure 1. Structural Diagram with labeled Components Figure 2. 3D Model of Rocket Figure 3. Motor Mount Fin Assembly (MMFA) Figure 4. Separation Coupler Figure 5. Avionics Coupler Figure 6. Avionics Sled Figure 7. Plywood Centering Ring Figure 8. Fin Motor Mount Assembly Figure 9. Fin Aerodynamic Planform Area Figure 10: Rocket Model Displaying CP, CG Figure 11. StratoLoggerCF Figure 12. Trackimo Universal GPS Tracker Figure 13. Jolly Logic Chute Release Figure 14. Avionics Bay Mounting Plate Figure 15. Avionics Bay Mockup Figure 16. Black Powder Casing and Bulkhead Figure 17. Quick Link Figure 18. Rocket Trajectory with Cesaroni L800-P Motor Figure 19. Cesaroni L800-P Motor Thrust Curve Figure 20. Drift Calculation Lower Section Figure 21. Drift Calculation Upper Section Figure inch Parachute Impact Energy Figure inch Parachute Impact Energy Figure 24. Risk Assessment Matrix Figure 25. Leading Payload Design vii

8 Figure 26. 3D Model of SASOR Figure 27. Leading Deployment Method Figure 28. Project Timeline Gantt Chart viii

9 L i s t o f T a b l e s Table 1. Body Tubing Material QFD Table 2. Fin Decision Matrix Table 3. Decision Requirement Matrix for Bulkheads Table 4. Shear Pin Shear Strength Calculations Table 5. Decision Matrix for Avionics Bay Bulkhead Table 6. Rocket Motor Performance Data Table 7. Lower Section Parachute Comparison Table 8. Upper Section Parachute Comparison Table 9. Personnel Hazard Analysis Table 10. Failure Modes and Effects Analysis Table 11. Environmental Concerns Table 12. Project Management Risks Table 13. Navy Rockets USLI Expected Costs Table 14. Navy Rockets Full Scale Itemized Budget Table 15. Navy Rocket s Travel Budget ix

10 1 S u m m a r y o f P D R R e p o r t 1.1 Team Summary Team Name: Institution: Mailing Address: Navy Rockets United States Naval Academy Aerospace Engineering Department United States Naval Academy ATTN: NASA Student Launch Capstone Mail Stop 11B 590 Holloway Road Annapolis, MD Project Mentors: Robert Utley (NAR Level 3) NAR # 71782, TRA # 6103 President, Maryland Delaware Rocketry Association Trip Barber (NAR Level 3) NAR# 4322, TRA# 2639 Former President National Association of Rocketry 1.2 Launch Vehicle Summary The launch vehicle will be a six inch diameter rocket and 104 inches long. It will weigh approximately 31.7 lbs and will be powered by a Cesaroni L800-P motor. The launch vehicle will separate into two sections, both of which will utilize a drogue chute and main parachute recovery system. The detailed milestone review flysheet can be found in Appendix A. 1.3 Payload Summary The Single Axle Self Orienting Rover (SASOR) is a two wheeled rover that will be deployed from the Payload portion using a stepper motor with a lead screw and two guide rails. It will utilize an accelerometer and an ultrasonic distance sensor to determine distance from the body of the rocket and drive a minimum of five feet before deploying a foldable solar panel utilizing a small servo mounted in the body. 1

11 2 C h a n g e s M a d e S i n c e P r o p o s a l 2.1 Changes Made to Vehicle Criteria Changes to the Vehicle Criteria include the length and weight of the rocket, with an increase from 93 inches to 104 inches, and an increase in weight from pounds to 39.1 pounds. Another significant design change came in the deployment method. The original proposed idea included a clam shell opening that allowed the rover to be deployed in a suitable orientation. Since the design utilizes an omnidirectional rover the clam shell was no longer necessary. 2.2 Changes Made to Rover Criteria Several changes have been made to the structural design, onboard sensors, and deployment method of the rover system. The new structural design for the payload is a two wheeled rover rather than the original design with four wheels. This design allows for easier deployment method possibilities because it is both self-righting and lighter. With two wheels, space will be maximized by having customized wheels formed in the shape of large hemispheres and slightly hollowed to allow space for the motors and gearing mechanisms. Changes have also been made to the rover s onboard systems by changing the sensors and actuators. The previous proposed design contained one direct current (DC) motor connected to a single axle to drive the rover, and one servo motor connected to the front axle to provide steering. The new design uses two DC motors, one for each of the two wheels to provide thrust and steering. The original design required two ultrasonic range sensors, one infrared (IR) distance sensor, two solar cells, and one accelerometer. The rover will not be using the IR sensor and will use just an accelerometer, a radio transceiver, and one ultrasonic range finder to navigate. The rover will carry and deploy one foldable solar panel that will be stored in an accordion fold style and stretched using a servo motor. A microprocessor will still be used to control all actuators and sensors on the rover. 2

12 The securing of the rover during flight and its deployment will be contingent on the axle of the rover, two rails on opposite sides of the inner body of the launch vehicle, and a stepper motor with a lead screw mounted to the bulkhead of the section divider in the rover section. The two rails will serve as a track for the axle of the rover to slide on to ensure no rotational movement during flight. Simultaneously, the lead screw will be actuated by the stepper motor and will screw the rover into the rocket through a threaded nut attached to the rover. After the rover has reached the desired secured position, the stepper motor will hold it in place until initiated again upon landing the launch vehicle. The reverse process will be used to deploy the rover: the stepper motor will spin the lead screw which will push the rover out of the body of the rocket, guided by the rails on the sides. 3

13 3 V e h i c l e C r i t e r i a 3.1 Selection, Design, and Rationale of Launch Vehicle Mission Statement, Requirements, and Success Criteria The mission of the Navy Rockets launch vehicle is to reach an apogee of 5,280 feet AGL, and subsequently deploy separate recovery systems for each of the two sections. The recovery system will enable each of the two sections to land with a kinetic energy that will not exceed 75 ft-lbf. Upon landing, an autonomous rover will be deployed. That rover will then travel a distance no less than five feet and autonomously deploy a solar panel. The mission will be considered a success if the rocket reaches an apogee of 5,280 feet AGL +/- 50 feet. To be deemed successful, the recovery system must deploy at the appropriate time to allow the rocket to touch down without damage. For the payload to be successful it must autonomously deploy from the rocket body, travel a minimum of five feet, and then deploy a solar panel. After all systems have been launched, or had an attempted launch, the launch vehicle shall be capable of re-launch within two hours. 3.2 Vehicle Design The overall vehicle design will be a two section rocket recoverable rocket. The lower section will be entirely used for propulsion. It will have a Motor Mount Fin Assembly (MMFA) which will fix to the rocket body tubing. The motor will slide into the MMFA and be rigidly attached using an Aero Pack retainer. Above the motor will be a fiberglass avionics coupler. The coupler will house the avionics sled that contains both altimeters and the global positioning system (GPS). Above the avionics coupler will be the recovery system of the lower section. Two rail buttons will be used to guide the rocket along the launch rail. They will be placed aft of the center of gravity. 4

14 Connecting the lower and upper sections of the rocket will be a separation coupler. The separation coupler will be rigidly attached to the lower section while the upper section will be attached with shear pins. The upper section of the rocket will contain the vehicle s payload. To allow a recovery system to deploy and the rover to deploy, the payload bay will be the aft-most section of the upper section. The payload bay will be designed to deploy the rover aft upon landing. This allows the rover to have a completely unimpeded exit path. Above the payload bay will be an avionics coupler identical to the lower section. Above the avionics bay will be the recovery system. The nosecone will be tethered to the upper section and will be fixed with shear pins. A complete layout of the vehicle design is shown in Figure 1. Figure 1. Structural Diagram with labeled Components. 5

15 A 3D model of the rocket is shown in Figure 2. Figure 2. 3D Model of Rocket. 3.3 Material Selection The structure of the rocket, shown in Figure 1, is designed to be as rugged and light as possible. The exterior structure of the rocket will be made up of four primary components: a nose cone, couplers, fins, and body tube sections. The nose cone will be fixed onto the forward body tube section with shear pins. There will be a total of four body tube sections that are all connected by couplers with ¼ inch steel screws. Two of the couplers will also make up a small portion of the exterior body of the rocket, since they enclose avionics housings that need to be able to transmit signals out of the rocket body. Carbon fiber acts as a faraday cage and blocks the transmission of radio frequency (RF) signals. Therefore it was necessary to create an avionics coupler that allowed the transmission of RF signals. These couplers provide transmission paths because they will be made of fiberglass which does not inhibit the transmission of the avionics. The third coupler will be entirely internal to the rocket body. It will connect the upper and lower sections of the rocket and will stay fixed to the lower section upon separation. The fins will protrude from 6

16 slots the aft body tube section, but will not be directly connected to it. The fins will be attached to the motor mount tube. The motor mount tube will have two centering rings attached to it so that it fits snuggly into the aft rocket body tube. This allows for the fins to be attached to the motor mount and the entire assembly slid into the rocket body. The interior structure of the rocket will consist of bulkheads of varying thickness and materials, avionics sleds, and the MMFA. The bulkheads facilitating the recovery mission will require a material and thickness that are suited to take on the loads associated with the recovery mission. Bulkheads associated with avionics and rover deployment will be made of materials and a thickness that is suited to their respective missions Rocket Body Rocket Body Tubing Carbon fiber was selected for its superior strength, low weight, and its availability to the team. The cost was significantly higher than alternative materials such as fiberglass, but the cost difference was not significant enough to push the design out of budget. Its low density allowed a reduction in launch vehicle dimensions and motor size necessary to reach the 5280 feet. Carbon fiber will be used throughout the majority of the launch vehicle body to decrease weight and increase strength. This material can be purchased commercially and is easy to modify with a high speed rotary tool. Carbon fiber and fiberglass are common materials used in high power rocketry. A house of quality was used to determine the strengths and weaknesses of using each material. The house of quality uses the Quality Function Deployment System (QFD). The QFD system allows important characteristics of materials to be compared and weighed against each other. The number values correspond to impact as follows: 1: Low importance 3: Medium importance 9: Critical importance 7

17 This weighing system allows important factors to outweigh less desired ones. If the material being judged met the description of the material factor, it was given a positive weight score. If it did not match the description of the material factor, it was given a negative weight score. If the material factor did not stand out as a factor, a negligible score of zero was given. The Material QFD for carbon fiber and fiberglass is shown in Table 1. Table 1. Body Tubing Material QFD. Material Factor Weighting Factor Carbon Fiber Fiberglass Low cost High availability Compact rocket size Low weight Easy production High tensile strength Highly compressive High stiffness High heat resistance High Young's modulus Large motor selection Total

18 The rocket body will be made from six inch diameter carbon fiber tubing. The outer diameter of the rocket body will be inches and will be 104 inches tall when fully assembled and the inner diameter of the rocket body will be 6.00 inches. The body sections will each be constructed of the carbon fiber tubing Motor Mount Fin Assembly The bottom section of the body will contain the Motor Mount Fin Assembly (MMFA). It will be held in place by a bulkhead at the top which is suited to take on the propulsion leads. The two centering rings of the MMFA are ½ inches thick and are attached with epoxy to the motor mount. The fins will be attached to the engine mount as part of the MMFA. There will be slits in the aft body section of the rocket to allow the MMFA to be slid up into the booster section of the rocket Fins Figure 3. Motor Mount Fin Assembly (MMFA). The launch vehicle s fins, shown in Figure 3, will be constructed from ¼ inch plywood. This material was chosen to fulfill several requirements, the most important being the simplicity of manufacturing and procurement. A decision matrix of the different materials considered for the 9

19 fins is shown in Table 2. The lightweight requirement needs to be met in order to give more flexibility for the team to design the payload which is critical to mission success. A house of quality was used to determine the strengths and weaknesses of using each material. The number values correspond to impact as follows: 0: Does not fulfill requirement 1: Partly fulfills requirement 2: Fulfills requirement This weighing system allows important factors to outweigh less desired ones. Table 2. Fin Decision Matrix. Plastic 3-D Printed Fiberglass over 1/4" Plywood Fiberglass over Foam Lightweight Easy to Make Easy to Duplicate Low Cost Durability/Withstands Impact Total The fins will be part of the MMFA, which will be secured into the aft carbon fiber body section with epoxy. The fins will be covered with fiberglass for added strength. 3-D printed ABS plastic is a common engineering material that was considered for the fin design. The ABS plastic is expensive and weights too much for the rocket design to handle as shown in Table 2. Fins constructed with fiberglass over foam did not offer the durability required for this mission. 10

20 Many fiberglass layers would be necessary to create the desired rigidity and durability, and in doing so significant weight would have been added to the fin design Nose Cone An ogive nose cone will be used in the final design. It will measure 18 inches tall from mating with the rocket to the point of the ogive. At the base, the nose cone will have a diameter of 6 in. The total weight of the nose cone will be lbs. The nose cone will be purchased commercially and will be made of fiberglass. This material provides strength and interior space that can potentially be filled with ballast if required Alternate nose cone designs used different materials which were inadequate in terms of the design constraints. Foam covered with fiberglass was considered, but the foam would not leave any room for ballast in the nose cone. Also the fiberglass covered foam design required the team to construct the nose cone to meet the dimensions of the rocket body tubes. Precisely constructing foam and fiberglass is a task best left to professionals who can produce fairly inexpensive products that have more precise dimensions. Wood turned on a lathe was also considered for the nose cone design. The wood would have to be turned until it was very thin, which means it would be brittle and more liable to crack upon landing. Fiberglass is more flexible and is able to withstand straining without breaking Bulkheads The bulkheads for this rocket design have to meet the requirements of the missions they support. The decision matrix shown in Table 4 shows how each potential material fulfills the various requirements. The number values correspond to impact as follows: 0: Does not fulfill requirement 1: Partly fulfills requirement 2: Fulfills requirement 11

21 The values are not totaled for the materials in this matrix because there are different bulkheads in the rocket with different requirements. Table 3. Decision Requirement Matrix for Bulkheads. 1/2" Plywood Aluminum Lightweight 1 1 Easy to Make 2 0 Low Cost 2 1 Axial Blast 1 2 Torsion Strength 1 2 For the recovery bulkheads, a requirement is that the bulkheads are able to withstand the blast from the black powder charges. The recovery bulkheads also need to have u bolts installed to attach recovery tethers to. The bolts go through the bulkheads and are secured by bolts that are screwed onto the other side. A spreader plate is tightened between the nut and the bulkhead so that the force of the tether pulling on the bolt and nut is spread evenly through the bulkhead Separation Coupler The separation coupler will be an entirely internal component. It will be made of vulcanized paper and will attach to the rocket body by both shear pins and rigid bolts. The rigid bolts will ensure the coupler stays attached to the lower section of the rocket while shear pins will allow the upper section to separate. This decision was made so that the separation coupler stays attached to the lower section of the rocket. This allows the rover an unimpeded path to exit the rocket body. The separation coupler is shown in Figure 4. 12

22 Figure 4. Separation Coupler Shear Pins Shear pins will be used to attach the separation coupler and the nosecone to the rocket body. For the connection between the nosecone and body tube two 4-40 nylon shear pins will be used to retain the nosecone during flight. For the separation coupler, two 4-40 nylon shear pins will be used keep the upper section and separation coupler in flight. The proper number of shear pins was determined by first calculating the minimum and maximum shear strength of standard shear pins. Using the pressure created by the black powder charge in the recovery section and the strength of an individual shear pin, a total number of required shear pins was determined. Using the diameter of the shear pin the cross sectional area was calculated using Equation 1. A = πd2 4 (1) Once the area was calculated the shear strength was calculated using Equation 2. Where F is shear strength, P is shearing pressure, and A is cross sectional area. F = PA (2) Shear strength was calculated given the minimum and maximum shear strength pressure for 13

23 nylon shear pins is 9600 and psi. Given the 4-40 nylon shear pin diameter is mm (0.112 in), the maximum shear strength is pounds as shown in Table 4. Using Equation 2 and assuming 10 psi of separation pressure (see section 3.7.2), the force distributed on the rocket from one black powder charge would be lbs. Dividing the distributed force by the individual shear pin strength, it was calculated that 2.73 shear pins are required, therefore two 4-40 shear pins will be used for each recovery section. Table 4. Shear Pin Shear Strength Calculations. Screw Shear Strength Min Max Nylon 6/6 Shear Strength(psi) Areas (sq Shear Strength Screw Diameters(in) in) (lbs) Size Minor Minor Min Max M M Avionics Couplers The avionics couplers, shown in Figure 5, are designed to be made of fiberglass, which allows the avionics system to transmit outside of the rocket. Otherwise, the carbon fiber body would act as a faraday cage. A 3-D printed plastic avionics coupler was considered, but was too heavy for it to have similar strength properties to the carbon fiber body sections. The use of fiberglass for the couplers was justified because its strength is comparable to that of the carbon fiber body, and the two interlocking materials must have similar strength properties to avoid failure at the connections. Fiberglass also has a lower tensile modulus than carbon fiber making it 14

24 better suited to bend and strain without breaking, these are characteristic properties of rocket couplers because of the forces applied to the structure from propulsion and ground impact. Figure 5. Avionics Coupler. The coupler has a 1 inch midsection that is inches in diameter which is the same as the exterior diameter of the main rocket body. The small diameter section has a diameter of 5 ¾ inches and are each 6 ½ inches long Avionics Bays The avionics housings are formed by the fiberglass avionics couplers and the wooden avionics bulkhead which secures the sled guiding screws. A removable avionics sled is enclosed within each avionics bay and has holes for the two guiding screws to hold the sled in position. 15

25 Figure 6. Avionics Sled. The avionics sled, shown in Figure 6, will be made with eastern white pine wood. This type of wood is readily available in the USNA workshop, it s easy to shape, and, its most important attribute, it has high shear and torsional strength while being lightweight. 3-D printed Acrylonitrile Butadiene Styrene (ABS) plastic and fiberglass avionics sled designs were also considered. A house of quality was used to determine the strengths and weaknesses of using each material. Table 5. Decision Matrix for Avionics Bay Bulkhead. Eastern White Pine Plastic 3- D Printed Fiberglass Torsional strength Lightweight Easy to duplicate Easy to make Cost Total

26 3.4 Propulsion and Motor Choice The entire rocket was modeled on Open Rocket, a common high powered rocket software. Through modeling the rocket on Open Rocket it was possible to perform simulations on various configurations. These simulations produced the data shown in Table 1. The configuration of the rocket is shown in Figure 1. From this Open Rocket software the center of gravity, center of pressure and stability margin were all calculated. The stability margin was calculated to be 2.08 which is in accordance with (IAW) NASA SLI launch requirements. The launch vehicle will be powered by a commercially available Cesaroni L800-P Classic Rocket Motor. The primary goal of selection is to safely reach 5,280 feet above ground level at apogee of the flight. Using a Cesaroni L800-P Classic Rocket Motor the rocket reached 5,619 feet in an Open Rocket flight simulation. The Cesaroni L800-P was chosen due to a tolerance of 6.4% within the height of desired apogee. The tolerance is needed because the final weight of the rocket is predicted to be 31.7 pounds, and having this tolerance allows for changes to the design if needed during the full scale build. This tolerance will also allow the rocket to reach its assigned altitude given adverse weather conditions or variations in projected trajectory. The Cesaroni L800-P has a weight of five pounds, which is 24.42% of the total weight of the rocket. An L size motor was selected because the required total impulse to allow for this rocket to reach an apogee of 5,280 feet with 31.7 pounds was never achieved when running the largest size K motors in Open Rocket. The performance characteristics of a Cesaroni L800-P motor compared to other potential motors are given in Table 6. Flight simulations were run in Open Rocket with the same rocket configuration while varying the motor. Detailed flight simulation data is shown in Appendix B. 17

27 Table 6. Rocket Motor Performance Data. Configuration Apogee (ft) Flight Time (s) Max Acceleration (ft/s^2) Max Velocity (ft/s) L800-P L995-RL L890SS-P L805- WH The comparison of rocket motors allowed for a clear choice to be made in favor of the Cesaroni L800-P solid rocket motor due to its higher performance with the predicted rocket characteristics. Considerations were taken based on performance characteristics such as apogee height, maximum acceleration, and maximum velocity. Through the Open Rocket simulations, it was clear that the L800-P had the greatest apogee height, a low maximum acceleration and moderate maximum velocity in comparison to the other L sized motors. This showed that the motor would reach the desired altitude without imposing too great of an acceleration (7.45 g s) which could potentially damage the payload of the rocket. Cesaroni was chosen for the simplicity as compared to other motor brands. Cesaroni did not require a seal ring to be installed for the motor. The increased complexity of installing a seal ring was part of Aerotech motors of similar size. The seal ring is a potential point of failure if installed incorrectly. Therefore, it was decided that the motor selection would include a preinstalled seal ring. The propulsion assembly must be re-loadable within two hours in accordance with (IAW) NASA SLI project requirements. The motor will be loaded into the suggested casing from Cesaroni which is the Pro75 Classic motor mount. The motor inside the Pro75 Classic mount will then be loaded into an assembly already attached to the rocket body. The assembly consists of motor centering rings shown in Figure 7 and a motor sleeve as shown in Figure 8. 18

28 Figure 7. Plywood Centering Ring. The full assembly is shown in Figure 8. This assembly will be permanently installed into the rocket body by putting epoxy on the centering rings and on the interior of the rocket body. Figure 8. Fin Motor Mount Assembly. 19

29 3.5 Aerodynamic Analysis Fin Design Fins are used on this rocket in order to improve the stability of the vehicle. Fins accomplish this by moving the center of pressure (CP) aft, and by providing a restoring moment to the rocket when disturbances are experienced in flight. The launch vehicle will have three fins to provide correcting force for stable flight. The fins are separated by 120 degrees. Three fins are used because this design will produce less drag than designs with greater numbers of fins. Furthermore, three is the minimum number of fins that will provide a restoring moment against a disturbance in any direction. This design choice minimizes drag while maintaining rocket stability in all flight conditions. The fins will have a trapezoidal shape. This shape was chosen because this basic shape is easy to manufacture and can be quickly produced. This quality will be important since it allows fins to be readily replaced in the event of breakage or design shift. The length of the fin root is seven inches long and tapers at a 25 degree angle from the top extending out four inches to the tip. The height of the fin is six inches. Each fin has an area of 33 square inches, and will be constructed from ¼ inch plywood layered with fiberglass for added strength. The fin profile is shown in Figure 9. Figure 9. Fin Aerodynamic Planform Area. 20

30 The fin will extend through the vehicle body and be attached to the motor casing within. Each fin will be axially aligned with the rocket body, and the offset between the fins will be exactly 120 degrees. The full fin assembly is shown in Figure 3, which shows the interface between the fins, body tube, motor casing, and centering rings. Success criteria for the fins will depend on their ability to provide correcting forces to ensure straight upward flight. These criteria will be tested in a wind tunnel and implemented on the scale model launch vehicle Stability The stability of the launch vehicle is dependent on the position of the Center of Gravity (CG) relative to the Center of Pressure (CP). The stability margin of a rocket is defined as the distance between the CP and CG, measured in calibers, where one caliber is the maximum body diameter of the rocket. The NASA Student Launch requirements stipulate that the stability margin must be no less than two calibers. A computer model of the launch vehicle was created using the Open Rocket software suite. The model included every component that will be included on the final rocket, and made a scientific estimate of payload mass and location. Based on these estimates, the CP is located at inches and the CG is located at inches from the nosecone. The rocket body tube has a diameter of six inches and the resulting stability margin is 2.08 calibers. The rocket model is shown in Figure 10, where the CP location is represented with a red dot, and the CG location is represented with a blue dot. Figure 10: Rocket Model Displaying CP, CG 21

31 As the launch vehicle design evolves, the CG location may shift. The stability margin will be maintained by altering the area of the fins Drag Drag force is the result of the interaction between the rocket and air as the rocket travels through the air. The design seeks to minimize the impact of drag on the performance of the rocket by maintaining a smooth, uninterrupted outer shape and minimizing fin surface area. The skin friction component of drag can be minimized by having smooth aerodynamic surfaces. The profile drag is minimized by using the minimum number of fins with an area no greater than what is required to maintain the stability index. Refer to the Wind Tunnel Testing section for the methodology for calculating the launch vehicle drag Wind Tunnel Testing To model the aerodynamic performance of the rocket during flight, a full scale model of the rocket will be constructed and tested in the open circuit wind tunnel. This model will be constructed from the same materials that will be used to build the flight ready rocket, and will closely mimic the outer shape of this final product. The wind tunnel tests will be conducted in the Eiffel Wind Tunnel, which is located in the Rickover Hall Aerospace Lab at the United States Naval Academy. The model will be tested at a number of different velocities, and the drag force on the model will be measured. The range of these tests will be limited by the maximum air velocity that can be generated by the wind tunnel. A drag coefficient will be calculated for each tested velocity. Using the Open Rocket software suite, a model of the final rocket will be created and tested. This software can calculate the expected drag coefficient of a rocket at different velocities. Using this tool, the expected drag coefficient will be calculated for each velocity that was tested in the wind tunnel. The expected and experimental values for drag coefficient will be 22

32 compared to determine the accuracy of the Open Rocket tool. If the Open Rocket results are validated, a percent difference will be applied. This procedure of comparing drag coefficient values is being used because the Eiffel Wind Tunnel is unable to simulate the high velocities that the rocket will experience in flight. The wind tunnel is only capable of reaching a maximum velocity of 300 ft/s while the rocket will fly at varying speeds reaching a maximum of 650 ft/s. For this reason, the model will be tested over the range of velocities within the wind tunnel s capabilities, and these values will be used to determine the accuracy of the Open Rocket computer simulation. The Open Rocket software is expected to produce accurate results due to the simple external shape of the rocket. The rocket design does not include any protrusions or irregular features that would introduce inaccuracies in the Open Rocket results. The skin-friction of the wind tunnel model will closely resemble that of the final rocket, because the same materials will be used for construction. Both the model and the rocket will utilize carbon fiber tubing, a smooth fiberglass nosecone, and a smooth-finished fin set. Additionally, any minor surface irregularities in the rocket design will be reproduced on the wind tunnel model. In order to use the Eiffel Wind Tunnel in Rickover Hall at the USNA, a test plan must be submitted to the aerospace engineering department chair. A sample of the test plan is attached in Appendix C. The sample test plan is subject to change based on the suggestion of the department, but it will remain similar in structure and concept. The test plan shows the pertinent information of the testing process, including purpose, philosophy of operations, and the testing procedure. 3.6 Avionics The rocket separates into two independently recovered sections, and an avionics bay will control the recovery of each section. Therefore, the avionics system of the rocket consists of two separate, but identical avionics bays. These sections must be easily accessible, and have the 23

33 capability to be armed while the rocket is on the launch pad. The sections will contain redundant systems to activate the charges to separate the rocket sections, as well as GPS in order to track the rocket sections Electronics Altimeters will be used to activate separation charges at apogee. The altimeter chosen for use is the StratoLoggerCF. Other altimeter such as the StratoLogger SL100, Pnut Logging Altimeter, and the Affordable Precision Rocket Altimeter (APRA) were considered; however the APRA was quickly removed as an option because it cannot control recovery systems. The StratoLoggerCF was chosen because it was the smallest and lightest out of the options, while also having the ability to record and save flight information. The StratoLoggerCF uses barometric pressure in order to determine the altitude of the rocket. It features an extremely small size and weight, which is optimal in an altimeter. Each altimeter will be powered by a single nine volt battery. Each avionics bay will include two StratoLoggerCF altimeters in order to ensure redundancy of the recovery system. This ensures redundancy because it has two entirely separate recovery systems which could each individually ensure separation. One of the StratoLoggerCF altimeters will also be used as the scoring altimeter. The StratoLoggerCF altimeter is pictured in Figure 11. Figure 11. StratoLoggerCF The GPS chosen for use is the Trackimo Universal GPS Tracker. This GPS device will transmit its location to a ground station in order to ensure the discovery and recovery of the rocket after flight. The unit is powered by an on board Li-Ion battery, is small in size and self-sufficient. 24

34 Other GPS devices were considered, such as the AIM XTRA 2.0 GPS, however the previous Navy Rocket team utilized these units and it resulted in a premature detonation of the separation charges. The AIM XTRA 2.0 GPS is also larger than the Trackimo system. For this reason the Trackimo Universal GPS Tracker was selected. A single Trackimo Universal GPS Tracker will be included in each avionics bay. The Trackimo Universal GPS Tracker is shown in Figure 12. Figure 12. Trackimo Universal GPS Tracker In order to reduce the drift of the rocket, the main parachute will be deployed long after apogee. The first option to meet this goal was to use a dual deployment system that detonates a drogue and a main chute separately. However due to the complexity of packing the parachutes and executing this system, the more simple solution of using Jolly Logic Chute Release was chosen. This system features a unit that straps around the main parachute and at a chosen altitude the straps release, and the main parachute opens to slow the rocket s descent speed. This choice ensures simplicity, minimizes drift, and ensures a slow landing speed. In order to ensure redundancy, two Jolly Logic Chute Releases will be interlocked and used for each main parachute. This results in a system where if either Jolly Logic Chute Release releases, then main parachute deploys. This eliminates the risk of the main parachute not opening. A Jolly Logic Chute Release is included in Figure

35 Figure 13. Jolly Logic Chute Release Avionics Bay Structure The initial design for the avionics bay was a self-contained pod that contained all of the avionics components. This pod would then be able to slide in and out of the body tube to be able to be accessed. However, this initial design had a few problems; the first being that if the bay slides in from the top of the rocket body tube, it would be very difficult to reach and access if other rocket components are packed on top. Another issue is that the carbon fiber body tube could block the signals from the GPS and cause it to not function as intended. The chosen avionics bay consists of a tube of fiberglass that couples into the rocket body with rigid bolts. This design allows the entire avionics bay to be detached from the rest of the rocket to allow for easy maintenance, yet remain securely attached during flight. The choice of fiberglass allows the GPS signal to propagate. The coupler design contains a section that is of equal diameter to the exterior of the rocket; this section acts as an RF window that creates a path for RF signals to propagate. The interior diameter of the bay is constant, while the exterior diameter of the bay has two sections of smaller diameter to couple with the body tube. The length of the sections intended to couple with the rest of the rocket is equal to the diameter of the rocket, and in this case is six 26

36 inches. In the center of the coupler there is a one inch section of the same diameter as the exterior of the rocket. Drilled into this section will be four small holes to allow the altimeters to sample the barometric pressure, as well as two larger holes to house the rotary switches to arm the altimeters inside the bay. The rotary switches are connected to the altimeters, and once on the launch pad these switches can be turned and locked into the on position, and the altimeters will be on and armed. The proposed design is shown in Figure 5. Inside of this avionics bay will be a permanently attached ¼ inch thick wood plate that has two ¼ inch threaded rods embedded in it. The ¼ inch thickness was chosen because the mounting plate is only going to hold the weight of the avionics sled, and is not essential to the structural integrity of the rocket. The wood plate will be glued into the avionics coupler using epoxy glue. The ¼ inch threaded rods will be attached to the wood plate using epoxy as well as washers and nuts under the plate. This will be used as the attachment point for the removable avionics segment. This permanently fixed attachment point is shown in Figure 14. Figure 14. Avionics Bay Mounting Plate. The removable avionics segment will consist of two wooden disks joined in the center with a rectangular plate. All of the avionics components will be affixed to this removable avionics sled. Each disk will have two ¼ inch holes aligned to each other in order to allow the threaded rods to 27

37 pass through them. The disks are each 5.5 inches in diameter, and the rectangular plate is 6 inches in height. The design is shown in Figure 6. The GPS and altimeters will be mounted onto the avionics sled. The nine volt battery will be attached to the base disk of the sled, and the GPS and altimeters will be attached to the vertical rectangular section. In order to configure the system for flight, the removable sled is slid onto the two fixed threaded rods. A steel washer is placed on each ¼ inch steel rod and then a ¼ inch steel nut is fixed on the end in order to secure the sled in place. This configured avionics bay is then coupled securely with bolts to the rest of the rocket body in order to allow for flight. A full mockup of the avionics bay is shown in Figure 15. Figure 15. Avionics Bay Mockup Avionics Testing In order to ensure that the avionics will function properly in flight, the system will be tested before continuing to a flight ready model. The altimeter and recovery system will be tested by detonating a packed black powder charge during a drop test using a setup identical to what will occur in flight. If this test is successful, it will demonstrate that the altimeters for recovery will function properly. In order to test the GPS system, the GPS will be turned on and the avionics bay will be sealed as it will be in flight. The system will be moved a mile away, and the ground station will attempt to acquire the location of the avionics bay. If this test is successful, it will 28

38 demonstrate that the GPS is able to propagate its signal through the avionics bay exterior. A final test will be conducted to determine where the best position for the GPS unit will be. During the sub-scale launch, one of the GPS will be attached to the avionics sled, and the other will be tied to the parachute of one of the recovery units. This test will determine if the GPS unit functions while attached to the parachute, and if this is the case the fiberglass avionics bay to ensure signal transmission is unnecessary, and the avionics sled can be attached into the rocket body with an internal coupler. 3.7 Recovery Recovery System Summary The recovery system of the rocket consists of two independent sections; the lower and upper sections respectively. At apogee the altimeter in the lower section of the rocket will trigger the detonation of the primary black powder charge along the lower section / upper section separation seam. To protect the parachutes during the ejection blast, nomex blanket will be used. Exactly one second after the detonation of the lower section primary charge, the redundant charge in the lower section and the primary charge in the upper section will detonate to separate the nosecone. Next, exactly one second later the redundant charge in the upper section of the rocket will detonate. Detonation will break shear pins along the lower and upper seams of the rocket. Once the lower and upper sections of the rocket are separated from one another and the nosecone is separated from the upper section, drouge chutes will deploy from each section as a means of deploying the unopened main parachute from each respective recovery section. Additionally, the drogue chutes will slow the descent rate of each section as a means to prevent hazardous magnitudes of shock upon main parachute deployment. Once the main parachutes are pulled out of the internal structure of their respective subsections they will remain packed via the use of a Jolly Logic Device. The Jolly Logic devices will delay main parachute deployment until 800 ft AGL. Calculations used to determine the release altitude of 800ft AGL can be found in section Once the main parachutes are deployed the rocket will descend to the ground and land with a kinetic energy that will not exceed 75 ft-lbf. The kinetic energy upon landing was calculated to be ft-lbf and ft-lbf for the lower and upper sections respectively. A 60 29

39 inch diameter parachute will be used as the main parachute for the lower section, and an 84 inch diameter parachute will be used as the main parachute in the upper section. The main parachute size was determined via descent rate calculations to ensure the rocket does not land with more kinetic energy than 75 ft-lbf. Parachute area was calculated using Equation 1, descent rate was calculated using Equation 3, kinetic energy was calculated using Equation 4, and a table comparing parachute sizes to descent velocities and kinetic energy is displayed in Table 7 and Table 8. v d = 2W s C D ρ A (3) v d is descent velocity, W s is individual section weight, C D is coefficient of drag and was assumed to be 1.5, ρ is density of the air at standard sea level, and A is the area of the parachute. The assumption was made that the change in the air density would be negligible. KE = 1 2 mv d 2 (4) KE is the kinetic energy upon impact, and m is the mass of the section. Table 7. Lower Section Parachute Comparison. Diameter Chute Diameter (in) (ft) Stage Mass (slugs) Decent Velocity (ft/s) Landing KE (ft-lbf)

40 Table 8. Upper Section Parachute Comparison. Diameter Stage Mass Landing KE (ft- Chute Diameter (in): (ft) (slugs) Decent Velocity (ft/s) lbf) The purpose of separating the lower and upper subsections is in order to allow the rover to deploy from the internal structure of the rocket upon a successful descent and recovery. The purpose of separating the tethered nosecone from the upper section of the rocket is to allow the recovery system of the upper section to deploy. At apogee the onboard altimeter will send a current to the black power charges attached to the bulkhead, which will ignite the black powder housed in pieces of PVC pipe capped with masking tape. Figure 16 displays a bulkhead with black powder charges from a previous Navy Rockets team. The layout for black powder charges on a bulkhead will resemble the previously used design. 31

41 Recovery Flight Plan Phase 1: Launch Figure 16. Black Powder Casing and Bulkhead. Phase 2: Rocket reaches apogee and primary black powder charge ignites along the lower-cut section only separating the upper and lower stages of the rocket and shearing the two lower section shear pins. Phase 3: The drogue chute from the lower subsection deploys and pulls out the lower subsection main parachute with the Jolly Logic devices attached. Phase 4: The redundant charge in the lower section and the primary charge in the upper section detonate causing the tethered nose cone to separate from the upper payload section shearing the two upper section shear pins. Phase 5: The drogue chute for the upper subsection deploys and pulls out the main chute for the upper section with the Jolly Logic devices attached Phase 6: At lower section deployment altitude the Jolly Logic device deploys the 60 inch main parachute for the lower subsection 32

42 Phase 7: At upper section deployment altitude the Jolly Logic device deploys the 84 inch main parachute for the upper subsection Phase 8: The lower and upper subsections of the rocket land safely on the ground in such a manner that promotes a successful rover deployment. Recovery subsystem checklists are located in Appendix D Black Powder Charges These amounts of black powder were calculated for an ejection pressure of 10 psi [12]. Equation 5 was used to calculate the amount of black powder for the upper and lower section charges. X = C D 2 L SF (5) C is the pressurization constant (0.004 for 10 psi), D is the Diameter of the Rocket tube (six inches), L is the length of body tube where the recovery equipment is housed (12 in for lower section, 14 in for upper section), and SF is the safety factor (two) grams of black powder will be used for the lower section and 4.03 grams of black powder will be used in the upper section grams of black powder will be used for the redundant black powder charges. This increases the factor of safety to 2.60 and 2.23 for the lower and upper sections respectively Component Selection The components included in the recovery subsection were critically analyzed to determine the safest and most effective system. U-bolts were selected over eye-bolts for shock cord attachment on the recovery system bulkheads. An eye bolt has the advantage of requiring less parts when compared to a U-bolt as no washers or plates are required. However, the U-bolt has the advantage of being much lighter than the eye bolt, and weight is a much greater concern than number of parts for the mission at hand. Additionally, multiple attachment points on a U-bolt allow for the effective use of a spreader plate on the back of the attached bulkheads. The 33

43 spreader plate helps distribute loads over the entire bulkhead as opposed to singular screw points. The two alternatives for shock cord were half-inch nylon shock cord and one-inch shock cord respectively. Advantages of using the half-inch nylon shock cord include less weight and space required. A major disadvantage is the lower tensile strength when compared to one-inch nylon shock cord. Although the one-inch nylon shock cord will take up more space and add more mass to the rocket it is the better alternative for shock cord. In order to use a permanently mounted U-bolt, Quick Links were introduced as the method of attaching shock chord to the U-bolt. The shock cord will have a loop sewn on the end of the cord in order to attach to the Quick Link shown in Figure 17. Figure 17. Quick Link Recovery Redundancies The black powder charges for the lower and upper subsections will include two charges; each of which are wired completely independently from each other to ensure redundancy. The first charge to ignite will be the lower section primary charge. In the lower section of the rocket the primary black powder charge will detonate at apogee, while the redundant charges will detonate 34

44 in the upper and lower sections exactly one second after each respective primary charge detonates in the lower and upper sections. The recovery system will use two separate altimeters wired with two independent circuits. Since there are two recovery systems there will be a total of four altimeters and four black powder charges included in the rocket system. Furthermore, two Jolly Logic devices will be attached to the main parachutes in the lower and upper sections making a total of four Jolly Logic devices. The Jolly Logic devices will all be set to 800 ft AGL. The Jolly Logics will be connected such that only one of the two is required to release the main parachute. To ensure proper separation ground ejection tests will be performed prior to launch. Tests will be used to validate the quantity of black powder calculated for use in the lower and upper sections. The tests will also be used to demonstrate control of the electronic system. An ammeter will be used to verify continuity and proper wiring along the black powder ignition system prior to testing. Success in the ground ejection tests will be measured by the seamless separation of the subsections from each other in a manner that does not damage or degrade any portion of the rocket. 3.8 Mission Performance Predictions Flight Profile The vehicle s flight profile was simulated using Open Rocket software and is shown in Figure 18. The flight profile used a Cesaroni L800-P motor and accounted for the main parachutes opening at 800 ft AGL. 35

45 Figure 18. Rocket Trajectory with Cesaroni L800-P Motor. A motor thrust curve for the Cesaroni L800-P is shown in Figure 19. Figure 19. Cesaroni L800-P Motor Thrust Curve. Estimated component weights are displayed in Appendix B. 36

46 3.8.2 Stability The estimated CP of the rocket is located at inches aft of the nosecone and the CG is located at inches. The rocket body tube has a diameter of six inches and the resulting stability margin is 2.08 calibers. Visual representation is shown in Figure Kinetic Energy The kinetic energy of the upper and lower sections upon landing were calculated based on the parachute size and section mass. The lower and upper sections were calculated to have a kinetic energy of ft-lbf and ft-lbf respectively Drift Drift calculations were developed for the upper and lower sections of the rocket using Equations 6 and 7. The drift distance was plotted for wind speeds of 5-20 knots. The assumptions associated with the drift calculations include: 1.) Drogue chute does not contribute to the drag after main parachute deployment 2.) The kinetic energy on landing is only dependent on the vertical velocity 3.) The horizontal drift velocity of each section equals the velocity of the wind 4.) Vent hole is accounted for in drag coefficient 5.) The parachutes are perfect circles 6.) Mass of lower section is without any propellant Time to descend was calculated using (6), where Alt is the altitude AGL. t = Alt/v d (6) Drift distance was calculated using (7), where v w was the horizontal wind velocity. 37

47 Drift (feet) D = v w t (7) Graphical results from the drift calculations are plotted in Figure 20 and Figure 21. A horizontal line on each graph at 2500 ft represents the drift distance limit as per the NASA Student Launch Handbook. Raw data from drift calculations are located in Appendix E Knots 10 Knots 15 Knots 20 Knots Max Drift Linear (Max Drift) Release Altitude (ft) Figure 20. Drift Calculation Lower Section. 38

48 Drift (feet) Knots 10 Knots 15 Knots 20 Knots Max Drift Linear (Max Drift) Release Altitude (ft) Figure 21. Drift Calculation Upper Section Alternate Calculation Method As a method to confirm drift calculations. Online sources were used as a point of comparison. Our intended parachute vendor, Fruity Chutes, has their own online calculator that can be used to calculate energy upon landing and descent velocity. The figures included in this section show the results of the online calculations. 39

49 Figure inch Parachute Impact Energy. Converting joules to ft-lbf, the impact energy for a 60 inch parachute would be ft-lbf. This validates the calculated value of ft-lbf. Figure inch Parachute Impact Energy. 40

50 Converting joules to ft-lbf, the impact energy for an 84 inch parachute would be ft-lbf. This validates the calculated value of ft-lbf. 41

51 4 S a f e t y The Navy Rocket team will ensure that care and proper precautions are taken to mitigate the risks of high powered rocketry. The safety officer, MIDN Gomez, will be in charge of maintaining the overall safety of the project, and each individual team member will be made aware of the risks and hazards and how to mitigate them. If any activity is seen as unsafe it will be discontinued until the safety risk can be assessed and mitigated. 4.1 Personnel Hazard Analysis In order to manage risk in the project, either the likelihood or the consequences of any potentially dangerous activity will have to be reduced. Table 9 includes a list of potential risks to the project, and methods to mitigate them. The legend to interpret the likelihood and severity ratings is included in Figure 24. Figure 24. Risk Assessment Matrix. 42

52 Table 9. Personnel Hazard Analysis. Hazard Cause Effect Mitigation Probability/Severity Injury while Failure to wear Member injury. All team members will be given 1-C using protection. a lesson on proper equipment Manufacturing usage before beginning, as well Machinery as always wearing proper PPE Chemical Spills/ Improper handling of Environmental damage. Before any usage of chemicals, 2-D Contamination materials. the MSDS will be read. The chemicals will also be stored in a secure location Unexpected Failure to follow Motor launches before The combustibles will always be 2-C Combustion of procedures. necessary time. Rocket will stored in a safe and secure Motor or Black be damaged. location. They will also be Powder handled with extreme care. Before tests or launch a safe distance will be achieved. Collision of Foreign object in flight Damaged rocket and All of the FAA and NAR rules 1-D Rocket in Flight path. foreign object. and regulations will be followed at all times. 4.2 Failure Modes and Effects Analysis There are a number of failures that could occur with the proposed design of the rocket. Table 10 includes a number of possible failure modes in the rocket, ways to mitigate the possibility of these failures, and the likelihood and severity that these failures would affect mission success. 43

53 Table 10. Failure Modes and Effects Analysis. Hazard Cause Effect Mitigation Probability/Severity Misfire of Rocket Poor Weather on Launch Parachute does not Deploy Properly Altimeter and Software does not Function Rocket does not Separate Separation Coupler gets caught on Rover Rails GPS Signal does not Propagate Rover fails to Move Distance Sensors for rover Fail Bad ignitors or rocket motor is wet. Rocket does not launch Follow the launch procedures exactly. After a misfire a safety wait will be observed, and then the rocket will be disarmed and safely stored. An investigation will be conducted after. N/A Scrub launch day Check forecast leading up to launch, and if it changes assess the weather situation to decide launch. Parachute improperly packed or black powder charge too small. Improper program interfacing. Improper wiring. Dead power source. Black powder charge fails to separate shear pins. Misalignment of separation coupler. Weak power source. Propagation path blocked. Rover does not receive signal. Rover breaks upon landing. Sensor wiring breaks loose. Weak power source. Sensors fail upon landing. Rocket descends ballistic and is destroyed. Black powder charge does not detonate. Parachute does not deploy. Parachute does not deploy. Rover exit path impeded. Ensure that the parachute is packed properly, and that sufficient wadding is included. Testing of black powder prior to launch will be completed. Check batteries before launch, and test the systems multiple times before launch. Include redundant systems. Adequately test the separation charges on the ground prior to any launch. Carefully follow the instructions to pack the charges, and include redundant systems. Link and test the system with and without charges prior to any launch. Rocket is difficult to locate. Test the chosen GPS system in and out of the avionics bay to ensure that the signal will propagate. Rover does not complete its mission. Rover does not ensure 5ft radius from rocket body. Test the rover on a surface that it will land on before launch. Ensure the design can handle a variety of terrain. Test the rover and all of its sensors prior to launch. 1-D 3-D 1-C 1-D 1-D 2-C 3-D 3-C 4-C 44

54 4.3 Environmental Concerns The rocket can be affected by the environment negatively and can also affect the environment negatively. These possibilities are included in Table 11. Table 11. Environmental Concerns. Hazard Risk Mitigation Probability/Severity Rain High Winds The electronics could get wet, and the black powder could get damp and not combust. To mitigate all electronic and recovery work would be done under cover, and sealed while dry. The stability of the rocket both at launch and in flight could be compromised. Scrub launch. 2-D Scrub launch. 2-C Clouds The rocket could be lost from sight and reappear in an Ensure GPS is 3-D unexpected location. functioning properly. Motor lights grass The fire could spread and damage the area surrounding Ensure fire 2-D on Fire the launch area. To mitigate do not launch over easily extinguisher is readily combustible grass, and carry fire extinguishing accessible. equipment. 4.4 Project Management Risks Good project management is essential to the success of the project and the achievement of the mission. A number of risks to the project as a whole are included in Table

55 Table 12. Project Management Risks. Hazard Mitigation Probability/Severity Falling behind Schedule Create and strictly adhere to the deadlines and duties dictated in the Gant Chart. 3-B Over Budget Detailing a set budget long beforehand so no expenses surprise the team. 2-C Fabricating Parts Late Submit parts to the shop far ahead of schedule to allow for unexpected delays in fabrication. 3-C Ordering Unnecessary/Incorrect Parts Double check the dimensions and necessity of each part with the project lead prior to submission of each part. 3-C 4.5 NAR/TRA Procedures All of the NAR procedures for launch and design of the rocket will be strictly followed. These procedures are included in Appendix F. At each launch the Range Safety Officer will have the final say on the conditions and safety of the launch. Hazardous materials such as rocket motors and black powder for separation charges will be handled by qualified personnel only. All hazardous materials will be stored in a locked and fireproof container in the lab. 4.6 Safety Brief Prior to beginning the construction phase of the project, a safety brief will be given to the entire team to ensure that all team members are knowledgeable of the risks inherent to the project. The 46

56 brief will also include the methods and procedures that will be utilized to mitigate these risks. The brief will include items such as the proper personal protective equipment (PPE) for manufacturing equipment (glasses, gloves, hearing protection). The brief will also dictate to always work with a partner if a situation is potentially hazardous. The brief will also state that everyone should read and heed warning labels on any materials or equipment before use. 4.7 Pre-Launch Brief Before each launch a brief will be conducted to ensure the launch is conducted using the best safety practices. The pre-launch brief will include the following: a. Launch Goals Items to be tested this launch b. Performance Predictions Predicted altitude and performance of the rocket c. Conditions Weather, crowds, or any other factors that may affect the launch d. General Safety Safe distance and conduct during launch e. Safety and Accident Procedures Actions to take in case of a misfire or recovery failure f. Systems Check Ensure the avionics are properly charged, the recovery systems are properly packaged, and the launching system is behaving nominally g. Launch and recovery h. Post Launch Brief Review any notes or improvements that could be made to safety 4.8 Legal Consideration All of the FAA rules and regulations for the launch of high powered rockets will be followed at all times. Each of the following laws and rules will also be followed. The text of these scripts are included in Appendix G. Federal Aviation Regulations 14 CFR, Subchapter F, Part 101, Subpart C Amateur Rockets, Code of Federal Regulation 27 Part 55: Commerce in Explosives Fire prevention, NFPA 1127 Code for High Power Rocket Motors. 47

57 All launches will be done at MDRA facilities. These facilities comply with the FAA regulations. While at these facilities, the team will cooperate and obey with any suggestions or recommendations from the range safety officer (RSO). 4.9 Material Safety Data Sheets Many hazardous materials will be used for the project. Each material will be stored safely, and a material safety data sheet (MSDS) will be printed and stored alongside them. Each person using the material will be required to read and understand the MSDS sheets. The MSDS sheets for hazardous materials are included in Appendix H. 48

58 5 P a y l o a d C r i t e r i a 5.1 Payload Objective The objective of the payload is to successfully deploy a rover that will travel at least five feet away from the rocket and deploy a foldable solar panel. This experiment will provide information on the feasibility of deploying inexpensive autonomous rovers to perform a specific task. The success of this experiment will be determined by the complete deployment of the rover from the interior of the rocket, its ability to travel beyond five feet from the nearest section of the rocket, and a successful deployment of deployable solar panels. 5.2 Payload Systems Design Review The original design had four wheels with a thin body and is functional in either orientation. This design was mainly chosen for its simplicity and reliability. A four wheel rover is a very common design that has many resources available which help in the design process. Creating models and feedback controllers for four wheel car systems is common and simple to do. Additionally, the four-wheel design offers improved stability and traction which is an advantage over other body designs. Stability and traction provides a decreased chance that the rover becomes stuck after deployment in an adverse environment. The disadvantages of this design are that space is limited and providing a safe and effective deployment method is extremely difficult. In order for the rover to be operational both right side up and upside down, the wheels must be large and the body must be thin. The body would need to be long in order to accommodate the sensors and actuators on the rover. During flight the rover must be safely secured and reliably deployed after landing in a specific orientation; this requires several redundant mechanisms. The orientation of the rover must be controlled to ensure that it exits the body of the launch vehicle correctly, and the sensors and solar cell must be placed on the body of the rover in a way that is effective in either orientation. These restraints would be difficult to overcome, especially with the limited available space on the rover body. This design would also have excessive weight that could otherwise be used for other systems. 49

59 A spherical ball design was discussed because of its unique advantages of being omnidirectional and self-orienting. These advantages eliminate issues with controlling body orientation during deployment, because after leaving the body of the launch vehicle the rover will roll into the correct orientation. The disadvantages of this design are the required structural strength of the sphere, restrictive sensor positioning, and difficulty navigating across unpredictable terrain. The sphere casing runs the risk of collapsing due to in flight or landing forces if not constructed with sturdy material or an internal support system. Both of these options add significant weight and complexity. If the sphere material is too dense, many sensors that would be used on a rover would not be capable of propagating a signal beyond the walls of the rover. Radio direction finder (RDF) sensors, for example, would not be able to read distance from the rocket internally, and there would be no location to mount them externally on a sphere. Another notable disadvantage for this design is that it would have a difficult time navigating certain terrains due to the lack of a multiple wheel drive system. This risks the rover becoming stuck and not able to get proper traction. The final proposed design utilized two wheels and a single axle. This design has two DC motors on either end of the chassis, one for each wheel. It would be properly weighted on the underside of the axle to prevent the body from rotating without turning the wheels. The advantages of the two wheel design are decreased weight, safer and more reliable deployment options, and selforienting abilities. Because there are only two custom wheels, the tires can be larger and more robust but still lighter than a traditional four wheel rover design. For deployment, the two wheel rover has several different deployment options that are more secure and reliable than the four wheel rover. The single axle design provides high structural integrity, as there are no outlying moment arms that could cause undesired torque on the system. The weighted chassis provide its self-orienting ability, this is an advantage of the design. If the main body of the rover is weighted enough on one side, the moment created by gravity will align the chassis with the proper frame. The weight will also serve as a counterweight which will allow the rover to move. 50

60 5.3 System Design Research The advantages of the four wheel design are that it is easy to make and is very reliable on any surface [3], and contact avoidance can be easily implemented [2],[3]. The disadvantage to this design is that it requires significant vertical space which requires one of two options. The first option is to create a deployment system capable of correctly orienting the rover within the body of the launch vehicle before its deployment. The complexity of this system would increase the risk of failure enough to be removed from the list as a possible design The second option is to create a longer and flatter rover that has a body thinner than the diameter of the tires and is capable of operating in both orientations. This design would require more weight for a longer body and would be structurally weak in the middle, which could cause it to break before it ever leaves the launch vehicle. This design has advantages, however, it requires a complex deployment system and would not have the structural integrity to be chosen as the main design. Due to the complexities and weaknesses of this design it will not be used as the leading payload design. The design for a spherical rover was considered as an option because it is self-orienting, which eliminates many complications for deployment. Examples of this kind of rover are shown in reference [3]. This idea provides another unique advantage which is that after it has been deployed, the rover has the capability to travel in any direction. The disadvantages to this design are that it is not extremely stable, it is hard to control, and the sphere would limit the types of sensors that could be used to determine its distance from surrounding objects. In all of the examples shown in reference [3], the rovers are displayed moving over flat terrain, whereas this mission s terrain is unpredictable. It would be more structurally fragile than either of the other alternative designs. The two wheel design combines the advantages of the spherical design as well as unique advantages of its own. An example of this style of rover is shown in reference [4]. It will have treaded rubber wheels and the ability to navigate on difficult terrain like the four wheel design, and it will be self-orienting similar to the spherical design. Its unique advantages are that even though it will have hemispherical wheels it will not be enclosed and the same sensors that would be used on the four wheel design can still be used for range finding. The design will have two 51

61 actuators, making it easier to control than the spherical rover and allowing it to make sharper turns than the four wheel design. The disadvantages to this design are that there is limited space on the body of the rover for the sensors and actuators, and it will be more difficult to create models and an effective feedback controllers for a two wheel vehicle rather than four. Although there is not very much research on a two wheel rover used in this capacity, the proof of concept can be seen via online videos [4]. 5.4 Leading Payload Design The final design for the rover is the two wheel rover model. It will have two large hemispherical wheels to ensure that it can navigate tough terrain and roll if it lands on its side. These wheels will be hollowed slightly to allow the chassis to fit into the wheel providing more space for actuators and sensors. The wheels will be actuated by two DC motors mounted onto the axle itself, and connected to the wheels through a gearing system. The rover will have a battery mounted on the bottom of the axle in order to provide unbalanced weight which will be used for orientation. Due to the volatility of Lithium batteries when they are pierced, a set of nine volt Nickel Metal Hydride batteries will be used instead. The sensors that will be mounted on top of the chassis will be a nine axis Inertial Measurement Unit (IMU), a radio transceiver, a sound navigation and ranging (SONAR) distance sensor, DC motor encoders, and a foldable solar cell attached to a small servo. The IMU is to determine the direction of the force of gravity to ensure that the rover body has reached proper orientation before attempting to navigate. The radio transceiver will be utilized to activate the rover after deployment from the main body of the launch vehicle through a command from a computer. It will then autonomously drive straight until it leaves the payload body s immediate vicinity. The distance sensing will then shift over to the SONAR sensor to determine the distance from anything near it. After the rover has determined that it is at least five feet away from any part of the body of the launch vehicle, it will rotate a full 360 degrees to ensure proper clearance on all sides. After verification the rover will then complete its mission by deploying the solar cell. The solar cell will be placed on top of the axle of the rover and will be folded in half on itself. The bottom half will be firmly secured, while the top half is left hinged. Two simple servos will be used to rotate the unfixed half open like a flap. All of these systems will be wired utilizing a printed circuit board, to ensure strong 52

62 connections of wires. This circuit board will be mounted on the bottom of the axle to eliminate clutter on its topside. Figure 25 is the schematic for the leading rover design. The estimated horizontal width of the rover axle is 5.75 inches, and the estimated weight of the entire rover is 2.0 pounds. Figure 25. Leading Payload Design. 53

63 A 3D representation of the rover is shown in Figure 26. Figure 26. 3D Model of SASOR. 5.5 Deployment Design Multiple deployment systems were considered for SASOR. The biggest consideration was the ability to secure the rover in flight. Originally, it was discussed to have the rover roll out of the rocket body under its own power. Also discussed was using a spring to forcefully propel the rover from the rocket section or a conveyer belt system to move the rover axially. All these deployment methods made securing the rover a major liability. Using hinged doors as a way to secure the rover was considered, however a hinged door would still allow the rover to move while inside the rocket body. If the door were to fail, there would be no secondary rover securement. An alternate approach was to rigidly fix the rover to the rocket body during flight and then release the rover upon landing. One option was to attach the rover to a robotic arm inside the 54

64 rocket body. Upon landing, the arm would extend and drop the rover outside the body tubing to complete its mission. This design added too much complexity and weight. Continuing with the same approach, a stepper motor with a lead screw was considered. The rover would have a nut that would thread onto the screw. The rover would be rigidly attached to the rocket and upon landing the stepper motor would be activated and would unscrew the rover out of the rocket. This design allowed for simplicity and rigidly attached the rover to the rocket. However, it still did not provide a failsafe in the event that the rover nut failed. Therefore, a faux bulkhead with a threaded nut was added to the design. The bulkhead would thread onto the lead screw after the rover. The bulkhead would hold the rover from falling out of the rocket body if the rover s nut failed and would protect the rover from external forces such as the separation charge. 5.6 Leading Deployment Design The leading deployment design is a system consisting of two rails and a stepper motor with a lead screw. The system will be controlled by a microprocessor, powered by a battery and a voltage regulator, and initiated by a signal sent to a radio frequency module, all of which will be mounted in the payload section of the launch vehicle separate from the rover. The two rails be mounted to the main body of the launch vehicle opposite to each other and will act as guides for the rover axle to slide on. When the stepper motor is actuated, the lead screw will pass through a threaded piece mounted on the rover s axle located near the horizontal center. As the lead screw passes through the threaded piece the rails will restrict the rover from rotating, which will allow the stepper motor to screw the rover into the main body. Therefore, after the rover has been threaded onto the lead screw the only way that the rover can move is by actuating the stepper motor. To secure the rover before flight the rover axle will be aligned with the guide rails and the threaded pass-through piece on the rover will be aligned with the lead screw. The stepper motor will be actuated and will screw the rover into the body for a specified number of rotations, which will be applied by the microprocessor and tracked using the stepper motor s encoder. A mobile bulkhead will be designed to have notches to fit into the guide rails and will be threaded so that it 55

65 can be screwed into the main body after the rover. This faux bulkhead will serve two purposes: it will make sure no debris can enter the payload section of the rocket during flight and landing, and in the event that the rover s deployment nut fails and the rover becomes loose during flight it serves as an extra safety precaution to deny the rover the ability to fall out of the payload section of the launch vehicle. After the rocket has landed, the deployment system will be remotely activated by sending an initiation message to the radio frequency module, which will initiate the deployment process. The stepper motor will turn the lead screw in the opposite direction that was used to to secure the rover and the bulkhead before flight. The mobile bulkhead will exit the launch vehicle first and fall to the ground, and the rover will follow. The stepper motor will spin for more revolutions than it did to secure the rover to ensure that the rover is successfully ejected from the deployment system. Figure 27 shows the schematic for the leading deployment method system. The estimated weight of the deployment system to include the stepper motor, lead screw, and both rails is 3.0 pounds. 56

66 Figure 27. Leading Deployment Method. The biggest failure concern in the deployment method is a lead screw failure. If this occurs, the entire deployment method will fail. As a result, in the initial subscale launch, an accelerometer will be part of the rocket s payload. The accelerometer will record the large accelerations from 57

67 launch, separation, parachute deployment, and landing. This data will be used to calculate axial forces imposed on the stepper motor and lead screw. The resulting forces will be used to determine if the lead screw can survive flight. 5.7 Payload and Launch Vehicle Interfacing The SASOR will interface with the payload section of the rocket through the stepper motor and lead screw. The stepper motor will be mounted to a secure bulkhead in the body of the payload section using four separate bolts and a spreader plate to reduce the stress felt by the bolts and the stepper motor. The SASOR will have a large nut incorporated into the center of the axle for mounting onto the screw. The axle of the rover will extend slightly beyond the edge of the wheels on both sides and will fit into the guide rails on either side of the rocket body. These will guide the SASOR in and out of the body of the payload section. There will also be a detachable bulkhead with a nut in the center attached to the end of the screw after the rover is mounted. This will prevent the rover from falling if its internal nut fails. It also acts as a matter of ensuring that the lead screw is extended parallel to the deck by keeping it evenly spaced on all sides. 58

68 6 P r o j e c t P l a n 6.1 Requirements Verification The team s results verification matrices are shown in this section. Note: References to NASA Student Launch Handbook requirements section will be refered to as N-SLH MO# MISSION OBJECTIVES POC MISSION MO-1 The vehicle will deliver the payload to an apogee altitude of 5280 feet AGL. N-SLH 2.1 Actual MO-2 MO-3 MO-4 MO-5 MO-6 The launch vehicle will be designed to be recoverable and reusable. Reusable is defined as being able to launch again on the same day without repairs or modifications. Teams will design a custom rover that will deploy from the internal structure of the launch vehicle. At landing, the team will remotely activate a trigger to deploy the rover from the rocket After deployment, the rover will autonomously move at least 5 ft. (in any direction) from the launch vehicle Once the rover has reached its final destination, it will deploy a set of foldable solar cell panels. N-SLH 2.6 Actual N-SLH Actual N-SLH Actual N-SLH Actual N-SLH Actual MSC-# MISSION SUCCESS CRITERIA SOURCE POC MISSION MSC-1 The launch vehicle will be capable of being prepared for flight at the launch site within 3 hours of the time the Federal Aviation Administration flight waiver opens. MO-1 N-SLH 2.9 Actual MSC-2 The launch vehicle will be capable of remaining in launch-ready configuration at the pad for a minimum of 1 hour without losing the functionality of any critical on-board components. MO-1 N-SLH 2.10 Actual MSC-3 The launch vehicle will have a minimum static stability margin of 2.0 at the point of rail exit. Rail exit is defined at the point where the forward rail button loses contact with the rail. MO-1 N-SLH 2.16 Actual MCS-4 All teams will successfully launch and recover a subscale model of their rocket prior to CDR. MO-ALL N-SLH 2.18 Actual MCS-5 All teams will successfully launch and recover their full-scale rocket prior to FRR in its final flight configuration. The rocket flown at FRR must be the same rocket to be flown on launch day. A successful flight is defined as a launch in which all hardware is functioning properly (i.e. drogue chute at apogee, main chute at a lower altitude, functioning tracking devices, etc.). MO-ALL N-SLH 2.19 Actual 59

69 CATEGORY REFERENCE # REQUIREMENT SOURCE POC VERIFICATION METHOD VERIFICATION PLAN General Requirements N-SLH 1.1 Students on the team will do 100% of the project, including design, construction, written reports,presentations, and flight preparation with the exception of assembling the motors and handling black powder or any variant of ejection charges, or preparing and installing electric matches (to be done by the team s mentor). N-SLH Lewis Demonstation N/A N-SLH 1.2 N-SLH 1.3 N-SLH 1.4 N-SLH 1.5 N-SLH 1.6 N-SLH 1.7 The team will provide and maintain a project plan to include, but not limited to the following items: project milestones, budget and community support, checklists, personnel assigned, educational engagement events, and risks and mitigations. N-SLH Siprelle Demonstation Foreign National (FN) team members must be identified by the Preliminary Design Review (PDR) and may or may not have access to certain activities during launch week due to security restrictions. In addition, FN s may be separated from their team during these activities. N-SLH Lewis N/A The team must identify all team members attending launch week activities by the Critical Design Review (CDR). Team members will include: Students actively engaged in the project throughout the entire year. One mentor. No more than two adult educators. N-SLH Mariano Analysis The team will engage a minimum of 200 participants in educational, hands-on science, technology, engineering, and mathematics (STEM) activities, as defined in the Educational Engagement Activity Report, by FRR. An educational engagement activity report will be completed and submitted within two weeks after completion of an event. A sample of the educational engagement activity report can be found on page 31 of the handbook. To satisfy this requirement, all events must occur between project acceptance and the FRR due date. N-SLH Mullen Demonstation The team will develop and host a Web site for project documentation. N-SLH Mullen Demonstation Teams will post, and make available for download, the required deliverables to the team Web site by the due dates specified in the project timeline. N-SLH Mullen Demonstation The project manager will continually show updates to the project mentor. The team has no Foreign Nationals. Team members will all be attending launch week contingent on budget analysis. The team will participate in a large scale steam activity that will engage more than the required participants. The team's PAO will regularly meet with the project manager to show web site functionality and updates. The team's PAO will regularly meet with the project manager to show web site functionality and updates. N-SLH 1.8 All deliverables must be in PDF format. N-SLH Mariano Inspection N-SLH 1.9 N-SLH 1.10 In every report, teams will provide a table of contents including major sections and their respective sub-sections. N-SLH Lewis Inspection In every report, the team will include the page number at the bottom of the page. N-SLH Lewis Inspection All deliverables will go through the Project Manager who will inspect the file format before submission. The project manager will inspect reports for table of contents. The project manager will inspect reports for page numbers. 60

70 N-SLH 1.11 The team will provide any computer equipment necessary to perform a video teleconference with the review panel. This includes, but is not limited to, a computer system, video camera, speaker telephone, and a broadband Internet connection. Cellular phones can be used for speakerphone capability only as a last resort. N-SLH Nowotney Demonstation Logistics will ensure the team's teleconference room is ready for all presentations. N-SLH 1.12 N-SLH 1.13 N-SLH 1.14 All teams will be required to use the launch pads provided by Student Launch s launch service provider. No custom pads will be permitted on the launch field. Launch services will have 8 ft rails, and 8 and 12 ft rails available for use. N-SLH Moore Inspection Teams must implement the Architectural and Transportation Barriers Compliance Board Electronic and Information Technology (EIT) Accessibility Standards (36 CFR Part 1194) Subpart B-Technical Standards ( Software applications and operating systems Web-based intranet and Internet information and applications. N-SLH Gomez Demonstation Each team must identify a mentor. A mentor is defined as an adult who is included as a team member, who will be supporting the team (or multiple teams) throughout the project year, and may or may not be affiliated with the school, institution, or organization. The mentor must maintain a current certification, and be in good standing, through the National Association of Rocketry (NAR) or Tripoli Rocketry Association (TRA) for the motor impulse of the launch vehicle and must have flown and successfully recovered (using electronic, staged recovery) a minimum of 2 flights in this or a higher impulse class, prior to PDR. The mentor is designated as the individual owner of the rocket for liability purposes and must travel with the team to launch week. One travel stipend will be provided per mentor regardless of the number of teams he or she supports. The stipend will only be provided if the team passes FRR and the team and mentor attends launch week in April. N-SLH Lewis Demonstation The team will ensure only the proper launch pads are used. Avionics will deomstrate completion of this requirment to the project manager. The project manager will regularly meet and maintian contact with the team mentor throughout the lifetime of the project. 61

71 CATEGORY REFERENCE # REQUIREMENT SOURCE POC VERIFICATION METHOD VERIFICATION PLAN Aerodynamics Drag N-SLH 2.1 N-SLH The launch vehicle will be designed to minimize drag. N-SLH Siprelle Analysis The subscale model will have an overall drag coeeficient the same as the full scale launch vehicle. N-SLH Siprelle Analysis Wind tunnel testing will be performed and analyzed to find areas to minimize drag. The model will be designed to match the drag of the analyzed full scale design. N-SLH 2.1 Drag coefficients on the launch vehicles will be accurately calculated. N-SLH Siprelle Analysis Wind tunnel tests will be compared to Open Rocket and RockSim to ensure accurate calculations. Stability N-SLH 2.1 N-SLH 2.16 Wind tunnel tests will be conducted on the launch vehicle. N-SLH Siprelle Test The launch vehicle will have a minimum static stability margin of 2.0 at the point of rail contact. N-SLH Siprelle Analysis The tests will be performed. Using wind tunnel tests and computer analysis, the stability margin will be calculated. Fins will be designed to meet this requirement. Misc N-SLH 2.17 The center of pressure will be accurately calculated. N-SLH Siprelle Analysis N-SLH 2.18 The center of gravity will be accurately calculated. N-SLH Siprelle Analysis N-SLH The launch vehicle will not utilize forward canards. N-SLH Siprelle Inspection Wind tunnel tests will be compared to Open Rocket and RockSim to ensure accurate calculations. The calculated CG location will be compared to the CG of the actual rocket. Inspection will ensure no forward canards are used. N-SLH The launch vehicle will not exceed Mach 1 at any point during flight. N-SLH Siprelle Analysis Flight simulations using Open Rocket and RockSim will help select a motor that does not propell the rocket beyond Mach 1. 62

72 CATEGORY REFERENCE # REQUIREMENT SOURCE POC VERIFICATION METHOD VERIFICATION PLAN Structures N-SLH Bulkheads remain fixed to inside of rocket at all times N-SLH Moore, Nowotny Test Roket body sections will go through a series of drop tests to ensure bulkheads remain rigidly fixd to the body tubing. N-SLH N-SLH Rocket fins remain attached to the rocket body throughout flight and landing. N-SLH Moore, Nowotny Demonstrate Rocket body remains intact throught flight and landing. N-SLH Moore, Nowotny Testing Roket body fin section will go through a series of drop tests to fins remain rigidly fixd to the body tubing. Roket body sections will go through a series of drop tests to ensure the tubing does not fail. N-SLH N-SLH Expoxy/resin is applied such that no bubbles are visible. Internal Moore, Nowotny Inspection Parachute attachment point stays rigidly attached throughout flight. Internal Moore, Nowotny Analysis Expoxied joins will be visually inspected. Materials used for parachute attachment will be analyzed to ensure they will remain intact. 63

73 CATEGORY REFERENCE # REQUIREMENT SOURCE POC VERIFICATION METHOD VERIFICATION PLAN RECOVERY N-SLH The parachute shock cord must be folded and tucked completely separate from the parachute shock cords. N-SLH Mullen Inspection Inspect the parachute and shock chord before launch. N-SLH 2.19 The launch vehicle will stage the deployment of its recovery devices, where a drogue parachute is deployed at apogee and a main parachute is deployed at a lower altitude. Tumble or streamer recovery from apogee to main parachute deployment is also permissible, provided that kinetic energy during drogue-stage descent is reasonable, as deemed by the RSO. N-SLH Mullen Inspection Black powder tests will be performed before the launch. N-SLH 3.2 Each team must perform a successful ground ejection test for both the drogue and main parachutes. This must be done prior to the initial subscale and full-scale launches N-SLH Mullen Demonstration Ejection tests will be conducted before the launch. N-SLH 3.7 Motor ejection is not a permissible form of primary or secondary deployment. N-SLH Mullen N/A N/A N-SLH 3.8 N-SLH 3.9 MO-2 MO-2 Removable shear pins will be used for both the main parachute compartment and the drogue parachute compartment. N-SLH Mullen Analysis Recovery area will be limited to a 2500 ft. radius from the launch pads. N-SLH Mullen Analysis The parachute and shock cords must have inherent tensile strengths capable of withstanding 1.5 times the anticipated loads they will experience during flight N-SLH Mullen Analysis The parachute and shock cords must be securely attached to the rocket with the use of a trilene knot. N-SLH Mullen Inspection Shear pin calculations will determined before testing. Drift calculations will be determined. Strength tests will be conducted on the parachute and shock chord. The knots will be inspected. MIDN Mullen MIDN Mullen MIDN Mullen MIDN Mullen MIDN Mullen MIDN Mullen MIDN Mullen MIDN Mullen The main parachute shall have a small circular hole cut in the center of the parachute with a hole area not to exceed 10 percent of the total main parachute surface area. N-SLH Mullen Inspection The connections between the main parachute cords and the main parachute shall be reinforced with a light coat of epoxy. N-SLH Mullen Inspection Black powder charges not to exceed 15g shall initiate the rocket subsection separation and parachute deployment. N-SLH Mullen Analysis All shock cord protectors shall be fixed with canvas protectors which cover the shock cord completely. N-SLH Mullen Inspection A canvas sheet with a minimum surface area of 1 square foot shall be placed beneath the main parachute, drouge parachute, and shock cord as a means to shield recovery components from the heat of the motor. This canvas sheet will be intertwined with and attached to the shock cord. N-SLH Mullen Inspection The black powder charges shall have suficient force to shear the 4 shearing pins installed prior to every flight. N-SLH Mullen Analysis The main parachute shall be packed in such a way that facilitates a quick, reliable, and efficient manner of opening upon being ejected from the internal structure of the rocket. N-SLH Mullen Inspection The recovery system components shall not interfere with the packing and merging of rocket subsections. N-SLH Mullen Analysis The hole will be thoroughly measured and cut exactly. The recovery team will ensure that the connections are secure. Black powder tests will be performed before the launch. The recovery team will ensure that the canvas protectors are secured. The recovery team will properly place the entire recovery subsytem into the rocket. Black powder tests will be performed before the launch. Proper packing procedures will be followed for the parachute. The recovery team will work with the other teams to make sure spacing is consistant. 64

74 CATEGORY REFERENCE # REQUIREMENT SOURCE POC VERIFICATION METHOD VERIFICATION PLAN Avionics N-SLH 2.2 One commercially availible barometric altimeter will be used to measue the altitude of the rocket for scoring N-SLH Gomez, Davidson Inspection The avionics team will include a altimeter in the avionics bay. N-SLH 2.3 Each altimeter must have a dedicated arming switch that is accessible from the exterior of the airframe when on the launch pad N-SLH Gomez, Davidson Inspection The design of the avionics bay will be easily accessible. N-SLH 2.4 Each altimeter will have a dedicated power supply N-SLH Gomez, Davidson Inspection Avionics will design the altimeters to have their own power supply. N-SLH 2.5 Each altimeter arming switch must be able to be locked in the ON position at launch N-SLH Gomez, Davidson Demonstration The avionics team will demonstrate that the key can be locked in the on position. N-SLH 2.1 N-SLH 3.1 The rocket must be able to sit on the launch pad for a minimum of 1 hour without losing any functionality N-SLH Gomez, Davidson Analysis Altimeters deploy the seperation charge for the main chute deployment at apogee N-SLH Gomez, Davidson Analysis N-SLH 3.1 Altimeters must deploy the main chute at 800 ft N-SLH Gomez, Davidson Testing N-SLH 3.4 N-SLH 3.5 N-SLH 3.6 N-SLH 3.1 N-SLH N-SLH 3.11 N-SLH 3.11 Recovery electronics will be independent of payload electrical circuits N-SLH Gomez, Davidson Inspection Recovery electronics will be powered by commercailly availible batteries N-SLH Gomez, Davidson Inspection Recovery system will contain redudant, commercially avaible altimeters N-SLH Gomez, Davidson Inspection Electronic tracking device must be attached to the launch vehicle and transmit the position to the ground reciever N-SLH Gomez, Davidson Inspection Electronic tracking device must be fully functional on the day of launch N-SLH Gomez, Davidson Analysis Recovery altimeters must be located in a separate, shielded compartment from any other electronics so that they cannot be interfered with N-SLH Gomez, Davidson Analysis Hole in separated chamber of altimeter to ensure proper altitude is recorded N-SLH Gomez, Davidson Demonstration All possible outcomes will be analyzed for the rocket sitting on the launch pad. A test will be conducted to make sure the coding runs correctly for the altimeters. Jolly Logics will be tested to ensure proper use. Th design of the subsystem will ensrue independence. The rover team will make sure commercially bought batteries are available. The rover team will make sure commercially bought altimeters are used. GPS tests will be conducted prior to full scale launch. Testing will be conducted the week leading up to testing. Multiple designs will be considered prior to a final selection being made. The avionics team will drill a hole in the body of the rocket. 65

75 CATEGORY REFERENCE # REQUIREMENT SOURCE POC VERIFICATION METHOD VERIFICATION PLAN Rover MO-5 The rover will operate under its own power. N-SLH Holley, Medina MO-5 The body of the rover will maintain its structural integrity for use during multiple flights. N-SLH Holley, Medina Demonstration Test Rover team will demonstrate the rover operating under its own power. Drop tests will be performed to ensure the rover is capable of withstanding in-flight forces. MO-5 The rover will remain secured during flight. N-SLH Holley, Medina MO-5 MO-5 The rover will be capable of sending and receiving remote signals. After exiting the main body of the rocket, the rover will be able to detect and track its position in relation to the rocket. N-SLH N-SLH Holley, Medina Holley, Medina MO-5 Rover will be able to navigate unpredictable terrain. N-SLH Holley, Medina MO-6 Solar cells will be deployed in proper orientation. N-SLH Holley, Medina MO-5 After receiving a remote signal the rover will operate autonomously. N-SLH Holley, Medina Test Demonstration Demonstration Test Demonstration Demonstration Drop tests will be performed to ensure the rover remains secured througout flight. Rover team will demonstrate the rover sending and receiving signals. Rover team will demonstrate the rover tracking its position in relation to the rocket body. Rover mobility will be tested in varying terrain. Solar cell deployment will be demonstrated. Autonomous operation will be demonstrated. 6.2 Timeline Displayed in Figure 28 is the project timeline. It shows the entire process from project proposal to post launch assessment review. 66

76 Figure 28. Project Timeline Gantt Chart. 67

77 6.3 Budget and Funding Currently Navy Rockets has one anticipated source of income for the USLI. The team has historically received funding from the USNA, and the requested amount for the fiscal year was $21,000. The funds released for the November-December phase was $4000. The preliminary budget has been designed around an assumed funding level of $21,000. The expected expenditures of the USLI competition are listed in Table 13. These costs are presented in depth in Table 14 and Table 15 with an itemized budget of the full scale design and of competition related expenses. Table 13. Navy Rockets USLI Expected Costs. Full Scale $4, Subscale $1, Testing and $ Development Support $1, Travel and Lodging $10, Outreach $ Total $17,

78 Table 14. Navy Rockets Full Scale Itemized Budget. Subsystem Item Unit Cost Quantity Total Cost Rocket Structure 6 diameter carbon fiber tube $ $ Resin $ $82.05 Hardener $ $33.35 Internal Structures and Payload $ $ Bay Miscellaneous Fasteners $ $50.00 Avionics and Trackimos GPS $ $ Recovery PerfectFlite Stratologger $ $ Altimeter Fruity Chutes 60 Standard $ $ Parachute Fruity Chutes 96 Standard $ $ Parachute 24 Elliptical Parachute $ $ fiberglass tube $ $ Tubular White Nylon 1" $ $57.50 Miscellaneous Fasteners and $ $35.00 Components Propulsion Cesaroni L800-P Classic Rocket $ $ motor 75mm Motor Casing $ $ Rover Solar panels $ $20.00 Body frame $ $ Battery $ $20.00 Control computer $ $47.50 Total $4,

79 Table 15. Navy Rocket s Travel Budget. Subsystem Item Unit Cost Quantity Total Cost Lodging Hotel Rooms (per room, per $ $2, night) Travel Plane Fares $ $5, RentNal Car Rental Car (per day) $ $ Per Diem Per Diem (per person, per day) $ $1, $10, Material Acquisition Plan The subscale vehicle will be constructed by 17 November, The required parts are contained in the Navy Rocket Team inventory. No parts will be purchased for the subscale vehicle. The full scale launch vehicle will be completed by 2 February, To complete this building process, parts will be ordered from online sources using the procedure outlined below. Any part that must be purchased will receive the approval of the Project Manager and the Project Mentor. Following approval, an acquisitions request form will be created and submitted to the Naval Academy Aerospace Engineering Purchase Card Authority. Following approval, the parts will be ordered and charged to the Navy Rocket Team account. The time between acquisitions request and part delivery is expected to be between three and six weeks. All mission critical parts will be ordered a minimum of six weeks before that part is required for the vehicle assembly. All parts required for the full scale vehicle build will be ordered no later than 15 December,

80 R e f e r e n c e s [1] Schenker, Paul S., Terrance L. Huntsberger, Paolo Pirjanian, Eric T. Baumgartner, Hrand Aghazarian, Ashitey Trebi-Ollennu, Patrick C. Leger, Yang Cheng, Paul G. Backes, Edward Tunstel, Steven Dubowsky, Karl D. Iagnemma, and Gerard T. Mckee. "." Intelligent Robots and Computer Vision XX: Algorithms, Techniques, and Active Vision, doi: / [2] Medina, Alberto & Mollinedo, Luis & Kapellos, Konstantinos & Crespo, Carlos & Poulakis, Pantelis. (2015). Design and Realization of a Rover Autonomy Testbed. [3] Crossley, Vincent A., Department of Mechanical Engineering Carnegie Mellon University, "A Literature Review on the Design of Spherical Rolling Robots." [4] How to make Self Balancing R/C robot; [Online]; October 21, H9Dj0lSb8 [5] NFPA Code 495, Explosives Materials Code, National Fire Protection Association, 1 Batterymarch Park, Quincy, MA [6] NFPA Code 1122, Code for Model Rocketry. NFPA Code 1127, Code for High Power Rocketry. Code of Federal Regulations, Title 14, Part 101, Federal Aviation Regulations by the FAA for unmanned rockets. [7] Code of Federal Regulation, Title 16, Part (a)(8), Consumer Product Safety Commission exemption for model rockets. [8] Code of Federal Regulations, Title 27, Part 55, Bureau of Alcohol, Tobacco, and Firearms regulations. 71

81 [9] Code of Federal Regulations, Title 49, Parts , Department of Transportation hazardous material shipping regulations. [10] Model Rocket Safety Code, National Association of Rocketry. [11] High Power Rocketry Safety Code, National Association of Rocketry. [12] How to Size Ejection Charge; [Online]; Huntsville Area Rocket Association. 72

82 A P P E N D I X A : M i l e s t o n e R e v i e w F l y s h e e t Milestone Review Flysheet Institution United States Naval Academy Milestone PDR Total Length (in) Diameter (in) Vehicle Properties Motor Properties Motor Brand/Designation Cesaroni/L800 Max/Average Thrust (lb.) / Gross Lift Off Weigh (lb.) Airframe Material(s) Fin Material and Thickness (in) Coupler Length/Shoulder Length(s) (in) 31.7 Carbon Fiber, Fiberglass 1/4 Plywood w/ Fiberglass 13/6 Total Impulse (lbf-s) Mass Before/After Burn (lb.) 7.74/3.78 Liftoff Thrust (lb.) Motor Retention Method Motor Sleeve with Aeropac Retainer Velocity at Deployment (ft/s) Terminal Velocity (ft/s) Recovery Harness Material Recovery Harness Size/Thickness (in) Harness/Airframe Interfaces Kinetic Energy of Each Section (Ft-lbs) Recovery Harness Length (ft) Altimeter(s)/Timer(s) (Make/Model) Redundancy Plan and Backup Deployment Settings Stability Analysis Center of Pressure (in from nose) Center of Gravity (in from nose) Static Stability Margin (on pad) Static Stability Margin (at rail exit) Thrust-to-Weight Ratio Rail Size/Type and Length (in) Rail Exit Velocity (ft/s) Recovery System Properties Manufacturer/Model Size/Diameter (in or ft) Altitude at Deployment (ft) Pad Stay Time (Launch Configuration) Drogue Parachute Fruity Chutes Section 1 Section 2 Section 3 Section 4 Recovery Electronics inches /53.00 Tubular Nylon Steel U-bolt, steel quick link, tubular nylon 4 PerfectFlight Two altimeters Stratologgers wired in parallel with main and redundant black powder charge. Both altimeters will be capable of detonating both N/A N/A 1 hour /144in Altitude at Deployment (ft) 800 Velocity at Deployment (ft/s) Kinetic Energy of Each Section (Ft-lbs) Maximum Velocity (ft/s) Maximum Mach Number Rocket Locators (Make/Model) Transmitting Frequencies (all - vehicle and payload) Energetics Mass - Drogue Chute Lower (grams) Energetics Mass - Main Chute (grams) Section 1 Section 2 Section 3 Section N/A N/A Recovery Electronics Primary Backup Primary Backup Primary Backup 46.31/53.00 Terminal Velocity (ft/s) 18.52/15.14 Recovery Harness Material Tubular Nylon Recovery Harness Size/Thickness (in) 1 Recovery Harness Length (ft) 12 Harness/Airframe Interfaces Ascent Analysis Maximum Acceleration (ft/s^2) Predicted Apogee (From Sim.) (ft) Recovery System Properties Manufacturer/Model Size/Diameter (in or ft) Main Parachute ction System Energetics (ex. Black Powd Energetics Masses - Drogue Chute Upper (grams) Fruity Chutes 60 inch/84 inch Steel U-bolt, steel quick link, tubular nylon Trakimo GPS ***Required by CDR*** Black Powder N/A N/A

83 Milestone Review Flysheet Institution United States Naval Academy Milestone PDR Payload Overview Payload 1 (official payload) The Single Axle Self Orienting Rover (SASOR) is a two wheeled rover that will be deployed from the Payload portion using a stepper motor with a lead screw and two guide rails. It will utilize an accelerometer and an ultrasonic distance sensor to determine distance from the body of the rocket and drive a minimum of five feet before deploying a foldable solar panel utilizing a small servo mounted in the body. Overview Payload 2 (non-scored payload) N/A Test Plans, Status, and Results Ejection Charge Tests Black powder testing is planned to be conducted on 6NOV17. It will test detonating black powder charges and the charge's ability to break nylon shear pins for separation. Sub-scale Test Flights Subscale rocket will be launched with scaled payload weight and motor size. An accelerometer will be attached to measure forces that the rover will experience in flight. It will provide information reguarding the survivability of the rover and lead screw. Full-scale Test Flights Full scale rocket will be launched with in order to ensure readiness for FRR. 74

84 Milestone Review Flysheet Institution United States Naval Academy Milestone PDR Additional Comments 75

85 A P P E N D I X B : O p e n R o c k e t D e s i g n w i t h D e t a i l e d P a r t s L i s t 76

86 77

87 78

88 79

FLIGHT READINESS REVIEW TEAM OPTICS

FLIGHT READINESS REVIEW TEAM OPTICS FLIGHT READINESS REVIEW TEAM OPTICS LAUNCH VEHICLE AND PAYLOAD DESIGN AND DIMENSIONS Vehicle Diameter 4 Upper Airframe Length 40 Lower Airframe Length 46 Coupler Band Length 1.5 Coupler Length 12 Nose

More information

Critical Design Review

Critical Design Review Critical Design Review University of Illinois at Urbana-Champaign NASA Student Launch 2017-2018 Illinois Space Society 1 Overview Illinois Space Society 2 Launch Vehicle Summary Javier Brown Illinois Space

More information

CRITICAL DESIGN REVIEW. University of South Florida Society of Aeronautics and Rocketry

CRITICAL DESIGN REVIEW. University of South Florida Society of Aeronautics and Rocketry CRITICAL DESIGN REVIEW University of South Florida Society of Aeronautics and Rocketry 2017-2018 AGENDA 1. Launch Vehicle 2. Recovery 3. Testing 4. Subscale Vehicle 5. Payload 6. Educational Outreach 7.

More information

NASA SL - NU FRONTIERS. PDR presentation to the NASA Student Launch Review Panel

NASA SL - NU FRONTIERS. PDR presentation to the NASA Student Launch Review Panel NASA SL - NU FRONTIERS PDR presentation to the NASA Student Launch Review Panel 1 Agenda Launch Vehicle Overview Nose Cone Section Payload Section Lower Avionic Bay Section Booster Section Motor Selection

More information

NASA USLI PRELIMINARY DESIGN REVIEW. University of California, Davis SpaceED Rockets Team

NASA USLI PRELIMINARY DESIGN REVIEW. University of California, Davis SpaceED Rockets Team NASA USLI 2012-13 PRELIMINARY DESIGN REVIEW University of California, Davis SpaceED Rockets Team OUTLINE School Information Launch Vehicle Summary Motor Selection Mission Performance and Predictions Structures

More information

Auburn University. Project Wall-Eagle FRR

Auburn University. Project Wall-Eagle FRR Auburn University Project Wall-Eagle FRR Rocket Design Rocket Model Mass Estimates Booster Section Mass(lb.) Estimated Upper Section Mass(lb.) Actual Component Mass(lb.) Estimated Mass(lb.) Actual Component

More information

Illinois Space Society Flight Readiness Review. University of Illinois Urbana-Champaign NASA Student Launch March 30, 2016

Illinois Space Society Flight Readiness Review. University of Illinois Urbana-Champaign NASA Student Launch March 30, 2016 Illinois Space Society Flight Readiness Review University of Illinois Urbana-Champaign NASA Student Launch 2015-2016 March 30, 2016 Team Managers Project Manager: Ian Charter Structures and Recovery Manager:

More information

Presentation Outline. # Title # Title

Presentation Outline. # Title # Title CDR Presentation 1 Presentation Outline # Title # Title 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 Team Introduction Vehicle Overview Vehicle Dimensions Upper Body Section Payload

More information

Georgia Tech NASA Critical Design Review Teleconference Presented By: Georgia Tech Team ARES

Georgia Tech NASA Critical Design Review Teleconference Presented By: Georgia Tech Team ARES Georgia Tech NASA Critical Design Review Teleconference Presented By: Georgia Tech Team ARES 1 Agenda 1. Team Overview (1 Min) 2. 3. 4. 5. 6. 7. Changes Since Proposal (1 Min) Educational Outreach (1 Min)

More information

Project NOVA

Project NOVA Project NOVA 2017-2018 Our Mission Design a Rocket Capable of: Apogee of 5280 ft Deploying an autonomous Rover Vehicle REILLY B. Vehicle Dimensions Total Length of 108 inches Inner Diameter of 6 inches

More information

Presentation Outline. # Title

Presentation Outline. # Title FRR Presentation 1 Presentation Outline # Title 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 Team Introduction Mission Summary Vehicle Overview Vehicle Dimensions Upper Body Section Elliptical

More information

Jordan High School Rocketry Team. A Roll Stabilized Video Platform and Inflatable Location Device

Jordan High School Rocketry Team. A Roll Stabilized Video Platform and Inflatable Location Device Jordan High School Rocketry Team A Roll Stabilized Video Platform and Inflatable Location Device Mission Success Criteria No damage done to any person or property. The recovery system deploys as expected.

More information

Auburn University Student Launch. PDR Presentation November 16, 2015

Auburn University Student Launch. PDR Presentation November 16, 2015 Auburn University Student Launch PDR Presentation November 16, 2015 Project Aquila Vehicle Dimensions Total Length of 69.125 inches Inner Diameter of 5 inches Outer Diameter of 5.25 inches Estimated mass

More information

Preliminary Design Review. California State University, Long Beach USLI November 13th, 2017

Preliminary Design Review. California State University, Long Beach USLI November 13th, 2017 Preliminary Design Review California State University, Long Beach USLI November 13th, 2017 System Overview Launch Vehicle Dimensions Total Length 108in Airframe OD 6.17in. ID 6.00in. Couplers OD 5.998in.

More information

Flight Readiness Review

Flight Readiness Review Flight Readiness Review University of Illinois at Urbana-Champaign NASA Student Launch 2017-2018 Illinois Space Society 1 Overview Illinois Space Society 2 Launch Vehicle Summary Javier Brown Illinois

More information

CRITICAL DESIGN PRESENTATION

CRITICAL DESIGN PRESENTATION CRITICAL DESIGN PRESENTATION UNIVERSITY OF SOUTH ALABAMA LAUNCH SOCIETY BILL BROWN, BEECHER FAUST, ROCKWELL GARRIDO, CARSON SCHAFF, MICHAEL WIESNETH, MATTHEW WOJCIECHOWSKI ADVISOR: CARLOS MONTALVO MENTOR:

More information

UC Berkeley Space Technologies and Rocketry Preliminary Design Review Presentation. Access Control: CalSTAR Public Access

UC Berkeley Space Technologies and Rocketry Preliminary Design Review Presentation. Access Control: CalSTAR Public Access UC Berkeley Space Technologies and Rocketry Preliminary Design Review Presentation Access Control: CalSTAR Public Access Agenda Airframe Propulsion Payload Recovery Safety Outreach Project Plan Airframe

More information

GIT LIT NASA STUDENT LAUNCH PRELIMINARY DESIGN REVIEW NOVEMBER 13TH, 2017

GIT LIT NASA STUDENT LAUNCH PRELIMINARY DESIGN REVIEW NOVEMBER 13TH, 2017 GIT LIT 07-08 NASA STUDENT LAUNCH PRELIMINARY DESIGN REVIEW NOVEMBER TH, 07 AGENDA. Team Overview (5 Min). Educational Outreach ( Min). Safety ( Min) 4. Project Budget ( Min) 5. Launch Vehicle (0 min)

More information

Tacho Lycos 2017 NASA Student Launch Critical Design Review

Tacho Lycos 2017 NASA Student Launch Critical Design Review Tacho Lycos 2017 NASA Student Launch Critical Design Review High-Powered Rocketry Team 911 Oval Drive Raleigh NC, 27695 January 13, 2017 Table of Contents Table of Figures:... 8 Table of Appendices:...

More information

Team Air Mail Preliminary Design Review

Team Air Mail Preliminary Design Review Team Air Mail Preliminary Design Review 2014-2015 Space Grant Midwest High-Power Rocket Competition UAH Space Hardware Club Huntsville, AL Top: Will Hill, Davis Hunter, Beth Dutour, Bradley Henderson,

More information

Statement of Work Requirements Verification Table - Addendum

Statement of Work Requirements Verification Table - Addendum Statement of Work Requirements Verification Table - Addendum Vehicle Requirements Requirement Success Criteria Verification 1.1 No specific design requirement exists for the altitude. The altitude is a

More information

Illinois Space Society University of Illinois Urbana Champaign Student Launch Maxi-MAV Preliminary Design Review November 5, 2014

Illinois Space Society University of Illinois Urbana Champaign Student Launch Maxi-MAV Preliminary Design Review November 5, 2014 Illinois Space Society University of Illinois Urbana Champaign Student Launch 2014-2015 Maxi-MAV Preliminary Design Review November 5, 2014 Illinois Space Society 104 S. Wright Street Room 321D Urbana,

More information

The University of Toledo

The University of Toledo The University of Toledo Project Kronos Preliminary Design Review 11/03/2017 University of Toledo UT Rocketry Club 2801 W Bancroft St. MS 105 Toledo, OH 43606 Contents 1 Summary of Proposal... 6 1.1 Team

More information

PROJECT AQUILA 211 ENGINEERING DRIVE AUBURN, AL POST LAUNCH ASSESSMENT REVIEW

PROJECT AQUILA 211 ENGINEERING DRIVE AUBURN, AL POST LAUNCH ASSESSMENT REVIEW PROJECT AQUILA 211 ENGINEERING DRIVE AUBURN, AL 36849 POST LAUNCH ASSESSMENT REVIEW APRIL 29, 2016 Motor Specifications The team originally planned to use an Aerotech L-1520T motor and attempted four full

More information

Student Launch. Enclosed: Preliminary Design Review. Submitted by: Rocket Team Project Lead: David Eilken

Student Launch. Enclosed: Preliminary Design Review. Submitted by: Rocket Team Project Lead: David Eilken University of Evansville Student Launch Enclosed: Preliminary Design Review Submitted by: 2016 2017 Rocket Team Project Lead: David Eilken Submission Date: November 04, 2016 Payload: Fragile Material Protection

More information

PRELIMINARY DESIGN REVIEW

PRELIMINARY DESIGN REVIEW PRELIMINARY DESIGN REVIEW 1 1 Team Structure - Team Leader: Michael Blackwood NAR #101098L2 Certified - Safety Officer: Jay Nagy - Team Mentor: Art Upton NAR #26255L3 Certified - NAR Section: Jackson Model

More information

Tacho Lycos 2017 NASA Student Launch Flight Readiness Review

Tacho Lycos 2017 NASA Student Launch Flight Readiness Review Tacho Lycos 2017 NASA Student Launch Flight Readiness Review High-Powered Rocketry Team 911 Oval Drive Raleigh NC, 27695 March 6, 2017 Table of Contents Table of Figures... 9 Table of Appendices... 11

More information

Preliminary Design Review. Cyclone Student Launch Initiative

Preliminary Design Review. Cyclone Student Launch Initiative Preliminary Design Review Cyclone Student Launch Initiative Overview Team Overview Mission Statement Vehicle Overview Avionics Overview Safety Overview Payload Overview Requirements Compliance Plan Team

More information

Wichita State Launch Project K.I.S.S.

Wichita State Launch Project K.I.S.S. Wichita State Launch Project K.I.S.S. Benjamin Russell Jublain Wohler Mohamed Moustafa Tarun Bandemagala Outline 1. 2. 3. 4. 5. 6. 7. Introduction Vehicle Overview Mission Predictions Payload Design Requirement

More information

Flight Readiness Review Addendum: Full-Scale Re-Flight. Roll Induction and Counter Roll NASA University Student Launch.

Flight Readiness Review Addendum: Full-Scale Re-Flight. Roll Induction and Counter Roll NASA University Student Launch. Flight Readiness Review Addendum: Full-Scale Re-Flight Roll Induction and Counter Roll 2016-2017 NASA University Student Launch 27 March 2017 Propulsion Research Center, 301 Sparkman Dr. NW, Huntsville

More information

Overview. Mission Overview Payload and Subsystems Rocket and Subsystems Management

Overview. Mission Overview Payload and Subsystems Rocket and Subsystems Management MIT ROCKET TEAM Overview Mission Overview Payload and Subsystems Rocket and Subsystems Management Purpose and Mission Statement Our Mission: Use a rocket to rapidly deploy a UAV capable of completing search

More information

NASA - USLI Presentation 1/23/2013. University of Minnesota: USLI CDR 1

NASA - USLI Presentation 1/23/2013. University of Minnesota: USLI CDR 1 NASA - USLI Presentation 1/23/2013 2013 USLI CDR 1 Final design Key features Final motor choice Flight profile Stability Mass Drift Parachute Kinetic Energy Staged recovery Payload Integration Interface

More information

University Student Launch Initiative

University Student Launch Initiative University Student Launch Initiative HARDING UNIVERSITY Flight Readiness Review March 31, 2008 Launch Vehicle Summary Size: 97.7 (2.5 meters long), 3.1 diameter Motor: Contrail Rockets 54mm J-234 Recovery

More information

Critical Design Review Report

Critical Design Review Report Critical Design Review Report I) Summary of PDR report Team Name: The Rocket Men Mailing Address: Spring Grove Area High School 1490 Roth s Church Road Spring Grove, PA 17362 Mentor: Tom Aument NAR Number

More information

AUBURN UNIVERSITY STUDENT LAUNCH. Project Nova. 211 Davis Hall AUBURN, AL Post Launch Assessment Review

AUBURN UNIVERSITY STUDENT LAUNCH. Project Nova. 211 Davis Hall AUBURN, AL Post Launch Assessment Review AUBURN UNIVERSITY STUDENT LAUNCH Project Nova 211 Davis Hall AUBURN, AL 36849 Post Launch Assessment Review April 19, 2018 Table of Contents Table of Contents...2 List of Tables...3 Section 1: Launch Vehicle

More information

Flight Readiness Review March 16, Agenda. California State Polytechnic University, Pomona W. Temple Ave, Pomona, CA 91768

Flight Readiness Review March 16, Agenda. California State Polytechnic University, Pomona W. Temple Ave, Pomona, CA 91768 Flight Readiness Review March 16, 2018 Agenda California State Polytechnic University, Pomona 3801 W. Temple Ave, Pomona, CA 91768 Agenda 1.0 Changes made Since CDR 2.0 Launch Vehicle Criteria 3.0 Mission

More information

University of Illinois at Urbana-Champaign Illinois Space Society Student Launch Preliminary Design Review November 3, 2017

University of Illinois at Urbana-Champaign Illinois Space Society Student Launch Preliminary Design Review November 3, 2017 University of Illinois at Urbana-Champaign Illinois Space Society Student Launch 2017-2018 Preliminary Design Review November 3, 2017 Illinois Space Society 104 S. Wright Street Room 18C Urbana, Illinois

More information

NASA SL Critical Design Review

NASA SL Critical Design Review NASA SL Critical Design Review University of Alabama in Huntsville 1 LAUNCH VEHICLE 2 Vehicle Summary Launch Vehicle Dimensions Fairing Diameter: 6 in. Body Tube Diameter: 4 in. Mass at lift off: 43.8

More information

University of Notre Dame

University of Notre Dame University of Notre Dame 2016-2017 Notre Dame Rocketry Team Critical Design Review NASA Student Launch Competition Roll Control and Fragile Object Protection Payloads Submitted January 13, 2017 365 Fitzpatrick

More information

NASA Student Launch W. Foothill Blvd. Glendora, CA Artemis. Deployable Rover. November 3rd, Preliminary Design Review

NASA Student Launch W. Foothill Blvd. Glendora, CA Artemis. Deployable Rover. November 3rd, Preliminary Design Review 2017 2018 NASA Student Launch Preliminary Design Review 1000 W. Foothill Blvd. Glendora, CA 91741 Artemis Deployable Rover November 3rd, 2017 Table of Contents General Information... 9 1. School Information...

More information

NASA s Student Launch Initiative :

NASA s Student Launch Initiative : NASA s Student Launch Initiative : Critical Design Review Payload: Fragile Material Protection 1 Agenda 1. Design Overview 2. Payload 3. Recovery 4. 5. I. Sub-Scale Predictions II. Sub-Scale Test III.

More information

USLI Critical Design Report

USLI Critical Design Report UNIVERSITY OF MINNESOTA TWIN CITIES 2011 2012 USLI Critical Design Report University Of Minnesota Team Artemis 1/23/2012 Critical Design Report by University of Minnesota Team Artemis for 2011-2012 NASA

More information

NASA USLI Flight Readiness Review (FRR) Rensselaer Rocket Society (RRS)

NASA USLI Flight Readiness Review (FRR) Rensselaer Rocket Society (RRS) 2016-2017 NASA USLI Flight Readiness Review (FRR) Rensselaer Rocket Society (RRS) Rensselaer Polytechnic Institute 110 8th St Troy, NY 12180 Project Name: Andromeda Task 3.3: Roll Induction and Counter

More information

AUBURN UNIVERSITY STUDENT LAUNCH PROJECT NOVA II. 211 Davis Hall AUBURN, AL CDR

AUBURN UNIVERSITY STUDENT LAUNCH PROJECT NOVA II. 211 Davis Hall AUBURN, AL CDR AUBURN UNIVERSITY STUDENT LAUNCH PROJECT NOVA II 211 Davis Hall AUBURN, AL 36849 CDR January 10, 2019 Contents List of Tables...7 List of Figures...9 1 CDR Report Summary...12 1.1 Payload Deployable Rover...12

More information

The University of Toledo

The University of Toledo The University of Toledo Project Cairo Preliminary Design Review 10/08/2016 University of Toledo UT Rocketry Club 2801 W Bancroft St. MS 105 Toledo, OH 43606 Contents 1 Summary of Preliminary Design Review...

More information

SpaceLoft XL Sub-Orbital Launch Vehicle

SpaceLoft XL Sub-Orbital Launch Vehicle SpaceLoft XL Sub-Orbital Launch Vehicle The SpaceLoft XL is UP Aerospace s workhorse space launch vehicle -- ideal for significant-size payloads and multiple, simultaneous-customer operations. SpaceLoft

More information

CNY Rocket Team Challenge. Basics of Using RockSim 9 to Predict Altitude for the Central New York Rocket Team Challenge

CNY Rocket Team Challenge. Basics of Using RockSim 9 to Predict Altitude for the Central New York Rocket Team Challenge CNY Rocket Team Challenge Basics of Using RockSim 9 to Predict Altitude for the Central New York Rocket Team Challenge RockSim 9 Basics 2 Table of Contents A. Introduction.p. 3 B. Designing Your Rocket.p.

More information

Pegasus II. Tripoli Level 3 Project Documentation. Brian Wheeler

Pegasus II. Tripoli Level 3 Project Documentation. Brian Wheeler Pegasus II Tripoli Level 3 Project Documentation Brian Wheeler Contents: A. Design Overview B. Booster Construction C. Electronics Bay (Mechanical) Construction D. Nose Cone Construction E. Recovery System

More information

NORTHEASTERN UNIVERSITY

NORTHEASTERN UNIVERSITY NORTHEASTERN UNIVERSITY POST-LAUNCH ASSESSMENT REVIEW NORTHEASTERN UNIVERSITY USLI TEAM APRIL 27TH 2018 Table of Contents 1. Summary 2 1.1 Team Summary 2 1.2 Launch Summary 2 2. Launch Vehicle Assessment

More information

MASSACHUSETTS INSTITUTE OF TECHNOLOGY Department of Aeronautics and Astronautics

MASSACHUSETTS INSTITUTE OF TECHNOLOGY Department of Aeronautics and Astronautics MASSACHUSETTS INSTITUTE OF TECHNOLOGY Department of Aeronautics and Astronautics 16.00 Introduction to Aerospace and Design Problem Set #4 Issued: February 28, 2002 Due: March 19, 2002 ROCKET PERFORMANCE

More information

Presentation 3 Vehicle Systems - Phoenix

Presentation 3 Vehicle Systems - Phoenix Presentation 3 Vehicle Systems - Phoenix 1 Outline Structures Nosecone Body tubes Bulkheads Fins Tailcone Recovery System Layout Testing Propulsion Ox Tank Plumbing Injector Chamber Nozzle Testing Hydrostatic

More information

NASA SL Flight Readiness Review

NASA SL Flight Readiness Review NASA SL Flight Readiness Review University of Alabama in Huntsville 1 LAUNCH VEHICLE 2 Vehicle Overview Vehicle Dimensions Diameter: 6 fairing/4 aft Length: 106 inches Wet Mass: 41.1 lbs. Center of Pressure:

More information

Rocketry Projects Conducted at the University of Cincinnati

Rocketry Projects Conducted at the University of Cincinnati Rocketry Projects Conducted at the University of Cincinnati 2009-2010 Grant Schaffner, Ph.D. (Advisor) Rob Charvat (Student) 17 September 2010 1 Spacecraft Design Course Objectives Students gain experience

More information

Tripoli Rocketry Association Level 3 Certification Attempt

Tripoli Rocketry Association Level 3 Certification Attempt Tripoli Rocketry Association Level 3 Certification Attempt Kevin O Classen 1101 Dutton Brook Road Goshen, VT 05733 (802) 247-4205 kevin@back2bed.com Doctor Fill Doctor Fill General Specifications Airframe:

More information

Northwest Indian College Space Center USLI Critical Design Review

Northwest Indian College Space Center USLI Critical Design Review 2012-2013 Northwest Indian College Space Center USLI Critical Design Review Table of Contents, Tables, and Figures I.0 CDR Report Summary... 1 I.1 Team Summary... 1 I.2 Launch Vehicle Summary... 1 I.2a

More information

Notre Dame Rocketry Team. Flight Readiness Review March 8, :00 PM CST

Notre Dame Rocketry Team. Flight Readiness Review March 8, :00 PM CST Notre Dame Rocketry Team Flight Readiness Review March 8, 2018 2:00 PM CST Contents Overview Vehicle Design Recovery Subsystem Experimental Payloads Deployable Rover Payload Air Braking System Safety and

More information

Critical Design Review Report NASA Student Launch Florida International University American Society of Mechanical Engineers (FIU-ASME)

Critical Design Review Report NASA Student Launch Florida International University American Society of Mechanical Engineers (FIU-ASME) Critical Design Review Report 2014-2015 NASA Student Launch Florida International University American Society of Mechanical Engineers (FIU-ASME) Florida International University Engineering Center College

More information

HPR Staging & Air Starting By Gary Stroick

HPR Staging & Air Starting By Gary Stroick Complex Rocket Design Considerations HPR Staging & Air Starting By Gary Stroick 1. Tripoli Safety Code 2. Technical Considerations 3. Clusters/Air Starts 4. Staging 5. Summary 2 1. Complex High Power Rocket.

More information

SAE Baja - Drivetrain

SAE Baja - Drivetrain SAE Baja - Drivetrain By Ricardo Inzunza, Brandon Janca, Ryan Worden Team 11 Engineering Analysis Document Submitted towards partial fulfillment of the requirements for Mechanical Engineering Design I

More information

NUMAV. AIAA at Northeastern University

NUMAV. AIAA at Northeastern University NUMAV AIAA at Northeastern University Team Officials Andrew Buggee, President, Northeastern AIAA chapter Dr. Andrew Goldstone, Faculty Advisor John Hume, Safety Officer Rob DeHate, Team Mentor Team Roster

More information

University Student Launch Initiative

University Student Launch Initiative University Student Launch Initiative HARDING UNIVERSITY Critical Design Review February 4, 2008 The Team Dr. Edmond Wilson Brett Keller Team Official Project Leader, Safety Officer Professor of Chemistry

More information

LEVEL 3 BUILD YELLOW BIRD. Dan Schwartz

LEVEL 3 BUILD YELLOW BIRD. Dan Schwartz LEVEL 3 BUILD YELLOW BIRD Dan Schwartz This entire rocket is built using the same techniques I use for my nose cones, a central airframe tube for compression strength and rings of high compression styrofoam

More information

First Nations Launch Rocket Competition 2016

First Nations Launch Rocket Competition 2016 First Nations Launch Rocket Competition 2016 Competition Date April 21-22, 2016 Carthage College Kenosha, WI April 23, 2016 Richard Bong Recreational Park Kansasville, WI Meet the Team Wisconsin Space

More information

Post Launch Assessment Review

Post Launch Assessment Review AIAA Orange County Section Student Launch Initiative 2011-2012 Post Launch Assessment Review Rocket Deployment of a Bendable Wing Micro-UAV for Data Collection Submitted by: AIAA Orange County Section

More information

Preliminary Design Review November 15, Agenda. California State Polytechnic University, Pomona W. Temple Ave, Pomona, CA 91768

Preliminary Design Review November 15, Agenda. California State Polytechnic University, Pomona W. Temple Ave, Pomona, CA 91768 Preliminary Design Review November 15, 2017 Agenda California State Polytechnic University, Pomona 3801 W. Temple Ave, Pomona, CA 91768 Agenda 1.0 General Information 2.0 Launch Vehicle System Overview

More information

Student Launch. Enclosed: Proposal. Submitted by: Rocket Team Project Lead: David Eilken. Submission Date: September 30, 2016

Student Launch. Enclosed: Proposal. Submitted by: Rocket Team Project Lead: David Eilken. Submission Date: September 30, 2016 University of Evansville Student Launch Enclosed: Proposal Submitted by: 2016 2017 Rocket Team Project Lead: David Eilken Submission Date: September 30, 2016 Payload: Fragile Material Protection Submitted

More information

Madison West High School Green Team

Madison West High School Green Team Madison West High School Green Team The Effect of Gravitational Forces on Arabidopsis Thaliana Development Flight Readiness Review The Vehicle Mission Performance Criteria Successful two stage flight Altitude

More information

Strap-on Booster Pods

Strap-on Booster Pods Strap-on Booster Pods Strap-On Booster Parts List Kit #17052 P/N Description Qty 10105 AT-24/12 Slotted (Laser Cut) Tube 2 10068 Engine Mount (AT-18/2.75) Tube 2 13029 CR 13/18 2 13031 CR 18/24 4 14352

More information

USLI Flight Readiness Review

USLI Flight Readiness Review UNIVERSITY OF MINNESOTA TWIN CITIES 2011 2012 USLI Flight Readiness Review University Of Minnesota Team Artemis 3/26/2012 Flight Readiness Report prepared by University of Minnesota Team Artemis for 2011-2012

More information

Innovating the future of disaster relief

Innovating the future of disaster relief Innovating the future of disaster relief American Helicopter Society International 33rd Annual Student Design Competition Graduate Student Team Submission VEHICLE OVERVIEW FOUR VIEW DRAWING INTERNAL COMPONENTS

More information

Critical Design Review

Critical Design Review AIAA Orange County Section Student Launch Initiative 2011-2012 Critical Design Review Rocket Deployment of a Bendable Wing Micro-UAV for Data Collection Submitted by: AIAA Orange County Section NASA Student

More information

Florida A & M University. Flight Readiness Review. 11/19/2010 Preliminary Design Review

Florida A & M University. Flight Readiness Review. 11/19/2010 Preliminary Design Review Florida A & M University Flight Readiness Review 11/19/2010 Preliminary Design Review 1 Overview Team Summary ~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~ ~~~~~~~~ Vehicle Criteria ~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~ ~~~~~~~~

More information

Deployment and Flight Test of Inflatable Membrane Aeroshell using Large Scientific Balloon

Deployment and Flight Test of Inflatable Membrane Aeroshell using Large Scientific Balloon 1 Deployment and Flight Test of Inflatable Membrane Aeroshell using Large Scientific Balloon Kazuhiko Yamada, Takashi Abe (JAXA/ISAS) Kojiro Suzuki, Naohiko Honma, Yasunori Nagata, Masashi Koyama (The

More information

Table of Content 1) General Information ) Summary of PDR Report ) Changes Made Since Proposal ) Safety... 8

Table of Content 1) General Information ) Summary of PDR Report ) Changes Made Since Proposal ) Safety... 8 Table of Content 1) General Information... 3 1.1 Student Leader... 3 1.2 Safety Officer... 3 1.3 Team Structure... 3 1.4 NAR/TRA Sections... 4 2) Summary of PDR Report... 5 2.1 Team Summary... 5 2.2 Launch

More information

Northwest Indian College Space Center USLI Post Launch Assessment Review

Northwest Indian College Space Center USLI Post Launch Assessment Review Northwest Indian College Space Center USLI Post Launch Assessment Review 2012-2013 Table of Contents I. Team Summary... 1 Team Name: Northwest Indian College RPGs... 1 II. Launch Vehicle Summary... 1

More information

AKRONAUTS. P o s t - L a u n c h A ss e s m e n t R e v i e w. The University of Akron College of Engineering. Akron, OH 44325

AKRONAUTS. P o s t - L a u n c h A ss e s m e n t R e v i e w. The University of Akron College of Engineering. Akron, OH 44325 AKRONAUTS Rocket Design Team Project P o s t - L a u n c h A ss e s m e n t R e v i e w The University of Akron College of Engineering 302 E Buchtel Ave Akron, OH 44325 NASA Student Launch Initiative April

More information

THE UNIVERSITY OF AKRON

THE UNIVERSITY OF AKRON THE UNIVERSITY OF AKRON College of Engineering 302 E Buchtel Ave Akron, OH 44325 September 20, 2017 NASA Student Launch Initiative Table of Contents 1. Adult Educators and Advisors... 4 2. Team Officials...

More information

University Student Launch Initiative Preliminary Design Review

University Student Launch Initiative Preliminary Design Review UNIVERSITY OF MINNESOTA TWIN CITIES 2012 2013 University Student Launch Initiative Preliminary Design Review Department of Aerospace Engineering and Mechanics 3/18/2013 2012-2013 University of Minnesota

More information

Rocket Design. Tripoli Minnesota Gary Stroick. February 2010

Rocket Design. Tripoli Minnesota Gary Stroick. February 2010 Rocket Design Tripoli Minnesota Gary Stroick February 2010 Purpose Focus is on designing aerodynamically stable rockets not drag optimization nor construction techniques! Copyright 2010 by Gary Stroick

More information

Reentry Demonstration Plan of Flare-type Membrane Aeroshell for Atmospheric Entry Vehicle using a Sounding Rocket

Reentry Demonstration Plan of Flare-type Membrane Aeroshell for Atmospheric Entry Vehicle using a Sounding Rocket AIAA ADS Conference 2011 in Dublin 1 Reentry Demonstration Plan of Flare-type Membrane Aeroshell for Atmospheric Entry Vehicle using a Sounding Rocket Kazuhiko Yamada, Takashi Abe (JAXA/ISAS) Kojiro Suzuki

More information

Flight Readiness Review Report NASA Student Launch Florida International University American Society of Mechanical Engineers (FIU-ASME)

Flight Readiness Review Report NASA Student Launch Florida International University American Society of Mechanical Engineers (FIU-ASME) Flight Readiness Review Report 2014-2015 NASA Student Launch Florida International University American Society of Mechanical Engineers (FIU-ASME) Florida International University Engineering Center College

More information

NASA University Student Launch Initiative (Sensor Payload) Final Design Review. Payload Name: G.A.M.B.L.S.

NASA University Student Launch Initiative (Sensor Payload) Final Design Review. Payload Name: G.A.M.B.L.S. NASA University Student Launch Initiative (Sensor Payload) Final Design Review Payload Name: G.A.M.B.L.S. CPE496-01 Computer Engineering Design II Electrical and Computer Engineering The University of

More information

University of North Dakota Department of Physics Frozen Fury Rocketry Team

University of North Dakota Department of Physics Frozen Fury Rocketry Team University of North Dakota Department of Physics Frozen Fury Rocketry Team NASA Student Launch Initiative Flight Readiness Review - Report Submitted by: The University of North Dakota Frozen Fury Rocketry

More information

Introduction: Problem statement

Introduction: Problem statement Introduction: Problem statement The goal of this project is to develop a catapult system that can be used to throw a squash ball the farthest distance and to be able to have some degree of accuracy with

More information

First Revision No. 9-NFPA [ Chapter 2 ]

First Revision No. 9-NFPA [ Chapter 2 ] 1 of 14 12/30/2015 11:56 AM First Revision No. 9-NFPA 1127-2015 [ Chapter 2 ] Chapter 2 Referenced Publications 2.1 General. The documents or portions thereof listed in this chapter are referenced within

More information

Pre-Flight Checklist for SLIPSTICK III

Pre-Flight Checklist for SLIPSTICK III Advanced Planning 1 Schedule a Check that waivers are available at the intended launch site and date. b Check weather forecast for wind and temperature conditions at the site. c Have TAP members approved

More information

Electronic Deployment

Electronic Deployment Electronic Deployment and a little bit of recovery too! By: Gerald Meux, Jr. NAR and TRA Level 3 1-3-11 8/28/2014 Electronic Deployment - Gerald Meux, Jr. 1 Table of Contents 8/28/2014 Electronic Deployment

More information

Critical Design Review

Critical Design Review Critical Design Review 1/27/2017 NASA Student Launch Competition 2016-2017 California State Polytechnic University, Pomona 3801 W Temple Ave, Pomona, CA 91768 1/27/2017 California State Polytechnic University,

More information

Remote Control Helicopter. Engineering Analysis Document

Remote Control Helicopter. Engineering Analysis Document Remote Control Helicopter By Abdul Aldulaimi, Travis Cole, David Cosio, Matt Finch, Jacob Ruechel, Randy Van Dusen Team 04 Engineering Analysis Document Submitted towards partial fulfillment of the requirements

More information

Rocket Activity Advanced High- Power Paper Rockets

Rocket Activity Advanced High- Power Paper Rockets Rocket Activity Advanced High- Power Paper Rockets Objective Design and construct advanced high-power paper rockets for specific flight missions. National Science Content Standards Unifying Concepts and

More information

Deployment and Drop Test for Inflatable Aeroshell for Atmospheric Entry Capsule with using Large Scientific Balloon

Deployment and Drop Test for Inflatable Aeroshell for Atmospheric Entry Capsule with using Large Scientific Balloon , Germany Deployment and Drop Test for Inflatable Aeroshell for Atmospheric Entry Capsule with using Large Scientific Balloon Kazuhiko Yamada, Takashi Abe (JAXA/ISAS) Kojiro Suzuki, Naohiko Honma, Yasunori

More information

Modified shock-cord mount and cables (cables are shown pushed into motor mount here)

Modified shock-cord mount and cables (cables are shown pushed into motor mount here) Building the Ariel Builder: Ray Wilkinson This is Ray Wilkinson's own rocket, but will mostly reside at UH, and will be used for display purposes as well as being flown. It's built from a kit made by PML

More information

Name: Space Exploration PBL

Name: Space Exploration PBL Name: Space Exploration PBL Students describe the history and future of space exploration, including the types of equipment and transportation needed for space travel. Students design a lunar buggy and

More information

XIV.C. Flight Principles Engine Inoperative

XIV.C. Flight Principles Engine Inoperative XIV.C. Flight Principles Engine Inoperative References: FAA-H-8083-3; POH/AFM Objectives The student should develop knowledge of the elements related to single engine operation. Key Elements Elements Schedule

More information

Project WALL-Eagle Maxi-Mav Critical Design Review

Project WALL-Eagle Maxi-Mav Critical Design Review S A M U E L G I N N C O L L E G E O F E N G I N E E R I N G Auburn University Project WALL-Eagle Maxi-Mav Critical Design Review 2 Engineering Dr. Auburn, AL 36849 January 6th, 205 Table of Contents SECTION

More information

Pre-Launch Procedures

Pre-Launch Procedures Pre-Launch Procedures Integration and test phase This phase of operations takes place about 3 months before launch, at the TsSKB-Progress factory in Samara, where Foton and its launch vehicle are built.

More information

NWIC Space Center s 2017 First Nations Launch Achievements

NWIC Space Center s 2017 First Nations Launch Achievements NWIC Space Center s 2017 First Nations Launch Achievements On April 18, 2017, we were on two airplanes to Milwaukee, Wisconsin by 6:30 am for a long flight. There were 12 students, 3 mentors, 2 toddlers

More information

To determine which number of fins will enable the Viking Model Rocket to reach the highest altitude with the largest thrust (or fastest speed.

To determine which number of fins will enable the Viking Model Rocket to reach the highest altitude with the largest thrust (or fastest speed. To determine which number of fins will enable the Viking Model Rocket to reach the highest altitude with the largest thrust (or fastest speed.) You are a mechanical engineer that has been working on a

More information

Project WALL-Eagle Maxi-Mav Flight Readiness Review

Project WALL-Eagle Maxi-Mav Flight Readiness Review S A M U E L G I N N C O L L E G E O F E N G I N E E R I N G Auburn University Project WALL-Eagle Maxi-Mav Flight Readiness Review 2 Engineering Dr. Auburn, AL 36849 March 6th, 205 Table of Contents Section

More information

Preliminary Design Review

Preliminary Design Review Preliminary Design Review November 16, 2016 11/2016 California State Polytechnic University, Pomona 3801 W Temple Ave, Pomona, CA 91768 Student Launch Competition 2016-2017 1 Agenda 1.0 General Information

More information