S AMUEL G INN C OLLEGE OF E NGINEERING. Auburn University. University Student Launch Initiative Critical Design Review

Transcription

1 S AMUEL G INN C OLLEGE OF E NGINEERING Auburn University University Student Launch Initiative Critical Design Review

2 Section 1: Summary of PDR Report Team Summary Launch Vehicle Summary Payload Summary... 1 Section 2: Changes Made since PDR Vehicle Design Changes Payload Criteria Project Plan... 4 Section 3: Vehicle Criteria Mission Statement Mission Requirements Mission Success Criteria Design Review and Analysis of Subsystems... 7 Nose Cone... 7 Recovery Section... 8 Booster Section Fins Motor Selection and Mass Estimation System Level Functional Requirements and Verification Workmanship Planned Testing Integrity of Design Suitability of Fin Shape Proper Use of Materials Proper Assembly Procedures... 26

3 Sufficient Motor Mounting and Retention Status of Verification Propulsion Verification a Launch System Verification b Motor Mounting System Verification System Stability Verification Mass Adjustments Verification Fin Mounting Verification Structural Verification Fuselage Verification Tube Coupling Verification Recovery System Verificatio Retention Verification Design Drawings Mass Estimations Mission Performance Criteria Flight Simulations Drift Simulations Flight Predictions Drag Calculations Stability Calculations Recovery preparation Motor preparation Igniter installation Setup on launcher Postflight inspection Safety Officer Preliminary Failure Mode Hazards Environmental Concerns Section 4: Payload Criteria Payload Overviews Threat Analysis Payload (TAP)... 55

4 4.1.2 Boost Phase Analysis Payload (BPAP) Corrosion Mitigation Payload (CMP) System Level Review Payload Requirements and Objectives Advantages and Disadvantages of Each Payload Selection Payload Subsystems Payload Design and Integration Boost Phase Analysis Payload Corrosion Mitigation Payload Verification Plan and Performance Metrics Verification Plan Performance Expectations and Metrics Payload Concept Features and Definition Creativity, Originality, and Uniqueness Level of Challenge Science Value Payload Objectives Payload Success Criteria Experimental Logic, Approach, and Method of Investigation Test and Measurement, Variables, and Controls Relevance of Expected Data and Accuracy/Error Analysis Preliminary Experiment Process Procedures Section 5: Project Plan Budget Plan Funding Plan Educational Engagement Section 6: Conclusion Appendix A: Educational Engagement School Handbook Appendix B: Sampling of MSDS

5 Appendix C: Collection of Risk Mitigation Tables

6 1.1 Team Summary Section 1: Summary of PDR Report School: Team Name: Location: Auburn University Project Nova Davis Hall 141Engineering Drive, Auburn, AL Mentor: Christopher Short TRA #10247 L3 TAP NAR #8300 L3CC 1.2 Launch Vehicle Summary The most important launch vehicle characteristics are given in the following table: Table 1.1: Launch Vehicle Summary Total Length (in) Diameter (in) Gross Lift Off Weight (lb) Airframe Material Rail Size Motor Manufacturer(s) Motor Designation(s) Max/Average Thrust (lb) Total Impulse (lbf sec) 108 in 5 in 67.9 lb Fiberglass 15/15, Large CTI N /509.8 lbf lb s 1.3 Payload Summary Threat Analysis Payload will serve as a threat detection system that will identify possible threats and hazards in a landing zone from two externally mounted cameras on the aft section of the rocket. An onboard computer system will analyze the images for possible threats using an integrated software package designed and programmed by the team. 1

7 The BPAP payload will provide information about the rocket during the boost phase through a series of different sensors providing data about critical subsystems of the rocket. The Corrosion Mitigation Payload (CMP) is designed to test the effectiveness of the of the NeverWet anti icing system at supersonic speeds, this was accomplished using a heat sensitive paint underneath the layers of both paint and NeverWet. 2

8 Section 2: Changes Made since PDR 2.1 Vehicle Design Changes A Second Motor was selected in case of the altitude switch from 20,000 feet to 10,000 feet. The fin design was switched to a three fin design for the reasons listed in the report. Various dimensions of the vehicle have also changed. Throughout the preliminary design, modifications were made to the original length of the vehicle s subsections. The first major change was the decrease in size of the payload section to eliminate empty space. By decreasing the length of the payload section, extra space was added within the booster section to accommodate the second major change: the addition of the static stability chamber. This was done on a further level for the CDR, and the mass numbers finally tuned. Finally, the most significant change in the vehicle design was the switch from CO2 canisters to the more traditional black powder. This was due to lack of production of commercially available CO2 canisters from Rouse Tech, since they are no longer producing additional canisters, and the team was doubtful that enough reloads could be purchased to fully test the system. 2.2 Payload Criteria The only significant change within the payload critieria was a further development of the risks and the downselection of nosecone materials if required. All other changes within the payload criteria only represent a further sophistication of the design moving forward towards the CDR and ultimately the flight of Project Nova. 3

9 2.3 Project Plan The project plan saw little difference other than a concretization of the travel and educational plans. In addition, the Boy Scouts activity was also added to the educational engagement plans. 4

10 Section 3: Vehicle Criteria 3.1 Design and Verification of Launch Vehicle Mission Statement The mission of the Auburn University Rocketry Association s launch vehicle, Project Nova, is to deliver a primary payload to a target altitude of 15,500 feet while achieving supersonic flight speed during ascent for data acquisition and testing of our Corrosion Mitigation Payload (CMP) and Boost Phase Analysis Payload (BPAP). During descent, the Threat Analysis Payload (TAP), consisting of two cameras attached symmetrically to the sides of the launch vehicle, will scan the ground in search of hazardous objects and report this to the ground station for assessment Mission Requirements In order for the mission to be successful, it must be carried out in accordance to all NASA University Student Launch Initiative competition rules and it must also adhere to all NAR safety regulations. In addition to the USLI competition rules and regulations, the following requirements must also be met: 1. The launch vehicle must provide adequate structure to safely and reliably deliver all payloads to the required altitude and back to the ground. 2. The launch vehicle fuselage must stay fully intact from launch to landing. This means no structural failures on the outside of the vehicle, such as cracks in the fuselage or nose cone, or broken fins; and no structural failures internally, such as broken bulkheads and motor mounts, stripped pins, and damaged payload bays. 5

11 3. All onboard electronic systems are fully operational at the end of the mission, with all data being recoverable as well. 4. The launch vehicle must be able to be assembled on the launch field in less than two hours. 5. The launch vehicle must be flight ready within one hour of a prior launch without any repairs or modifications. 6. The launch vehicle must be able to sit on the launch pad for a minimum of one hour, maintaining flight readiness of all systems during this time, including the launch system itself, BPAP, TAP, and all onboard avionics. 7. The launch vehicle must achieve a minimum flight speed of Mach An apogee of 15,500 feet must be reached, allowing for a margin of two hundred and fifty feet above or one hundred feet below Mission Success Criteria Mission success hinges on completion of all USLI competition objectives, as well as the following: 1. The launch vehicle reaches its target altitude of 15,500 feet within the margins listed in the Mission Requirements. 2. Temperature fluctuations on all thermocouple locations are recorded from launch to end of burn time and all data recorded is marginally close to calculated values. 3. Pressure fluctuations at all sensor locations are recorded from launch to apogee and all data recorded is marginally close to calculated values. 4. Strain forces are recorded at each location of a strain gauge within the fuselage of the launch vehicle, from launch to landing, and all data recorded is marginally close to calculated values. 5. All data recorded during the mission is recoverable. 6. The onboard recovery system deploys correctly and decreases the descent rate of the vehicle to its designed criteria. 6

12 7. The primary payload must be able to identify any large objects on the ground and report them as hazardous to the ground station. 8. All sections of the rocket must land on the ground while generating a kinetic energy less than the designated 75 ft lbf Design Review and Analysis of Subsystems The proposed launch vehicle is shown on the following page in Figure 3.1, with vehicle sections indicated. Color shown on the launch vehicle in Figure 3.1 are for visual purposes only, and are not indicative of actual vehicle color. As shown, the vehicle has three primary independent sections: nose cone, recovery section, and booster section. Each section separates completely during recovery system deployment and is connected together via shock cords. Each section contains the required internal components to achieve mission success including payloads, propulsion systems, avionics, and recovery systems. Nose cone Recovery Section Booster Section Figure 3.1: Overview of the vehicle design, indicating the three primary sections, along with the CG and CP locations represented by the blue and red dots, respectively. Nose Cone The nose cone of the launch vehicle, as indicated in the figure below, is a total of eighteen inches long with a diameter of five inches at the base. The shape of the nose cone is ogive, which creates a pointed tip and curved surface, ideal for supersonic speeds to minimize the creation of shockwaves at sharp corners. It tapers at the end to create a seamless transition from the nose cone to the recovery section for lowered drag and optimal aerodynamics. 7

13 Figure 3.2: Overview of the nose cone, with annotations providing dimensions. The nose cone is to be made of fiberglass with a thickness of inches. It will be created utilizing a similar process as outlined in the recovery and booster sections, with the only differences being the geometry of the mold. Alternatively, if the team cannot create a satisfactory part in the allotted time, given the challenging nature of the geometry of the mold, the team will down select to a commercial 5 plastic nosecone. The nose cone will be jettisoned from the rocket upon main parachute deployment. Pressure tests will be performed to ensure the blackpowder deployment system has the power required to break the seal between the nose cone and the payload section Recovery Section The length of the recovery section is a total of 21 inches and is constructed completely out of cardboard. It houses the main parachute, including the shroud lines and shock cords. It also includes the forward part of the electronics bay #1, which also acts as the tube coupler between the recovery section and booster section. The electronics bay #1 includes avionics equipment, as well as BPAP equipment including the data logger, pressure transducers, and 8

14 strain gauges, which will all be located within the electronics bay #1, as shown in the figure below. Main Parachute Bay Forward part of electronics bay #1 Figure 3.3: Dimensional overview of the recovery section. The main parachute has twelve should lines connecting the parachute to a shock cord located within the casing, just forward of the electronics bay #1. The main parachute is made of ripstop nylon and is one hundred and ninety three inches in diameter. It will deploy at an altitude of two thousand feet and has been sized to decrease the velocity of the vehicle to approximately fifteen feet per second. The shock cord of the main will be attached to the forward most bulkhead. To insure the rigidity of this bulkhead and to verify that the bulkhead will not dislodge from within the section, force tests will be performed to determine the maximum tensile stress the bulkheads are capable of withstanding, as outlined in the verification portion. Next is the electronics bay #1, which houses the avionics system including redundant barometric altimeter and GPS tracking device as well as part of the BPAP, specifically the pressure and strain gauge data logger and instrumentation associated with each. The pressure transducers will be situated in two locations along the outer surface of the recovery 9

15 section and the strain gauges will be located in varying locations along the inner surface of the vehicle. All wiring will run through either the forward or aft end of the electronics bay #1, through an orifice on the bulkheads, and to the instrumentation. The bay will be a cardboard tube with an inner diameter of 4.8 inches and an outer diameter of 4.9 inches. It will be twelve inches long and situated between two bulkheads, both made of fiberglass with a thickness of one quarter inch, and each acting as the anchor for the main and drogue parachute shock cords, respectively. Figure 3.4: Overview of recovery section showing distribution of equipment including parachutes, electronics bay #1, and bulkheads. 10

16 Booster Section The booster section of the vehicle consists of the aft part of the electronics payload bay #1, the electronics payload bay #2, and the propulsion system. The total length of the booster section is sixty nine inches, making it the largest of the three sections. It maintains a diameter of five inches. It is also made out of fiberglass. Aft portion of EB #1 Electronics bay #2 Motor Figure 3.5: Overview of the booster section showing the electronics bay #2, motor, and fin placement. From forward to aft, the first component is the aft portion of electronics bay #1. The second is electronics bay #2, which is housed in a fiberglass tube with a diameter of 3.5 inches. It is 9.4 inches long. The tube contains the electrical components for Threat Analysis Payload and the Booster Phase Analysis Payload. The drogue parachute, along with shock cord and shroud lines, is situated at the fore end of the motor section. The drogue is attached to a shock cord, which connects the recovery section to the booster section. The drogue parachute is 31.2 inches in diameter and has been sized to limit the descent speed of the vehicle to a maximum of feet per second. Just as with the main parachute, strength tests will be performed on the bulkheads holding the shock cord of the drogue to insure that the bulkhead will withstand the force. As according to the verification matrix. 11

17 The payload section and booster section will be held together using a tube coupler that is 4.75 inches in diameter and six inches long. It will be made from cardboard. As with the nose cone, a pressure test will be performed to insure that the blackpowder deployment system is powerful enough to separate the recovery section from the booster section. The Threat Analysis Payload (TAP) includes a camera system that scans the surface during descent in order to detect potential landing hazards. The camera system will consist of two Raspberry Pi s with camera attachments for both. The cameras will be houses in enclosing s symmetrically placed on the sides of the launch vehicle, maintaining a camera orientation towards the ground for optimal visualization of the ground surface below. Along with the Raspberry Pi s will be the electronics necessary to transmit the data gathered to be analyzed to a ground station in real time by a custom designed, onboard software package that will determine if landing hazards are present in the landing zone. The third component of the BPAP system is located in the electronics bay #2. It will test consist of thermocouples placed on the nozzle of the motor, as well as on the casing of the motor, and will record the temperature fluctuation of both during boosting phase of the flight. This data will be stored in a data logger located in electronics bay #2. This data logger also has the capability of recording ambient temperature, and will serve a secondary function recording any temperature fluctuations happening within the electronics bay #2. In the case that the optimal weight of the rocket is not reached at final design, a special feature has Figure 3.6: 2D view of the ballast tank showing proximity to the CG 12

18 been incorporated within the rocket behind the electronics bay #2, namely, a ballast chamber. This chamber will hold sand. Sand was chosen due to its ease of accessibility and ease of transport. It can mold itself into any shape allowing for maximum adaptability to the size and configuration of the ballast chamber. Also, sand is very stable, in the sense that its very abrasive, which causes it to move very little when not acted upon by some force. Being placed inside the chamber with no other forces than those generated by the motor will insure minimal movement within the ballast chamber. This will keep the stability variableness to a small margin. Also, the density of sand is high enough that it can increase the weight significantly without requiring large amounts of volume. The far aft of the booster section will house the propulsion system, a Cesaroni Technologies N2200 PK engine. The N2200 PK will be mounted flush to the aft of the vehicle. The engine is 3.86 inches in diameter and 39.8 inches in length. Also, an additional motor mount will be placed at the center of the motor casing, and a third one will be at the far aft end. The dissipation of the shear forces into the body tube through the bulkplate, coupled with the restrictive forces generated in the motor mounts will insure the motor does not break away from its holds and explode through the rocket upon ignition or during flight. 13

19 Fins A trapezoidal fin shape was chosen for the launch vehicle. A total of three fins are to be used, each situated symmetrically around the external surface of the launch vehicle, as shown in the figure to the right. The surface area size of all four fins equates to a total surface area of square inches and given that the fins are located three inches from the base of the rocket, this gives a stability caliber of 2.55 Figure 3.7: Back view of the rocket showing with the center of gravity position being symmetrical orientation of fins inches from the nose of the rocket and the center of pressure being inches from the nose. This ensures complete stability of the system during ascent to the desired target altitude of 15,500 feet. 14

20 Figure 3.8: Dimensional view of the fins The table below provides the overall dimensions of the fins. Note: the sweep angle is measured from the vertical, and LE is dictation for leading edge. Table 3.1: Fin Dimensions Root chord 8.40 in Tip chord 4 in Height 5.50 in Sweep Length 6 in Sweep angle

21 Motor Selection and Mass Estimation Motor selection was accomplished using the criteria needed for mission success, specifically the motor had to meet the following requirements: 1. The motor had to have enough power to accelerate the rocket up to supersonic speeds without going to far outside of the transonic region, ideally around Mach The motor could not deliver the rocket past the designated altitude limit of 20,000 feet, as set by the USLI competition rules. 3. The motor had to deliver the payload to teams target altitude of 15,500 feet given any mass increases on the order of ~20%. Given the possibility of the total rocket mass increasing 25 33% from preliminary design to final design, as stated in the USLI competition information packet, the motor had to be capable of still achieving the above requirements regardless of a mass increase. After running simulations, it was determined that the Cesaroni N2200 motor had the highest probability of meeting all requirements should the mass increase significantly. The data from the simulations using the 25 33% mass increase margin is given in Table 3.2 below. Table 3.2: Simulated Apogees with Mass Increases 0% Mass Increase 25% Mass Increase 33% Mass Increase Total Liftoff (lbm) Apogee Achieved (ft) Highest Mach # The data above shows that if the mass of the rocket increases above 25%, the target apogee will not be achieved and supersonic flight will likely also not be obtained. Further simulations were done to find the optimal weight of the rocket that would achieve our designated apogee, while still reaching Mach numbers above one. These simulations showed that the optimal weight is 68.5 lbm. This allows for a mass increase of 22% from our current 16

22 weight, providing a good margin of mass increase to work with from current design point to final design. The Cesaroni N2200 motor is available for purchase from What s Up Hobbies, located in Stockton, CA, and from Wildman Rocketry, located in Van Orin, IL. The chosen location for purchase will be based on proximity to competition site, capability to deliver motor to competition, and cost of shipping motor from shop location to Auburn University. The thrust curve for the N2200 motor is shown in the figure below. Below that are tables detailing the performance characteristics of the N2200 motor. Table 3.3: Cesaroni N2200 Data Total Impulse lb s Maximum Thrust lbf Average Thrust lbf Specific Impulse s Burntime 5.55 s 17

23 Figure 3.9: Thrust curve for Cesaroni N2200 motor. 18

24 If the launch site is changed resulting in a ceiling limit of 10,000 feet, the alternate motor will be a Cesaroni M1830. The estimated altitude with this configuration is 9174 feet. The CMP will be unachievable, however, due to lack of distance and minimized burntime of motor. The highest calculated Mach is The data for the Cesaroni M1830 is given below. Table 3.4: Cesaroni M1830 Data Total Impulse lb s Maximum Thrust lbf Average Thrust lbf Specific Impulse s Burntime 3.07 s Figure 3.10: Thrust curve for Cesaroni M1830 motor. 19

25 3.1.5 System Level Functional Requirements and Verification The table below lists all system level functional requirements necessary for mission success, along with the verification plan to insure the requirement is met and the status of the implementation of the plan. Table 3.4: System Level Requirements, Verification, and Status Requirement Verification Plan Status Shall deliver the research payload to a predetermined altitude appropriate for the associated payload. The Corrosion Mitigation Payload requires supersonic speeds so a class N motor was chosen for the design while keeping the Simulations have proven the design capable of meeting the requirement. The next phase will be fullscale test. apogee below 20,000 feet. Shall carry one commercially available, Altus Metrum Telemetrum Purchased and tested on the subscale. barometric altimeter for recording the official altitude used in the competition scoring. Shall be designed to be recoverable and reusable. The launch vehicle will be designed so that the recovery system will decrease the descent velocity of the vehicle so that upon ground impact, the total kinetic energy will be less than the designated 75 ft lbf. Simulations have proven the design capable of meeting the requirement. The next phase will be subscale and full scale test flights to determine if max kinetic energy at landing is less than 75 ft lbf. 20

26 Shall be capable of being prepared for flight at the launch site within 2 hours. Shall 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 component. Shall be capable of being launched by a standard 12 volt direct current firing system. Shall require no external circuitry or special groundsupport equipment to initiate launch. Shall use a commercially available solid motor propulsion system approved by NAR, TRA, and CAR. The launch vehicle will be design so that total vehicle assembly will take less than two hours. The launch vehicle is designed using a [INSERT BATTERY INFORMATION HERE] that will insure the rocket will remain launch ready for a minimum of one hour. All data loggers onboard also have integrated battery level indicators for ease of checking before vehicle is placed on launch pad. Standard igniters purchased from Chris Rocket Supplies will be used to ensure proper ignition on the pad. The launch vehicle is designed to have no external circuitry or ground support equipment for launch initiation. The Cesaroni N2200PK P, which the team has chosen for the launch vehicle, has been certified by NAR, TRA, and CAR, as referenced on Practice assembly runs will be done after construction of the subscale and the fullscale rocket are complete. All batteries will be charged prior to any launches, checked before installation into the vehicle, and check once more before vehicle is placed on the launch pad. Complete Complete Complete 21

27 Shall successfully launch and recover full scale rocket prior to FRR. there latest list of certified motors, accessible from NAR.com A test flight with the fullscale rocket will be performed prior to FRR. Incomplete Workmanship The workmanship that each team member puts forth into the project has to be of upmost importance. To insure this is done, the team works in open groups, allowing correction by fellow team members to make sure that any and all results are accomplished successfully. This applies to all aspects of the competition build including the overall design of the launch vehicle, the manufacturing of component parts, the assembling of subsystems and main systems, and testing of systems, as well as flight tests with subscale and full scale versions of the launch vehicle. The team seeks guidance from mentors to insure that all aspects of the build are being performed correctly. Having the ability to get feedback from professors, workshop technicians, and the team mentor insures that the completed, full scale launch vehicle will be of sound construction and integrity, and will deliver mission success Planned Testing Before the full scale launch vehicle will be fully assembled, a structural buckling test will be performed to insure that the fiberglass fuselage that will make up the body of the rocket can withstand the calculated normal forces that it can be expected to experience. Using OpenRocket to simulate flight data, estimated forces were calculated for the rocket during launch. The forces were calculated by summing the total forces exerted on the launch vehicle in the vertical direction, essentially thrust minus drag and gravitational forces. The entire 22

28 rocket can be expected to experience a max normal force of lbf right after launch, as depicted in the graph below. Figure 3.11: Graph of Thrust vs. Total Force vs. Time showing the point of maximum normal force A three foot long fiberglass fuselage, with a five inch inner diameter, will be placed on loading cells and loaded to the point of buckling failure. If the load at failure is at least one and 20% greater than the max normal force calculated, the structure will be deemed suitable for manufacturing commencement. 23

29 3.1.8 Status and Plans of Remaining Manufacturing and Assembly 24

30 25

31 3.1.9 Integrity of Design Suitability of Fin Shape The mission requires that the launch vehicle reach supersonic speeds. This means the fins have to be very aerodynamically efficient, creating the least amount of drag possible while still maintaining a stability caliber greater than 2.2 so that the vehicle ascends in at least a nearly vertical orientation. The current fin design meets all of the above criteria. The overall stability currently is Simulations show that the launch vehicle reaches a Mach number of 1.25, and that, even after increases in mass of up to 22%, the vehicle can be expected to reach local supersonic conditions along the nose cone, which is the payload that requires supersonic conditions be present Proper Use of Materials Within every lab the MSDS sheets for all materials are posted on the doors as well as around the workbenches and other areas where work is done in order to ensure that all materials are handled safely Proper Assembly Procedures Chris Short, the team s level 2 mentor will oversee the construction of the rocket. The team will verify that all materials are assembled according to the workmanship required to ensure that the rocket is soundly built. 26

32 Sufficient Motor Mounting and Retention The team has decided to utilize the 98 mm flanged aeropack container, shown right, mounted in a ½ bulkplate at the bottom of the rocket in order to retain the motor. The thickness of the bulkplate provides sufficient surface area to ensure that the force of the motor is transferred to the body tube of the wall, thus transferring the load into the structurally sound body tube Status of Verification The verification of the various systems has begun, or will begin, as outlined in the verification sections of each portion of the system review Propulsion Verification a Launch System Verification The launch system will be verified based on geometric analysis of the launch vehicle with respect to the launch pad. The launch vehicle must be perpendicular to the launch pad. Angular measurements will be made at three different spots on the launch vehicle: the top of the recovery section, the separation point of the recovery and booster section, and the base of the rocket. The launch pad will also be checked for parallelization to the horizon. This will be accomplished using a leveler. If all three points lie perpendicular to the launch pad and the launch pad is level, flight will be verified to be perpendicular from the point of lift off from the ground b Motor Mounting System Verification The motor retaining system must transmit the force due to thrust from the motor into the airframe as a normal force, not shear force. This system will be tested twice, once during a 27

33 subscale flight using a lower grade motor, and the second time during a full scale flight using the motor indicated in section Motor Selection and Mass Estimation System Stability Verification Mass Adjustments Verification The mass of the rocket will be the largest determining factor of its stability. A ballast tank has been implemented within the design. The ballast tank will be used to increase the weight, if necessary, to maintain a level of stability greater than 2.20 caliber. Two flight tests will also be done to insure the stability of the rocket matches the design calculations. Fin Mounting Verification The fins will be mounted through the rocket and down to the motor casing. The fins will need to be rigid to resist fluttering. The fins will also need to produce the calculated and simulated stability effects expected. Two flight tests will be done to verify the overall stability of the rocket and that the fins maintain structural rigidity during flight. Structural Verification Airframe Fins Verification Description The airframe will be statically loaded to the calculated maximum load, as well as a load approximately 20% above the calculated maximum load in order to ensure the airframe will not fail. The fins will be statically loaded to the calculated maximum load, as well as a load approximately 20% above the calculated maximum load in order to ensure the airframe will not fail. 28

34 Connection Points Launch Lug Ejection Charge Each connection point will be assembled and threaded with the nylon machine screw acting as the shear pin connection in order to verify that the connection points do not fail premature to the black powder charge forcing separation. The launch lug will be statically loaded to the maximum thrust attained on the thrust curve of the motor selected, in order to ensure that the lug is securely fastened and will not detach on the launch rail. The size of the black powder charge will be verified through a fullscale ground test of the components being fully assembled. This is in addition to the previous mock ground tests with the equivalently sized tubes to size the ejection charge. This test will also verify that the bulkplates are thick enough to withstand the force of the ejection charge. Fuselage Verification The fuselage of the rocket (fuselage meaning the recovery and booster section) will experience large compressive loads during ascent due to the weight of the rocket and the drag generated during flight pushing against the thrust generated by the motor. The fuselage of the rocket needs to withstand this force without any buckling/compressive failure. To test this, the recovery and booster sections will be independently tested for ultimate buckling failure using loading cells within the Polymer and Fiber Engineering Research lab. If the sections can withstand loading greater than one and one half times the total calculated force the rocket will experience, the structure of the rocket will be validated. Tube Coupling Verification Tube couplers are vital to the success of the recovery system. If the integrity of the coupling fails, the portions of the rocket will separate prematurely, resulting in premature ejection of the parachute, and most likely severe damage to the airframe of the rocket Recovery System Verificatio 29

35 Retention Verification The retention system is throughout the rocket is vital to the recovery of the rocket. If one of the bulkplate fails, the parachutes will fail to deploy, resulting in a failure to recover the rocket and a hazard of a large rocket falling without assistance from a parachute. Table 3.5: Summary of Vehicle Systems and Verification Metrics System Description Functional Requirements Verification Metrics The method and Launch System equipment necessary to properly orient the rocket on the The rocket must launch completely vertical with reference to the zenith. Geometric calculations, on site angular measurements, leveler launch pad. The motor must not Motor Mounting How the motor is held in place within the rocket be displaced from its mounted position in reference to the Ground testing, subscale and full scale flight tests airframe. The fins must be Fin Mounting The way the fins are mounted to the rocket completely rigid and properly orient the rocket during flight, Sound constructing methods, subscale and full scale flight tests maintaining full stability 30

36 The fuselage must Airframe/Fuselage The structure of the fuselage and its ability to withstand a load withstand at least 1.5x the calculated load that it will experience during Load calculations, lab testing using a loading cell, subscale and full scale flight tests flight The method of Tube Coupling attaching one section to another 31

37 Design Drawings Figure 3.12: 2D view of the internal components of the launch vehicle Figure 3.13: Dimensional overview of the launch vehicle. 32

38 Mass Estimations The mass estimations performed during the motor selection process were accomplished in tandem with OpenRocket, the open source rocket modeling and simulation software. In OpenRocket, all components can be assigned material specifications with designated densities for each material. By inputting the lengths and thicknesses for each component, the software calculates the total mass accurately. Masses for each main component of the rocket are listed below. Mass calculations for electronic systems and payloads are estimations currently, and will change as systems are defined and installed throughout the manufacturing process of the rocket. Given the progress made in the development of the rocket, as well as the results gathered from testing of the subscale, a mass increase past the current calculated mass of 56.1 will be minimal. Due to the inaccuracies OpenRocket simulates at supersonic speeds, the estimated apogee may be much lower than simulated. This will be proven after a full scale test flight. All efforts will be made to minimize the required added mass (sand in this case) into the ballast tank to bring the apogee near 15,500 feet. Table 3.6: Mass Estimations (units: lbm) Nose Cone 1.7 Recovery Section 15.2 Bulkheads 1.0 Electronics Bay #1 5.0 Recovery System 5.0 Booster Section 16.0 Bulkheads 2.0 Motor Mounting System 1.0 Electronics Bay #2 3.9 Total Weight (without motor)

39 Motor 25.0 Total Weight (with motor) Subscale Flight Results The subscale rocket was an 80% scale of the fullscale vehicle. This scale was chosen to simulate the stability conditions predicted through computer simulations to insure the overall structure of the rocket would be suitable for fullscale flight and would be able maintain the predicted stability. The vehicle was constructed using similar methods that will be used in the full scale vehicle construction. The fuselages were constructed from phenol, which were wrapped in two layers of carbon fiber to simulate the same surface texture that will be present on the full scale rocket. The full scale rocket will not have the inner phenol, but instead be complete fiberglass. Figure 3.14: One section of the fuselage entering vacuum bagging 34

40 The nosecone was commercially made and provided in a Level 2 rocket kit that had been purchased for the Auburn University Rocketry Association. The nosecone for the full scale rocket will be made completely from fiberglass, as well. The fins were made from Figure 3.15: Subscale rocket fin dimensions. plywood and then wrapped in carbon fiber. The fins were made with dimensions that are slightly smaller than the full scale version dimensions. These were calculated from scaling down the rocket to 80% of full scale size. These dimensions are listed in the figure to the right. Overall dimensional data for the subscale is given below. Table 3.7: Size and Mass Length 86.4 in Diameter 4 in Dry Weight (no motor) 10.3 lbm Wet Weight 11.7 lbm Table 3.8: Motor Choice Manufacturer Aerotech Designation J270W 35

41 Loaded Weight Total Impulse Average Thrust Burntime 1.42 lbm lbf s lbf 2.6 s Table 3.9: Recovery System Main Parachute Diameter Shroud Line Length Number of Shroud Lines Shock Cord Length Drogue Parachute Diameter Shroud Line Length Number of Shroud Lines Shock Cord Length The internal layout of the rocket is shown below. The forward most component is the main parachute with its shock cord. Next is the avionics payload. After the separation point is the drogue parachute with its shock cord, and then finally the motor tube with motor. The internal configuration was modified from the full scale model for easier assembly. 36

42 Figure 3.16: Internal components overview of the subscale rocket. The flight of the subscale was a partial success. The rocket flew with no faults. The stability was consistent with all calculations and simulations that had been performed prior to the flight. It reached an apogee of 2066 feet with a maximum velocity of 157 ft/s. The tables below compare the simulated data with the actual flight data. Table 3.10: Simulated Flight Data versus Actual Flight Data for Subscale Rocket Simulated Flight Data Actual Flight Data Apogee (ft) Max Velocity (ft/s) Max Acceleration (ft/s 2 ) Time to Apogee (s) Flight Time (s) The rocket was a partial success due to a failure in the recovery system. The black powder charge that was to separate the rocket into two sections to allow the drogue to deploy actually ignited the shock cord of the drogue because it had did not have fire retardant employed far enough along it. This caused the cord to break upon drogue deployment. The rocket did not sustain any structural damage upon landing, however. 37

43 3.3 Recovery Subsystem The recovery system consists of 3 stages: deployment of a drogue parachute at apogee, deployment of a main, larger parachute at a lower, set altitude, and the landing of the rocket at a safe speed, yielding a maximum kinetic energy no greater than 75 ft lbf. The original recovery system design involved the use of Rouse Technology CO2 canisters. Unfortunately, CO2 canisters in the size and quantity required for our testing and launches could not be sourced. The team decided to downgrade our system to black powder charges to replace the CO2 canisters. The entire recovery system has a mass of lbm and is primarily guided by two Altus Metrum Telemetrum altimeters with built in GPS. The system is designed for dual deployment, with parachute deploys at apogee and at 2,000 ft. The dual deploy system is designed to reduce drift and to avoid deployment of the main at high velocity. The second round of deploys adds additional risk of failure and a possibility of simultaneous deployment, but with our dual altimeter system in place and back up charges for each deploy, we feel confident that the advantages of the dual deploy system greatly outweigh the risks. Before deployment of parachutes, the rocket sections are held together by shear pins: #4 40 nylon machine screws, ¼ in long. Two pins hold the avionics and booster sections together and four pins hold the nose cone to the avionics section. At launch the pins experience lbf. The Nylon NSM pins have a shear strength of 10,000 psi. The pins have a cross sectional area of in 2. The pins will hold the rocket together until deployment of the parachutes with a margin of safety of (calculated below). Margin of safety for the four pins at the nose cone is

44 , 10, At apogee (target height: 15,500ft), the altimeter will trigger the first black powder charge. This charge will break the two shear pins holding together the avionics and booster sections of the rocket , A force of over 156 lbf is required for separation. 5 grams of black powder will provide 164 lbf which is sufficient for separation, assuming a rocket diameter of 5 in and a compartment length of 5 feet., There are redundant charges set up for each deployment. The first altimeter is set up to deploy the first drogue and main charges. The first drogue charge will be triggered when the first altimeter reaches apogee (approximately 15,500). The second drogue charge will be triggered by the second altimeter at a slightly lower altitude to avoid the chance of simultaneous deployment. The second charge will also be stronger, to ensure that if separation does not occur on the first attempt, it will definitely occur on the second. The second charge will be 5.5 grams black powder, providing a separation force of 180 lbf., If the first charge is successful, the second charge will deploy into the open atmosphere and will not affect the flight. The main deploy will be executed at 2,000 feet in the same manner as the drogue: two charges, the second activating slightly after the first to ensure deployment. The layout of the altimeters and their basic wiring can be seen in the figure below. 39

45 The drogue parachute is 17.3 in diameter and is made from ripstop nylon. The required parachute diameter is calculated using the following formula: 8 The parachute is attached to the fireproofed shock cord using 12 shroud lines. The shock cord is made from 60 in of 1 in diameter tubular nylon and is, in turn, attached to the bulkhead with steel eyebolts. The attachment of the parachute to the shock cord, and not directly to the bulkhead, reduces the stress on the rocket. The drogue parachute reduces the velocity of the rocket to 120ft/s, the highest velocity that can be achieved before main deployment, therefore reducing drift. The main parachute deploys at 900 feet, leaving enough time from the drogue deployment for payload 1 to collect data. The parachute is 138 in diameter and is made from ripstop nylon. The parachute diameter is calculated using the same formula as the drogue. The parachute has 12 shroud lines connecting the parachute to a fireproofed shock cord. The shock cord is made from 240 in of 1.5 in diameter tubular nylon. The shroud line is attached to the bulkhead located within the rocket casing, just before the payload bay. This bulkhead is thicker that other bulkheads to ensure that it will withstand the force of the main deploy. The main parachute reduces the rocket velocity to approximate ft/s, allowing all three rocket sections of the rocket, the nosecone, avionics bay, and booster section, to land with a kinetic energy less than 75ft lbf 40

46 3.4 Mission Performance Predictions Mission Performance Criteria The launch vehicle must achieve each objective listed below in order for the mission to be deemed a success. 1. The rocket flight must be completely stable and return safely to the ground. 2. The rocket must reach the target altitude of 15,500 feet. 3. The rocket must reach supersonic speeds during ascent. 4. The BPAP system must record data for pressure fluctuations, strain forces in the body of the rocket, and thermodynamic data at the nozzle exit. 5. The drogue parachute must deploy at apogee. 6. The main parachute must deploy at 2000 feet. 7. At main parachute deployment, the TAP system must begin surveying the ground beneath the rocket. 8. Each section of the rocket must land with a kinetic energy of less than 75 ft lbf, with no damage. 9. All data recorded during flight must be completely retrievable Flight Simulations OpenRocket was used to simulation the flight characteristics that can be expected during a full scale flight. All simulations were done using the current vehicle configuration as described in preceding sections, unless otherwise noted Drift Simulations The table on the following page shows the drift calculations and apogee points for the rocket if it were launched in 5 different wind speed conditions. The rocket reaches an apogee of 41

47 17803 feet and drifts a maximum of feet in 20 mph wind with a main parachute deployment at 2000 feet. Depending on conditions during the final launch, the main parachute deployment may be delayed till a lower altitude to minimize the drift distance. To accompany the table, a graph was created showing the flight side profile with respect to time during a wind speed of approximately 0 5 mph. This graph shows the altitude at which burnout occurs, as well as the altitude locations where the drogue and main parachute deploy. Table 3.11: Drift Calculations Wind Speed (mph) Drift Distance (ft) Altitude (ft) Main Deploy Burnout Drogue Deploy Figure 3.17: Flight side profile showing altitude versus position upwind. 42

48 3.4.4 Flight Predictions Further simulations at low wind speeds provided the following flight data. This flight data simulates the mission performance that could be expected from the rocket design if it were launched in average conditions. Table 3.12: Maximum Data Values from Flight Altitude Max Velocity Max Acceleration Average Thrust T/W (ft) (ft/s) (ft/s^2) (lbf) Ratio Table 3.13: General Flight Data Velocity off Time to Flight Velocity at Main Velocity at Ground rod (ft/s) Apogee (s) Time (s) Deployment (ft/s) Impact (ft/s) The velocity off the rod is calculated using a seven foot rod. Considering the separation of the launch vehicle into three different sections and given the velocity at ground impact, the total kinetic energy imparted on each individual section is less than 75 lbf ft. The following graph visualizes the altitude and vertical velocity versus time. Further simulations were performed at the optimum weight of 68.5 lbm. These simulations were performed using the same conditions that were used in the previous simulation done for the current weight of the rocket, which provided the data in the previous tables. 43

49 Table 3.15: General Flight Data Velocity off Time to Flight Velocity at Main Velocity at Ground rod (ft/s) Apogee (s) Time (s) Deployment (ft/s) Impact (ft/s) Table 3.14: Maximum Data Values from Flight Altitude Max Velocity Max Acceleration Average Thrust T/W (ft) (ft/s) (ft/s^2) (lbf) Ratio As shown, the optimum weight reaches the target altitude precisely while also reaching Mach Though this is low and still in the transonic region, it can be ascertained that supersonic conditions occur along the nosecone of the rocket, which is the location on the rocket that is most critical to experience supersonic conditions. It should also be noted that with the given configuration of the recovery system, the rocket still generates less than 75 lbf ft of kinetic energy upon ground landing. This is ascertained by making the assumption that the rocket has fully separated into its three separate sections and that each section weighs less than 27.7 lbm, which if each section increases in weight by 22% (from the current values given in Table 3.6), they will be below that mark Drag Calculations The overall drag is shown graphically on the following page. The graph shows the different drag coefficients, along with the overall drag coefficient, versus the change in Mach number. The total drag increases from 0.40 at Mach 0.1 to 0.59 at Mach Maximum drag 44

50 experienced at this Mach number is lbf. This was calculated using the following equation 2 (3.1) where is the atmospheric density, A is the cross sectional area of the rocket, and v is the velocity of the rocket. It should be noted that OpenRocket does not provide sound data at transonic and supersonic conditions. The drag coefficients in these realms will be higher than those calculated by OpenRocket due to frictional forces from abrasive surfaces, boundary layer conditions, Mach and shock waves, and turbulent wind conditions. Because of this, the rocket will experience much higher drag than calculated, and ultimately a greater load force, but this load force should still fall well below the buckling load of the fiberglass, cylindrical fuselage. Figure 3.16: Drag Coefficient (total, frictional, base, and pressure) versus Mach number. 45

51 3.4.6 Stability Calculations The Center of Gravity is given by the following equation: CG d w d w d W nc nc s1 s1 s2 (3.2) where d s and w s represent the distance from the center of gravity to the aft end and the weight of the section, respectively, for the nose cone, section one, and section two. Using Barrowmans Equations (The Theoretical Prediction of the Center of Pressure, 1966), the center of pressure for a rocket without any conical transitions, as measured from the tip of the nose cone, was calculated using X N C C C X C X N N N F F N N N F (3.3) where is the center of pressure coefficient for a conical nose cone ( 2, is the center of pressure for the nose cone, which for an ogive nose cone is equal to X N 0.466L (3.4) N where is the length of the nose cone. The center of pressure coefficient for the fins was calculated by C N F 2 S 4N R d 1 S R 2 2Lf 1 1 CR CT (3.5) In this equation, R is the radius of the body at the aft end, S is the fin semispan, N is the number of fins, d is the diameter at the base of the nose, is the length of the fin mid chord 46

52 line, is the fin root chord, and fin tip chord. The center of pressure for the fins is calculated using X CR 2C 1 R T CRCT XF XB CR CT 3 C C 6 C C R T R T (3.6) where is the distance from the nose tip to the fin root chord leading edge and is the distance between the fin root leading edge and the fin tip leading edge, parallel to the body. Overall, the final CG and CP positions are listed below, generating a stability caliber of Table 3.16: Stability of the Vehicle Stability (as measured from the tip of the nose cone) CG Position in CP Position in Figure 3.18: Ballast Tank showing location 47

53 Given the possibility that the weight of the final rocket design may be under the optimal weight discussed earlier, a ballast tank has been incorporated into the design. This would be used to increase the weight of the rocket should the final weight not be sufficient enough to decrease the maximum altitude achieved to 15,500 feet. The ballast tank has been located specifically near the current location of the center of gravity so that in the case that weight is added, the stability of the rocket will not be affected. Calculations and simulations were performed using OpenRocket to insure that the rocket design would meet all flight criteria. Simulations showed the rocket remained completely Figure 3.19: Vertical orientation versus time for multiple wind speeds stable throughout flight under the five different wind speeds, as shown graphically below, where the vertical orientation is plotted versus time. Another graph shows the change in stability as a correlation to the change in the center of gravity location towards the forward part of the rocket as the propellant in the motor is burned and as the center of pressure increases due to increase fluid pressure from flight. 48

54 The stability margin increases from 2.93 at launch to 4.75 at motor burnout, as shown on the graph Figure 3.20: Stability versus time, showing a correlation between change in CG and CP position With such high accelerations, the internal components of the rocket will be subjected to large gravitational forces. Using OpenRocket to calculate the acceleration speeds during ascent, the gravitational forces were calculated. The maximum acceleration achieved by the rocket is 324 ft/s 2. This correlates to a G force of It should be noted, however, that at the optimum weight of 68.5 lbm, the max acceleration decreases to 258 ft/s 2, resulting in a maximum G force of The range of G forces the rocket experiences from launch to motor burnout are shown in the graph below. 49

55 Figure 3.21: G forces exerted on the rocket during flight for the current weight and optimal weight. 3.5 Payload Integration The overall design of the rocket revolved heavily around the requirements each payload proposed and the necessity to deliver each one successfully. This required the design of the rocket to allow each payload to be easily accessible and for each to be safely secured within the frame of the rocket to prevent any damage or data corruption, which would result in mission failure. The integration plan for each payload is listed below: 1. Corrosion Mitigation Payload (CMP) a. Consists of three layers of paint: the innermost layer being a heat sensitive paint; the intermediate layer will be a coating of NeverWet; the exterior coating will be commercial paint. b. Each coat will be applied evenly to the nose cone of the rocket according to manufacturers specifications. Each coat will cover the entire surface of the nose cone with constant thickness. 2. Threat Analysis Payload (TAP) 50

56 a. Consists of externally mounted camera systems that share an onboard computer system located in electronics bay #2. b. Both cameras will be mounted on the aft section of the rocket, with each lens pointing in the aft direction. Each camera will have an aerodynamic housing around it to protect it during flight. c. Onboard computer will be mounted on a sled internally of electronics bay #2; camera connections will pass through small openings of electronics bay #2, pass thru openings cut into the side of the rocket, which will be covered by the aerodynamic housing. 3. Boost Phase Analysis Payload (BPAP) a. Consists of three analysis systems: temperature, pressure, and strain gauges. b. Strain and pressure gauge system are connected to a single Omega Bridge/Strain Gauge Data Logger located inside electronics bay #1. c. The data logger will be placed inside cylindrical, foam cutouts that will fit within the tube coupler to prevent movement of the data logger. d. Two pressure transducers and two strain gauges will record pressure and strain differentiations during ascent of vehicle. e. An Omega Temperature Data Logger will be located within electronics bay #2. It will also be located within a cylindrical, foam cutout that will fit within the electronics bay #2 tube to prevent movement of the data logger. f. Two thermocouples will be connected to the nozzle exit of the rocket. 51

57 3.6 Launch concerns and operation procedures 3.6.1Recovery preparation. In order to ensure recovery deployment, the parachute must be packed correctly. This will consist of folding the parachute along the shroud lines, in order to ensure that they line up, and unfurl properly. Then, in order to ensure that the shroud lines do not wrap around the parachute, they will be placed inside the fold of the parachute and wrapped inside of the parachute, ensuring that the parachute will unfurl before the shroud lines, giving the best possibility of full unfurling of the parachute. Finally, the parachutes will be packed in their respective compartments, as outlined in the previous subsystem reviews Motor preparation. To install the motor on site, the motor will be mounted constructed by the Level 3 mentor, Christopher Short, since no team member has the certification required to construct the motor. The motor will then be mounted within the motor retention tube. Finally, the team will install final retention in the form of the 94mm aeropack retaining fixture screwed into the bulkplate Igniter installation. The igniter will be installed while the rocket is on the Launchpad, in order to absolutely insure no accidental ignition of the motor in a hazardous location. After the rocket is secured on the Launchpad, the igniter will be installed as far into the motor as possible, in order to obtain the maximum rate of success in ignition of the motor. Finally, before leaving the Launchpad, continuity will be tested in order to make sure that the igniter is properly connected electrically to the controller Setup on launcher. In order to set the rocket on the launch pad, the rocket will be carried out to the launch pad by the team, accompanied by the team s level 3 mentor, Christopher Short. The rocket will be installed onto the launch rail using the large rail lugs installed onto the rocket in order to ensure that the rocket obtains enough velocity on the rail to be stable, as well as ensuring the launch angle of the rocket is the desired launch angle. The altimeter will then be armed once the rocket is secured on the rail, and verified that it is functioning using the groundstation and receiver supplied by Altus Metrum. Finally, the igniter will be installed, and continuity will be tested as per

58 3.6.5 Postflight inspection. Once the rocket lands under power of recovery, the team will inspect the rocket for any external damage. Then, the separate components will be individually taken apart and verified that no damage was sustained. Finally, the rocket will be dissembled and the relevant scientific data will be collected. 3.7 Safety Safety Officer The team s safety officer is Jacob Herrera, a Junior in Aerospace Engineering, and working towards a level one High Powered Rocketry certification as well as a full OSHA certification Preliminary Failure Mode The primary failure modes of the rocket design, payload integration, and launch operations have been listed and addressed in each sub system review. These sections cover all notable issues and how they have been assessed and mitigated to ensure the success and safety of the project Hazards Personnel hazards have been accounted for and can be seen in each sub system review. A sample of the MSDS s can be found in Appendix B of this document. More relevant MSDS s are currently being gathered and are actively being added to the comprehensive list on the safety portion of the website. NAR regulations and risk mitigation tables can be found in Appendix E where all primary concerns have been addressed Environmental Concerns Wind poses the most significant threat to the safety of the launch vehicle. If wind is a persistent factor when launching, proper mitigations will be made to delay the deployment 53

59 of the main parachute to limit the amount of drift that occurs. Drift calculations done within the airframe section are being used to ensure the launch vehicle returns within the flight box. 54

60 Section 4: Payload Criteria 4.1 Payload Overviews Threat Analysis Payload (TAP) TAP will serve as a threat detection system that will identify possible threats and hazards in a landing zone. It will accomplish this by collecting images from two externally mounted cameras on the aft section of the rocket. These cameras will relay the images to an onboard computer system that will analyze the images for possible threats using a integrated software package designed and programed by the team. Once the images are analyzed, the computer system will transmit the data via radio transmitters to a ground station where continued analysis may be conducted. All data transmissions will be conducted in real time. In order to ensure the survivability of the externally mounted cameras, shielding casings were designed to protect both cameras from the adverse effects of supersonic flight. The shields were designed to allow an unobstructed view of the landing zone while retaining the minimal drag characteristics of the rocket Boost Phase Analysis Payload (BPAP) BPAP is designed in order to verify certain parameters of the rocket and thus gain valuable information about its flight characteristics. In order to do this, a series of sensors will be placed on critical subsystems within the rocket in order to gather information regarding the motor, electrical systems, and airframes. This information will be collected by a DAQ system onboard the rocket, and offloaded once safely recovered Corrosion Mitigation Payload (CMP) CMP will analyze the effects of supersonic flight on paint and coatings. The selected coating was determined to be NeverWet, a commercially available anti icing/corrosion coating often 55

61 used on aircraft wings. Supersonic flight holds new obstacles to overcome as paint and coatings must withstand the harsh effects of supersonic flight. Therefore, NeverWet will be tested for its effectiveness after enduring supersonic flight. A layer of heat sensitive paint will be painted beneath a layer of NeverWet to determine the effectiveness of the coating after supersonic flight. The heat sensitive paint will undergo a color change when exposed to changes in temperature such as the introduction of water droplets or the presence of external heat. Therefore, this payload will also test the effectiveness of heat sensitive paint in determining the heat profile of the surface of a rocket while experiencing supersonic flight. 4.2 System Level Review Payload Requirements and Objectives All payloads must be recoverable and reusable 1. Threat Analysis Payload (TAP) a. Using a camera system, scan a landing zone to detect possible threats and hazards b. Data from the camera system must be analyzed by a onboard software package for any potential threats and hazards c. All data and results from the onboard software system must be transmitted in real time to a ground station for analysis 2. Boost Phase Analysis Payload (BPAP) a. Analyze both the structural and dynamic systems during boost phase b. Identify the conditions of the structural and dynamics systems of the rocket during boost phase c. Identify motor conditions during boost phase d. Store data for future analysis 3. Corrosion Mitigation Payload (CMP) a. Testing the effects of supersonic flight on NeverWet in order to determine its viability during transonic and supersonic flight 56

62 b. Adaptation of heat sensitive paint in order to test it s viability as a model of areas where adverse supersonic effects are greatest along the surface Advantages and Disadvantages of Each Payload Selection 1. Threat Analysis Payload a. Advantages i. Multiple cameras and power sources will provide system redundancy allowing for higher mission success. Onboard, integrated software will analyze images in real time in order to provide accurate threat analysis. b. Disadvantages i. Mounting the camera externally exposes the system to adverse effects during the ascent phase, especially during supersonic flight. Shielding must designed to protect the system, possibly increasing the drag force experienced by the rocket. Additionally, the material used in the construction of the rocket must not interfere with the data transmission to the ground station, limiting material choice unless an external transmitter is mounted. 2. Boost Phase Analysis Payload (BPAP) a. Advantages i. Provides valuable data during boost phase b. Disadvantages i. Complexity of integrating the sensor package and making sure all sensors remain properly calibrated during supersonic flight 3. Corrosion Mitigation Payload a. Advantages i. Simplicity. The payload requires no complex electrical components or integration. Payload is easily applied to the rocket and tested by visual and tactile inspection. b. Disadvantages 57

63 i. Because of the nature of visual and tactile inspection, analysis is variable and results may be inconsistent. Weather effects, impact during landing, and launch debris may all cause inconsistences in the analysis 4.3 Payload Subsystems Required Subsystems for Mission Success 1. Threat Analysis Payload a. Integrated camera system in order for image capture b. Onboard software for image analysis c. Integrated computer system for software housing and data interpretation d. Onboard radio transmitter for data transmission 2. Boost Phase Analysis Payload a. Properly calibrated sensor package b. Integrated computer system for data analysis and storage 3. Corrosion Mitigation Payload a. No major subsystems required since data analysis will be done from visual and tactile inspection of the rocket after recovery. 4.4 Payload Design and Integration Threat Analysis Payload The Threat Analysis Payload (TAP) will consist of externally mounted camera systems that will detect oncoming hazards as the rocket approaches the landing zone. The payload will consist of two externally mounted cameras linked to a shared onboard computer system. Both cameras will be mounted on the aft section of the rocket, with the each lens pointed downward, allowing for a clear view of the landing zone as the rocket descends. Since both cameras are mounted externally, visual interference and diffraction of light is minimized in the images Figure 4.1: Raspberry Pi [8]. collected. Both cameras will take images of the landing zone, 58

64 relaying them to the computer system where the integrated software package will analyze the data for threats and hazards. The software package will consist of a two dimensional array of floating point values that will overlay a snapshot. Edges and color differentiations will be identified in the picture in order to detect landing hazards. Potentially hazardous spots detected in the picture will be represented by higher numeric values. Therefore, a high confidence hazard such as a large building may be represented by the value 10. A lower confidence hazard such as a bird may be represented by the value 2. The onboard software will be implemented in Java and will run on a Raspberry Pi, as depicted in Figure B. During the rocket s descent, the software will transmit a continuous feed of data encoded in JSON to a ground computer. The computer on the ground will store the data in a Postgres database. This will allow the data to be analyzed at any point during or after the descent. In order to ensure, the success of the payload, redundant power systems will be placed along side the integrated computer system in the payload bay in the form of lithium ion batteries. At a minimum, two lithium ion batteries will be placed in the payload bay to serve as a primary and secondary power source for the TAP. 59

65 Figure 4.2: Camera System Integration and Shielding Design In order for the camera system to survive the ascent and supersonic flight, an aerodynamic heat shield will provide protection for the externally mounted cameras. This will ensure their survivability during the ascent phase, allowing for threat analysis as the rocket approaches the landing zone. In order to minimize the drag induced by this shielding, it has been designed as thin as possible without interfering with the camera s field of view. Therefore, the camera must be small enough to allow minimal aerodynamic effects resulting from the shielding Boost Phase Analysis Payload The second payload configuration will consist of one thermocouple attached to the motor casing near the nozzle, along with two pressure transducers recording pressure on the outside surface of the rocket and two strain gauges attached to the inside surface of the rocket, as shown in the figure below. 60

66 Strain Gauges Pressure Gauges Thermocouple Strain/Pressure Data Logger Thermocouple Data Logger The strain gauges are Omega KFH series pre wired gauges that are self adhesive. These will be placed on the inner walls of the vehicle in the locations shown above. They will be attached to a single Omega Bridge/Strain gauge data logger located within the payload bay that will sample the data at 20 Hz processing speed. The pressure transducers are miniature flush diaphragm transducers that will be placed within a hole drilled out in the spots indicated in the figure above. These will record pressures directly behind the nose cone where expansion fans may form in higher Mach number conditions, and also downstream in normal flow conditions. The transducers will also be attached to an Omega Bridge/Strain gauge data logger located within the payload bay and have data sampled at 20 Hz. The thermocouple will be Omega ready made insulated thermocouples with glass braid insulation and molded connectors. The thermocouple will be adhesively bonded to the circumferential surface of the motor casing at the nozzle exit. The wire will be incased in a sleeve from the point of contact to the surface of the motor to the avionics bay to protect it from high temperatures. The thermocouple will be connected to an Omega ambient temperature and thermocouple data logger within the avionics bay. Since this data logger 61

67 has the ability to record ambient temperature of its environment, it will be recording the temperature within the avionics bay for analysis of the environmental change that all avionics equipment must undergo during flight. All data will be sampled once per second. All data gathered during ascent will be transferred to a computer at the ground station via USB upon landing. This will be accomplished by having a USB splitter cable within the body of the vehicle, which will then be connected to the ground station computer. The thermocouple attached to the motor casing will read an increase in temperature as the motor burns. Assuming the temperature of the casing is equivalent to the temperature within the motor, the exit velocity of the exhaust of the motor can be calculated using the equation below: 2 where the exit temperature is the ambient temperature, which will be recorded from the ground station. Using isentropic equations, the temperature within the motor, as well as the ambient temperature and pressure (as recorded by the ground station), the total pressure within the motor can also be calculated. 62

68 Using the total pressure of the motor along with other motor parameters that we already know, such as specific impulse, further calculations can be made to find parameters such as mass flow rate, thrust, effective exhaust velocity, and thrust coefficient where A is the exit area of the nozzle and M is the Mach number at the nozzle, which can be calculated using the isentropic equations for temperature or pressure Corrosion Mitigation Payload The Corrosion Mitigation Payload (CMP) is designed to test the effectiveness of the of the NeverWet anti icing system at supersonic speeds. Additionally, NeverWet is used to prevent corrosion of surfaces exposed to moisture and salts. Furthermore, the effectiveness of heat sensitive paint as a means of determining areas of high heat concentration along a rocket s body will be tested as well. The design of this system is relatively simple as all it requires is coating the exterior of the rocket with layers of the respective paint and/or coating to be tested. 63

69 Figure 4.3: Corrosion Mitigation Payload Design and Integration The payload will consist of three layers of paint as illustrated in Figure X. The innermost layer will be a heat sensitive paint that will be used to determine the effectiveness of NeverWet after undergoing supersonic flight. The intermediate layer of will be a coating of NeverWet. The exterior coating will be commercial paint for aesthetic purposes. Once the rocket is recovered, water will be applied to the surface of the rocket. If the heat sensitive paint changes color, then the effectiveness of NeverWet has been compromised. Additionally, the heat sensitive paint will identify patches of the rocket s surface that encountered intense heat resulting from supersonic flight not only identifying the areas of NeverWet failure, but also profile the heat exposure of the surface of the rocket. Visual and tactile inspection of the rocket will be conducted in order to verify the mission success of the payload. The payload may be repeated on site given time for reapplication of the necessary coatings to affected areas. 64

70 If NeverWet is found to maintain its integrity after supersonic flight, it could serve as an alternative to complex heating systems in order reduce the damage from icing and long turn exposure to moisture and salt. In conjunction with heat sensitive paint, the two could be used as a cost effective alternative to prevent damage from moisture but also provide a heat profile of the flight or space system. 4.4 Verification Plan and Performance Metrics Verification Plan In order to ensure the requirements of each payload are fulfilled, a verification plan was established. Table X outlines how each stated requirement as listed in Section 4.1 will be fulfilled and how that fulfillment will be verified. Table 4.1: Payload Verification Payload Requirement Outlined Requirement Verification Identification Requirement Fulfillment Method TAP will, using a Mounting and Threat Analysis Payload Requirement A camera system, scan a landing zone to detect possible integrating both external cameras with their shielding Analysis and Testing threats and hazards components 65

71 TAP will collect data from the camera Integrating a Threat Analysis Payload Requirement B system and analyze it using an onboard software package in order to analyze any Raspberry Pi computer system with onboard software into the Analysis and Testing potential threats and payload bay hazards Threat Analysis Payload Requirement C TAP will transmit all data analyzed by the onboard software to the ground station for analysis in real time Radio Transmitters will be used to transmit data to the ground station in real time Analysis and Testing Corrosion Mitigation Payload Requirement A Testing the effectiveness of NeverWet under supersonic conditions Applying NeverWet to the surface of the rocket and verifying that it endures the effects of supersonic flight Inspection and Testing 66

72 Applying heat Corrosion Mitigation Payload Requirement B Testing of heat sensitive paint as a viable way of analyzing the adverse effects of supersonic flight on a rocket's surface sensitive paint beneath the layer of NeverWet to determine the highest concentrations of heating effects associated with Inspection and Testing supersonic flight All systems are designed to withstand the All Payloads Must be recoverable and reusable adverse effects of supersonic flight, Analysis and Testing impact from recovery, and other possible scenarios Performance Expectations and Metrics In addition to the verification each payload s requirements, performance metrics have been designed in order to assess the payload s performance and whether or not the mission criteria was fulfilled. Threat Analysis Payload Performance Expectations Camera system must be able to endure the effects of supersonic flight. Camera system must have unobstructed view of the landing zone. 67

73 Onboard software package must correctly identify threats and hazards on the ground. Potential threats and hazards must be relayed to the ground station in real time. Performance Metrics Testing both camera s image capture ability and detection capability of potential threats. Testing shielding ability to protect camera system from supersonic effects. Testing integrated software for consistency in ability to perceive and identify threats and hazards. Testing radio transmitter s ability to properly relay real time information to a ground station. Corrosion Mitigation Payload Performance Expectations Effectiveness of NeverWet after enduring supersonic flight. Effectiveness of heat sensitive paint in determining the effects of supersonic flight. Performance Metrics Testing and inspection of NeverWet s anti icing and corrosion ability by applying water before and after launch in order to determine whether or not it was able withstand supersonic flight and still perform. Testing the ability of heat sensitive paint to identify areas of highest heat concentrations experienced along the surface of the rocket. 4.5 Payload Concept Features and Definition Creativity, Originality, and Uniqueness The Threat Analysis Payload provides a simple solution in order to effectively analyze hazards as the rocket approaches the landing zone. The design is mounted externally in 68

74 order to minimize any compromise in structural integrity normally associated with transparent sections. Additionally, the camera system is designed to withstand the effects of supersonic flight without mounting the cameras within the body of the rocket. Furthermore, onboard custom software will analyze the data and relay the information to the ground station in order for future analysis. Unique shielding techniques have to be designed in order to protect the camera system from the effects of supersonic flight. The proposed shielding design is also designed to minimize aerodynamic drag while still providing optimal protection for the Threat Analysis Payload to perform. Boost Phase Analysis Payload The data gathered will represent various flight performance parameters and characteristics of the rocket. All performance data will be analyzed and put through a Matlab code, which will generate graphics representing the flight parameters mentioned above for further analysis to insure the flight data matches closely with the data generated during simulations under flight conditions, such as wind speed, temperature, and barometric pressure. The Corrosion Mitigation Payload (CMP) is unique in its implementation and its simplicity. There are no complicated electronic systems to integrate or design. It simply uses commercially available coatings and paints in order to test their effectiveness after undergoing supersonic flight. If determined to be viable, the coatings could be used a cost effective alternative to complex heating systems to prevent icing and damage from moisture Level of Challenge The Threat Analysis Payload has several design challenges in the integration of the system. Mounting the cameras to the exterior of the rocket presents a design challenge in how a shielding process would need to be developed in order to protect the system from the effects of supersonic flight. Designing this shielding is a design challenge because the shielding must be substantial enough to protect the payload while maintaining minimal increases in drag. Additionally the custom made software provides a design challenge, as the software team 69

75 must build an integrated software package from scratch that will analyze potential threats and relay that information back into to the ground station. Designing the structure of the rocket to ensure that there is no interference when data is transferred to the ground station presents another challenge, as material selection will have to accommodate for data transfer. The Corrosion Mitigation Payload has limited challenges given its simplicity. The biggest challenge faced will be minimizing any damage to the coatings from launch debris, weather effects, and damage from impact. Therefore, the team must ensure a slow enough descent rate that the coating will not be affected during landing. This will insure accuracy and consistency of mission success evaluations. 4.5 Science Value Payload Objectives The Threat Analysis Payload (TAP) is an experimental software package that could be used to identify threats and hazards in spacecraft landing zones. The software package could eventually be integrated with the propulsion system in order to perform evasive maneuvers to avoid any identified threats or hazards. The objectives of the TAP is to identify potential threats and hazards in a landing zone through onboard analysis of captured images and transmitting potential threats to a ground station. The Corrosion Mitigation Payload (CMP) is testing the effectiveness of the commercially available NeverWet under supersonic flight conditions. If deemed viable, NeverWet and similar products could prove to be a cost effective alternative to complex and expensive heating systems that prevent icing and damage from moisture. In addition, the heat sensitive paint used to determine the effectiveness of NeverWet after enduring supersonic flight could be used as a method in determining the heat profile an air or spacecraft undergoing supersonic flight. 70

76 4.5.2 Payload Success Criteria Table 4.2 describes the payload mission success criteria. There are three levels of success. Complete mission success corresponds with a numerical value of 3, partial mission success corresponds with a numerical value of 2, and complete mission failure corresponds with a numerical value of 1. Decimal values may be assigned based on the results collected to increase the accuracy of mission success determination. For example, if the TAP is only able to identify 80 percent of the landing zone threats, but accomplishes the remaining success criteria, the payload success rating could be rated as a 2.7, allowing for more accurate descriptions of mission success. Table 4.2: Payload Success Criteria and Rating Payload Level of Success Success Criteria Payload is able to identify all hazards or threats with in the Threat Analysis Payload Complete Mission Success (3) landing zone, relay information back to the ground station, and remain undamaged for immediate reuse. 71

77 Payload was not able to identify some potential hazards or threats, some information was not Partial Mission Success (2) relayed to the ground station or was indecipherable due to interference, or payload was damaged requiring repairs. Payload was not able to identify any hazards or threats, no information Complete Failure (1) was relayed to the ground station, or payload was damaged beyond repair. Corrosion Mitigation Payload Complete Mission Success (3) Payload was able to determine the effectiveness of both NeverWet and heat sensitive paint during supersonic flight. Results are clear and well defined. 72

78 Payload was partially able to determine the effectiveness of either Partial Mission Success (2) NeverWet and heat sensitive paint during supersonic flight. Results were available but may be inconsistent. Payload was not able to determine the effectiveness of both Complete Failure (1) NeverWet and heat sensitive paint during supersonic flight. Results were either unclear or not available Experimental Logic, Approach, and Method of Investigation The Threat Analysis Payload s primary function is to detect and identify threats in the landing zone and then transmitting those results to the ground station in real time. In order to ensure that the images are properly analyzed, an onboard software package is integrated into a dedicated computer system used solely for the purpose of threat detection and hazard identification. This dedicated computer system ensures that all its processing power will be used for the sole purpose of analyzing image data and transmitting those results to the ground station, eliminating any possible issues arising from data interference from other payloads or the loss of processing power that could happen if there was only one computer system handling all the subsystems of the rocket. Additionally, redundant power systems provide a level of security against mission failure in the case of an electrical power failure. 73

79 The camera system selected is to be mounted externally in order to eliminate the need for a transparent section of the rocket, which could possibly lead to structural instability. Additionally, the implementation of an external camera system reduces the need for complicated, internal system integration. Shielding for the external camera systems provides protection to the system while maintaining the aerodynamic profile of the rocket. The Corrosion Mitigation Payload takes a more qualitative approach to meet its primary function. In order to test the viability of NeverWet as an appropriate corrosion prevention system on supersonic systems, it must be applied to the exterior of the rocket and flown at supersonic speeds. Testing will occur after recovery of launch vehicle by adding water droplets to the exterior of the rocket to determine whether or not the NeverWet system has been compromised. If it has been compromised, the underlying coat of heat sensitive paint will experience a color change indicating the presence of water on the surface of the rocket. Additionally, supersonic flight effects could be so great that NeverWet would erode away. The heat sensitive paint would indicate where such areas occurred allowing for a thermal profile of the rocket to be analyzed. Results will be verified qualitatively by a visual and tactile inspection of the rocket after recovery Test and Measurement, Variables, and Controls Every subsystem of the rocket will be tested to ensure mission success. Testing will mostly take place in the 3 wind tunnels here at Auburn University. The supersonic wind tunnel will be used to test the effectiveness of the Corrosion Mitigation payload. From testing, variations of paint and coating layers may be determined to be most optimal. The Threat Analysis Payload will be tested via demonstration trials that will test the software s ability to identify potential threats in the landing zone. Additionally the radio transmitter will be tested in order to ensure accurate data transmission. 74

80 Scaled flight test and wind tunnel testing will serve as controls for comparison to full scale launches. Having scaled test flights will ensure the viability of the rocket design before integrating the expensive electrical subsystems. These subsystems will be tested in the lab first through simulations before integration onto the full scale design. Measurements will be taken during wind tunnel testing and electronic simulation in order to ensure that each subsystem is fulfilling the mission criteria. Variables to testing include inclement weather, availability of equipment, and budget constraints. However, these variables will be mitigated by careful project planning and precise measurements of subsystem performance to ensure efficiency of the design process Relevance of Expected Data and Accuracy/Error Analysis The Threat Analysis Payload s accuracy and error analysis will be governed largely by the onboard, integrated software package. Given the experience of the software team, the software package should be able to analyze all the images for possible threats without major issues. Possible error could occur in identification of smaller hazards, but overall the system should perform as designed given all equipment is properly integrated into the rocket. Accuracy of the results will be evaluated by the performance metrics listed in Section and the success criteria listed in Section Any resulting error will be analyzed at each subsystem in order to identify the issue and then develop solutions to mitigate future errors. Since the Corrosion Mitigation Payload is evaluated by a largely qualitative approach, consistency of results will be determined on the team s adherence to the success criteria and performance metrics listed in Section and Preliminary predictions would indicate that NeverWet could fail at certain areas where supersonic flight creates enough friction and heat to erode the coating away. However, only supersonic testing will confirm or disprove these predictions Preliminary Experiment Process Procedures 75

81 The Threat Analysis Payload will be tested in 3 stages, with each stage representing another phase in design and integration level readiness. The first stage will be the assembly of a subscale payload in order to test the viability of the threat analysis software at identifying threats under controlled conditions. A subscale payload will then be developed and integrated onto a scaled rocket for preliminary flight testing. The final stage will be the integration of the full payload onto a full scale rocket for a full scale launch to test the payload s performance before competition. This will allow for any adjustments to occur before final design integration. The preliminary experiment process procedure of the Boost Phase Analysis Payload will be to test all equipment, including thermocouples, pressure transducers, and strain gauges, to verify that data acquisition is possible, and if so, that data gathered is consistent over multiple runs to solidify the relevance of the data recorded. The Corrosion Mitigation Payload will be tested in 3 stages, with each stage representing different levels of design and integration level readiness. The first stage will be the testing of the coating layers under supersonic conditions. These conditions would simulated via Auburn s supersonic wind tunnel and models placed with the wind tunnel. The second stage would be the integration of the payload onto scaled flight test. The final stage of testing would be the implementation of the payload onto a full scale test flight in order to evaluate the payload s performance prior to competition. This will allow for any adjustments to occur before final design integration. 4.6 Safety and Environment 76

82 Risk Matrix Analysis for TAP Payload 9 6 Probability Impact on Project 1. Shielding on the outside of the rocket could fail, resulting in a loss of the external camera. 2. Camera fails to initialize during the appropriate time of the descent. 3. Batteries could fail during the flight, resulting in the raspberry pi failing to initialize. 4. G forces during ascent could cause instrumentation to be displaced or dislodged from any holdings or fixtures inside the rocket resulting in corrupt data. 77

83 5. Software works improperly, failing to deliver the mission critical data to the groundstation. 6. Displacement of instrumentation could result in critical damage of other necessary components parts such as altimeters, leading to failure of recovery system or causing the recovery system to deploy improperly (too late or too early). 7. Size of instrumentation could be larger than anticipated resulting in internal components being moved from original design location or causing insufficient room within the bay areas. 8. Camera begins recording information prematurely during ascent. 9. Instrumentation could be damaged during transport or setup of rocket. 10. Recovery system deployment could damage or destroy instrumentation resulting in corrupt data or inability to transmit data. Mitigation for Risks Risk Number Mitigation Shielding will be statically loaded to calculated maximum in flight force in order to verify that the shielding will not fail. The setup will be ground tested prior to installation into the rocket in order to verify that the system is functioning as intended. Batteries will be ground tested prior to installation in to the rocket in order to verify that the batteries are charged, and spare batteries will be on site in the event the batteries are not charged. The bulkplates will be fastened securely in the airframe in order to minimize vibrations within the payload bay, minimizing the risk to the fragile electrical components. Software will be ground tested on site prior to installation into the rocket to verify that the system is working as intended. The bulkplates will be fastened securely in the airframe in order to minimize vibrations within the payload bay, minimizing the risk to the internal components of the rocket. The payload will be built well in advance in order to compensate for any unanticipated size increases 8 Ground Station will be prepared to handle extra data from the system 9 The entire system will be packaged carefully to avoid damage during transport The bulkplates will be thick enough to widstand the effects of the charges deploying 10 the parachutes. 78

84 Risk Matrix Analysis for BPAP Payload Probability Impact on Project 1. Thermocouple could become unattached from motor casing / nozzle causing data analysis to be interrupted or fail. 2. Data logger for the thermocouples fails to record data because connection of thermocouple to data logger was not placed securely. 79

85 3. Any of the data loggers could lose charge from batteries due to being left on the pad for extended periods of time, heat elements decreasing the overall battery life, or faulty batteries from original manufacturing. 4. G forces during ascent could cause instrumentation to be displaced or dislodged from any holdings or fixtures resulting in corrupt data. 5. Heat from motor casing could melt the thermocouple wiring or corrupt the data. 6. Displacement of instrumentation could result in critical damage of other necessary components parts such as altimeters, leading to failure of recovery system or causing the recovery system to deploy improperly (too late or too early). 7. Data logger for pressure transducers and strain gauges fails to record data due to connection being insecurely fastened. 8. Size of instrumentation could be larger than anticipated resulting in internal components being moved from original design location or causing insufficient room within the bay areas. 9. G forces during ascent could cause damage to instrumentation / data loggers. 10. Data loggers fail to begin recording data at lift off. 11. Data loggers seize recording information prematurely during ascent. 12. Instrumentation could be damaged during transport or setup of rocket. 13. Recovery system deployment could damage or destroy instrumentation resulting in corrupt data or inability to retrieve data. 14. Data gathering during post flight phase could fail due to damage to USB cords within the rocket during flight from heat melting the wiring or to high of G forces. 15. Incorrect instrumentation could be shipped from selling company resulting in loss time. 16. Instrumentation could be lost during transport from construction facilities to testing areas or competition area. 17. Instrumentation could be faulty, such as having defects from seller. 18. Pressure transducer could fail from large pressure fluctuations during ascent. 19. Recovery system deployment during descent could cause jettison of analysis equipment due to insecure placement within the vehicle. 20. Instrumentation could record invalid data during testing or competition flight due to internal defects, hardware issues, etc. 21. Backup instrumentation may not be readily available if damage occurs. 22. Required instrumentation may not be acquired due to insufficient budget. 23. The ground station computer could fail to retrieve data from data loggers due to damage, hardware issues, etc. 80

86 Risk Mitigation for BPAP Payload Risk Number Mitigation The thermocouple will be securely attached and grountd tested in order ensure that it is securely mounted. The setup will be ground tested prior to installation into the rocket in order to verify that the system is functioning as intended. Batteries will be ground tested prior to installation in to the rocket in order to verify that the batteries are charged, and spare batteries will be on site in the event the batteries are not charged. The bulkplates will be fastened securely in the airframe in order to minimize vibrations within the payload bay, minimizing the risk to the fragile electrical components. The setup will be tested and the casing will be ground tested in order to verify that the heat from the motor casing will not melt the thermocouple. The bulkplates will be fastened securely in the airframe in order to minimize vibrations within the payload bay, minimizing the risk to the internal components of the rocket. The bulkplates will be fastened securely in the airframe in order to minimize vibrations within the payload bay, minimizing the risk to the internal components of the rocket. The payload will be built well in advance in order to compensate for any unanticipated size increases The bulkplates will be fastened securely in the airframe in order to minimize vibrations within the payload bay, minimizing the risk to the internal components of the rocket. 10 The setup will be ground tested prior to installation into the rocket in order to verify that the system is functioning as intended. 11 Ground Station will be prepared to handle extra data from the system 12 The entire system will be packaged carefully to avoid damage during transport 13 The bulkplates will be thick enough to widstand the effects of the charges deploying the parachutes. 14 The setup will be tested and the casing will be ground tested in order to verify that the heat from the motor casing will not melt the USB The payload will be built well in advance and tested in order to compensate for any defects in manufacturer. The entire system will be packaged carefully to avoid damage or loss during transport 81

87 The payload will be built well in advance and tested in order to compensate for any defects in manufacturer. The pressure transducer will be purchased to accommodate the largest flucuations possible. The bulkplates will be fastened securely in the airframe in order to minimize vibrations within the payload bay, minimizing the risk to the internal components of the rocket. The payload will be built well in advance in order to compensate for any unanticipated size increases The team will treat the payload carefully to avoid unanticipated extras, as well as certain extra 21 sets of hardware ordered so that if mistakes due occur, there are backups. 22 Sufficient funds will be raised to ensure that the payload is built. 23 Software will be ground tested on site prior to installation into the rocket to verify that the system is working as intended. 82

88 Section 5: Project Plan 5.1 Budget Plan The budgeting plan has some significant changes, as the preliminary estimates of some of the costs of items was low. Given that the funding plan still vastly exceeds the costs of the project, the team is still confident that the funding should not prove problematic during the project. Table 5.1: Estimated Competition Expenses Item Price Rocket Planning and Manufacturing $7, Total Education Outreach Expenses $3, Total Travel Expenses $ Total Estimated Expenses $14, Table 5.2: Itemized Budget Subsystem Item Price Quantity Total Price Rocket Cytek FiberGlass 37.99/yard 2 $300 Nomex Honeycomb 222/sheet 2 $ Release Fabric 38.99/yard 0.5 $19.50 Breather Fabric 90.99/yard 0.5 $45.50 Vacuum Bag Sealant Tape 15.99/roll 2 $

89 Perforated Release Film 17.99/yard 2 $35.98 Motor 1 $ Non Perforated Release Film 18.99/yard 2 $37.98 General Hardware (nuts, bolts, washers, etc) N/A $ PVC Mandrel 1 $ Assorted Electronic Equipment (wires, sodering tools, etc.) N/A $ Epoxy N/A $ Shear Pins N/a $50.00 Recovery Parachute Shock Cord 4.77/yard $84.26 Ripstop Nylon Cloth 1.5/sq. yard $ Altimeter N/A 1 $ Rouse Tech CD3 system N/A $200 Shroud Line Tubular Nylon $7.37/ ft 20 $150 Payload 1 GoPro Camera N/A 1 $ Rasberry Pi N/A 1 $35.00 Lithium Ion Battery N/A 2 $65.84 Payload 2 Strain Gauge w/ Leads (10 pcs) $ Pressure Transducer $ Thermocouple $

90 Shielding $40.00 Temperature Data Recorder $ Pressure Data Recorder $ Strain Data Recorder $ Payload 3 Paint N/A 1 $50.00 Sealer/Primer N/A 1 $30.00 Clear Top Coat N/A 1 $30.00 Hydrophobic NeverWet Coating N/A 1 $75.00 Hydrochromatic Undercoating N/A 1 $ OverHead for Error Estimated Total Price $ Travel Plane tickets to Utah $ $ Hotel Rooms $ $1,650 Transportation of Rocket $100 1 $ Funding Plan The funding Plan remains, in its most integral form, essentially the same. With only one additional sponsor since the proposal, and most of AURA s needs met, the only significant change was the amount being invested by the Alabama Space Consortium, and was thus updated to reflect both changes. 85

91 Table 5.3: Estimated Funding Source Amount Alabama Space Consortium / Engineering Dept $15, NASA TBA Auburn University Rocketry Association Account $2, AIAA Solid Rocketry $ Total Estimated Funding $18,

92 5.3 Timeline 87

93 5.4 Educational Engagement The educational outreach program consists of three components that will target three different student age groups and reach approximately 800+ middle school students in the Auburn community. The three components: 8 th Grade Rocketry Unit A small group of students in the 8 th grade at Auburn Junior High School have a short enrichment class period during which the students are allowed to study different topics that interest them outside of the general curriculum. AURA has been given permission to implement a physics and rocketry unit with the students for a week at the end of March. One to two lesson plans are currently being constructed that will cover Newton s Laws of Motion and how they apply to rocketry. The students will also build and launch Alpha rockets during the unit at the end of the week. This small scale rocketry unit will double as a practice run for the larger scale rocketry unit that will be implemented for an entire class of 7 th graders th Grade Rocket Week A large scale rocketry outreach program similar to the 8 th grade rocketry unit will be carried out with the entire 7 th grade class at Drake Middle School. This represents a body of approximately 775 students. The school informed AURA that the 7 th grade science curriculum focuses primarily on life science, and a request was made to relate the rocketry unit to the life science curriculum. The program will last three days with a fourth day scheduled in the case of weather or other miscellaneous interruptions or delays. All parts of the program will take place during the 7 th grade science classes through the each day. On a date to be determined, the science teachers at Drake Middle School have agreed to attend a workshop during a school day in the middle of March. This workshop will cover the outline of the 88

94 program, the lesson, Alpha rocket construction, and launch day procedures. Tentatively scheduled for the first week of April, the first day of the program will feature a short lesson on topics like animals in space, G forces, and what happens to the human body in flight and in space. The students will be given a short homework assignment to complete that will engage them to review some things they learned in the lesson. After the lesson, AURA team members will begin guiding the students in the construction of the Alpha rockets. On the second day, the students will finish up with construction. As the students work on finishing up construction, AURA team members will be emphasizing how important teamwork is and how the activities they are participating in relate the real world of engineering. Safety guidelines and procedures for launch day will also be covered with the students. On day three, all sciences classes will meet in a field located on Drake Middle School s campus to have their rockets launched. There will be multiple launch pads set up and plenty of AURA members present to ensure that all students get to see their rockets launched. Finally, feedback inventory forms will be distributed to the teachers and administration who participated in order for them to assess the program and make suggestions for things that could be improved the next year. 2. Auburn University College of Engineering E Day On February 28, 2014, Auburn University College of Engineering hosts E Day, an open house preview day for prospective Auburn engineering students to come explore all that the college has to offer, including engineering student organizations. AURA will have a kiosk set up for students to stop by and learn about Auburn aerospace engineering, the student rocket organization, and USLI. This exposure is one of the greatest opportunities for High School students to gain an appreciation of the real challenges of engineering, as well as get some hands on experience as to the products delivered from an engineering endeavor. To that effect, the handbook in Appendix A demonstrates the guidelines given to the instructors of the classes. 89

95 Boy Scouts Merit Badge University In addition to the program being developed for the middle school, the team is also participating in the Merit Badge University program being held at Auburn University, teaching the space exploration merit badge. This involves teaching the following criteria, taken from the Boy Scouts Requirement Handbook: 1) Tell the purpose of space exploration and include the following: a. Historical reasons b. Immediate goals in terms of specific knowledge c. Benefits related to Earth resources, technology, and new products. d. International relations and cooperation 2) Design a collector's card, with a picture on the front and information on the back, about your favorite space pioneer. Share your card and discuss four other space pioneers with your counselor. 3) Build, launch, and recover a model rocket.* Make a second launch to accomplish a specific objective. (Rocket must be built to meet the safety code of the National Association of Rocketry. See the "Model Rocketry" chapter of the Space Exploration merit badge pamphlet.) Identify and explain the following rocket parts: a. Body tube b. Engine mount c. Fins d. Igniter e. Launch lug f. Nose cone g. Payload 90

96 h. Recovery system i. Rocket engine 4) Discuss and demonstrate each of the following: a. The law of action reaction. b. How rocket engines work c. How satellites stay in orbit d. How satellite pictures of Earth and pictures of other planets are made and transmitted. 5) Do TWO of the following: a. Discuss with your counselor a robotic space exploration mission and a historic crewed mission. Tell about each mission's major discoveries, its importance, and what was learned from it about the planets, moons, or regions of space explored. b. Using magazine photographs, news clippings, and electronic articles (such as from the Internet), make a scrapbook about a current planetary mission. c. Design a robotic mission to another planet or moon that will return samples of its surface to Earth. Name the planet or moon your spacecraft will visit. Show how your design will cope with the conditions of the planet's or moon's environment. 6) Describe the purpose and operation of ONE of the following: a. Space shuttle or any other crewed orbital vehicle, whether government owned (U.S. or foreign) or commercial b. International Space Station 7) Design an inhabited base located within our solar system, such as Titan, asteroids, or other locations that humans might want to explore in person. Make drawings or a model of your base. In your design, consider and plan for the following: a. Source of energy 91

97 b. How it will be constructed c. Life support system d. Purpose and function 8)Discuss with your counselor two possible careers in space exploration that interest you. Find out the qualifications, education, and preparation required and discuss the major responsibilities of those positions. 92

98 Section 6: Conclusion At the end of the design phase, the Auburn University USLI team is extremely confident in the design of Project Nova. The time between PDR and CDR has resulted in a further sophistication of the design, to the point where construction of the rocket can commence. At the end of the design phase, the team is confident that the manufacturing process will unfold on time and on budget, and will provide a successful mission in May. Given the significant challenges already faced, the team is confident that any challenges encountered in the manufacturing process can be overcome, and then deliver a highly successful Flight Readiness Review. 93

99 Appendix A: Educational Engagement School Handbook 77

100 S AMUEL G INN C OLLEGE OF E NGINEERING 2014 Student Outreach Program Official Handbook

101 Section 1: Auburn University Rocketry Association 1.1 Statement of Purpose The Auburn University Rocketry Association (AURA), along with the Department of Aerospace Engineering at Auburn University, is entering an exciting new era of growth, influence and leadership, as a devotion for the future advancement of aeronautical and astronautical engineering swells throughout the department. AURA will be participating in NASA s University Student Launch Initiative (USLI), which is a national design build launch rocketry competition. Also, the USLI competition requires teams to plan and execute educational engagement activities as a component of the overall project. However, AURA does not seek to fulfill the requirement for the sake of the competition. Just as NASA and the USLI competition has instilled the spirit of rocketry in AURA s team members, AURA truly aspires to instill the spirit of science, technology, engineering, mathematics and rocketry in young students here on the Plains. The solutions to the world s problems lie in the minds of generations to come. Through this program, the team hopes to inspire younger students in the same ways that AURA is inspired every day. 1.2 Program Leaders Outreach Program Team Lead: AJ Pollard, Undergraduate Student Contact Information: (256) , dap0018@auburn.edu Team Captain: Hugh (Luke) Humphreys, Undergraduate Student Contact Information: (904) , hnh0004@auburn.edu

102 Faculty Advisor: Dr. Joe Majdalani, Department Chair Contact Information: (334) , Faculty Advisor: Dr. Roy Hartfield, Professor Contact Information: (334) , Outreach Advisor: Rob Kulick, M. Ed., Student Services Contact Information: (334) , Safety Officer: Jacob Herrera, Undergraduate Student Contact Information: (256) , 1.3 Program Timeline Official Dates to come! Early March: Faculty Rocketry Workshop 96

103 Late March: 8 th Grade Rocketry Unit for one week Early April: 7 th Grade Rocket week Launch Day: Last day of 7 th Grade Rocket week Mid April: Feedback inventories due 97

104 Section 2: Program Overviews facilities Auburn in order University to design has and agreed build our to let rocket. AURA One use such several facility of the is engineering The Davis 2.1 Faculty Rocketry Workshop A one hour rocketry workshop will be held for all teachers and administrators who will be participating in the implementation of the 7 th Grade Rocket Week program in early April. During this time, AURA will outline the Rocket Week program covered in this handbook. The learning objectives and lesson plans will be discussed at the beginning of the session, and as time permits, the teachers and administrators present will begin constructing an Alpha rocket under the guidance of AURA. At the end of the program overview, attendees will be able to express concerns and ask questions until the end of the workshop. If necessary and agreed to, a follow up workshop may be held soon after to continue constructing rockets and to continue answering questions. 2.2 AJHS 8 th Grade Rocketry Unit The 8 th Grade Rocketry Unit is designed for a small group of 8 th grade students at Auburn Junior High School. Throughout this unit, the students will learn some basic rocketry terms and ideas that relate to Newton s 3 rd Law of Motion. Additionally, students will have the opportunity to build and launch model Alpha rockets under the guidance of AURA team members and teachers. During the assembly process, there will be open discussion for students to ask any questions or discuss any topic with AURA team members regarding rocketry, aerospace engineering, or any other STEM related topics in general. The purpose of this unit is to enrich and interest students in science, math, technology and engineering related fields in the hopes that they might consider a future in one of those careers. This unit will also provide these students with a fun activity that they may not have the opportunity to take advantage of every day. This hands on approach in teaching young students about rockets could perhaps inspire some to pursue a fulfilling journey in science, math, technology and engineering. 2.3 DMS 7 th Grade Rocket Week The 7 th Grade Rocket Week will be held at Drake Middle School during (dates TBA) for the entire 7 th grade class. This program is designed to enrich the life science curriculum with 98

105 an exciting rocketry focus. More specifically, the 7 th grade class will be learning about G forces that the human body experiences at high accelerations and microgravity environments. Students will learn about what can happen to the human body when it is exposed to prolonged high G force operations. After the lesson, an AURA team member will lead the class in the construction of a small model rocket. The students will have been placed on previously determined teams organized by the teachers, and each student will have a position on the team. Each position on the team will have certain assignments pertaining to the construction of the model rocket. Students will collaborate together to complete the project in a timely and organized manner. Model rocket construction should only take about 1.5 class periods, beginning immediately after the lesson and ending by the end of class the day after. On the third day, the science classes will meet in the field located on Drake Middle Schools campus during regular class time for the rocket launch. In the event that weather disrupts the scheduled launch day or that a significant number of rockets were incomplete before launch, a fourth day may be added to make up for the launch delays. 2.4 Outreach Feedback Inventory Will be included in print version of handbook. 99

106 Section 3: Prerequisite Faculty Workshop 3.1 Faculty Workshop Outline Will be included in next draft 100

107 Section 4: AJHS 8 th Grade Rocketry Unit 4.1 Introduction The 8 th Grade Rocketry Unit is designed for a small group of 8 th grade students at Auburn Junior High School. Throughout this unit, the students will learn some basic rocketry terms and ideas that relate to Newton s 3 rd Law of Motion. Additionally, students will have the opportunity to build and launch model Alpha rockets under the guidance of AURA team members and teachers. During the assembly process, there will be open discussion for students to ask any questions or discuss any topic with AURA team members regarding rocketry, aerospace engineering, or any other STEM related topics in general. The purpose of this unit is to enrich and interest students in science, math, technology and engineering related fields in the hopes that they might consider a future in one of those careers. This unit will also provide these students with a fun activity that they may not have the opportunity to take advantage of every day. This hands on approach in teaching young students about rockets could perhaps inspire some to pursue a fulfilling journey in science, math, technology and engineering. 101

108 4.2 8 th Grade Lesson Plan 8 th Grade Rocketry Lesson, Provided by EstesEducator.com Learning About Motion and Flight with a Model Rocket Objective of the Lesson: The student will be able to: Identify and trace the basic path of a rocket from launch to recovery. Describe how Newton s Third Law of Motion relates to launching a model rocket. Recognize and define vocabulary. Pass out vocabulary handouts for students to complete throughout the lecture and turn in at the end of the week. BACKGROUND FOR THE TEACHER Thrust is the upward force that makes the rocket accelerate upward. This is a demonstration of Newton s Third Law of Motion: For every action there is an equal and opposite reaction. The action is the gas escaping through the nozzle. The reaction is the rocket accelerating upward. The rocket will continue to accelerate until all of the propellant in the rocket engine is used up. 102

109 The casing of a model rocket engine houses the propellant. At the base of the engine is the nozzle, a heat resistant, rigid material. The igniter in the rocket engine nozzle is heated by an electric current supplied by a battery powered launch controller. The hot igniter ignites the solid rocket propellant inside the engine which produces gas while it is being consumed. This gas causes pressure inside the rocket engine, which must escape through the nozzle. The gas escapes at a high speed. This produces thrust. Above the propellant is the smoke tracking and delay element. Once the propellant is used up, the engine s time delay is activated. The engine s time delay produces a visible smoke trail used in tracking, but no thrust. The fast moving rocket now begins to decelerate (slow down) as it coasts upward toward apogee (peak altitude). The rocket slows down due to the pull of gravity and drag. Drag is the force that resists the forward motion of an object through the air. When the rocket has slowed enough, it will stop going up and begin to arc over and head downward. This high point is the apogee. At this point the engine s time delay is used up and the ejection charge is activated. The ejection charge is above the delay element. It produces hot gases that expand and blow away the cap at the top of the engine. The ejection charge generates a large volume of gas that expands forward and pushes the parachute out of the top of the rocket. The parachute now opens and provides a slow, gentle and safe landing. The rocket can now be prepared to launch again! 103

110 VOCABULARY Accelerate: Speed up. Gravity: The force that pulls all objects to the center of the Earth. Apogee: The peak altitude a rocket reaches when it is farthest from the surface of the earth. 104

111 Igniter: An electrical device that ignites the combustion of the propellant in a model rocket engine. Decelerate: Slow down. Launch: The lift off of a model rocket following the ignition of the engine. Delay Element: Ignites after the propellant burns out and is an aid in tracking the rocket and in providing a time delay during which the rocket coasts to apogee. Propellant: A mixture of fuel and an oxidizer which is the source of motive energy in a rocket. Drag: The force that resists the forward motion of an object as it moves through the air. Recovery System: The device in a model rocket whose purpose is to return the rocket to the ground safely by creating excess drag or by creating lift. Ejection Charge: Ignited by the delay element and produces expanding gases which activate or eject a recovery device. Thrust: The force that makes the rocket accelerate upward as the propellant is burning. STRATEGY 105

112 MOTIVATION: Show the students the rocket you have constructed. (It is essential that the teacher construct the specific rocket before beginning this unit.) Ask: What is this object? How does it work? (Allow the students to discuss how rockets are used in space specifically and other ideas they may have, such as rockets used to launch missiles and for launching fireworks.) Using a blank overhead transparency, begin to put the outline of the events of a model rocket launch in order as the students contribute ideas. Begin with the launch and end with the recovery. Then display the overhead transparency (1: Flight Sequence of a Model Rocket provided in the back of the handbook). Use both transparencies to demonstrate the areas that need clarification. Label the appropriate parts of the engine as you describe the flight sequence. Closure: Review with the students the concepts of thrust, launch, apogee, delay element, ejection charge, drag and recovery by asking them to read the definitions from their vocabulary sheet. Transition the lesson to the AURA team members in the classroom in order to begin constructing the rockets. 106

113 4.3 Model Rocket Construction Each student will be given an Alpha model rocket to assemble collaboratively under the supervision and guidance of AURA team members and the teacher. The students will be supplied with all materials needed in order to construct the rockets. Construction should take about 2 3 class periods since the class time is shorter for these students. While the students assemble their rockets, AURA team members will be talking with them about the importance of working together on teams and discussing science, technology, engineering and mathematical topics with the students that spark the students interests. AURA team members will do their best to answer any questions that the students may have regarding rocketry, aerospace 107

114 engineering, or STEM related subjects in general. As the students draw near to completing the construction of the rockets, launch safety procedures will be covered, and the students will sign a safety acknowledgement form. Finally, launch day procedures will be covered with the students in order to prepare them for the last day of the unit th Grade Launch Day Since there is a small number of students participating in this launch, each student will have the opportunity to launch and recover his or her own rocket in the football field located on Auburn Junior High School s campus. Before the launch, AURA team members will collect each rocket and install the motors in all of the rockets. Each rocket will be placed one at a time on the launch pad and connected to the electrical igniter system. When all is clear for the launch, an AURA team member will disable the safety mechanism on the launch switch. A five second countdown will begin out loud, and the prepared student will launch his or her rocket at the end of the countdown. An AURA team member will enable the safety mechanism on the launch switch once again after the rocket has been launched, and when given permission, the student will retrieve his or her rocket to keep. 108

115 Section 5: DMS 7 th Grade Rocket Week 5.1 Introduction Will be included in next draft 5.2 Daily Schedule Day 1: 1. Present G forces lecture (20 25 minutes). 2. Distribute and explain homework assignment. 3. Begin model rocket construction. Work until 5 minutes before end of class. 4. Have students clean their workspaces and store rockets and supplies in appropriate areas. 5. Prepare for next class. Day 2: 1. Turn in homework. 2. Continue working on model rockets until completion. Also, discuss safety guidelines and have students sign the Safety Procedures Acknowledgement Form. 3. Discuss launch day procedures and assign launching orders minutes before the end of class, have students clean their areas and store rockets and supplies in appropriate areas. Day 3: 1. After roll call for each class period, teachers and AURA members lead students to the launch site on campus. All rockets will have already been moved to the launch site. All science classes during each period will meet at the launch site for each of the seven class periods. 2. Each team in each class, along with each teacher and staff, will have the opportunity to launch his or her rocket only when cued by the appropriate AURA team member. One student per team will have the 109

116 responsibility of pressing the electronic ignition switch at the end of a countdown. 3. When all launches are complete, students will return to the classrooms to conclude the program and allow for the next group of classes to launch. Day 4: MAKEUP DAY th Grade Lesson Plan 7 th Grade G forces and Microgravity Lesson Some material provided by for kids.com Learning About the Effects of Gravity on the Human Body in Space Flight Objectives of the Lesson: The student will be able to: Explain what g forces are and how high g forces affect the human body Understand the behaviors of the heart and circulatory system of the human body in microgravity A G force is essentially a force with an acceleration that is equivalent to that of Earth s gravity. In a more technical sense, a G force is a numerical ratio that relates the weight of an object or body on Earth to some multiple of gravitational exerted on that object or body. For example, imagine that you are in an elevator. Have you ever noticed that for a split second when the elevator starts moving upward, you feel like you are being pushed into the floor, feeling heavier? Or for a split second when it starts moving downward, it feels like the floor is dropping out from underneath you? In a more extreme example, imagine that you are on a roller coaster. When you hit the bottom of a big hill, you feel like you are being pushed down in your seat, and at the top of a hill, you feel like you are coming out of your seat. These are just two examples of what it is like to experience g forces. When you feel like you are being pressed down or that you are heavier, those are called positive g forces, and when you feel like you are weightless or coming out of your 110

117 seat, those are negative g forces. G forces are expressed numerically in multiples of the force of gravity. When you are not changing speeds or direction (accelerating), you only experience 1 G, the force of gravity. Any forces greater than the force of gravity would be higher number Gs, and vice versa. For example, if you weigh 150lb, and you experienced an acceleration that was two times greater than gravity, you would be experiencing 2 Gs. In this case, with 2 Gs of force, it would feel like you weighed 300lb. During the ascension of a rocket into space, astronauts are pressed into their seats with more than three times the force of gravity, or more than 3 Gs. Crew members on the Russian Soyuz rocket endure four times the force of gravity. The Mercury capsules launched by America s Atlas booster rocket reached a peak acceleration of 8 Gs during ascent to orbit, then decelerated during re entry to the Earth at loads as high as 7.8 Gs. America s Titan rockets launched the Gemini space flights at 7.25 Gs, and NASA s Saturn 5 rocket, the largest rocket ever built, peaked at 4 Gs. However, the Apollo rocket capsules returning from the Moon re entered the atmosphere at over 6 Gs. What do you think that would feel like? How do you think the human body would react to the sudden acceleration? Present video clip of G force training. The ascent phase of a rocket launch into orbit is rather. The space shuttle typically took about 8 minutes to reach orbit and travel at a speed of 17,000 mph. A 200lb astronaut would feel like he or she weighed close to 800lb for 8 minutes on the trip to space. Prolonged high G forces affect the circulatory system and our internal organs. As G forces increases, visual effects include loss of color vision, followed by tunnel vision (where peripheral vision is lost, retaining only the center vision). If G forces increase further, complete loss of vision will occur, while consciousness remains. These effects are due to a reduction of blood flow to the eyes before blood flow to the brain is lost, because the extra pressure within the eye counters the blood pressure. This is what we saw happening to the pilots in training in the video. A further increase in g forces will cause G LOC where consciousness is lost. It is also important the direction of the high G force; the human body is considerably more 111

118 able to survive G forces that are perpendicular to the spine. That is why astronauts lie on their backs during launches rather than sitting or standing upright. Now, the ascent phase of our rocket launch and the deployment of the capsule are shorter than a giant rocket launch, but it will still create significant pressure on the rocket and would put plenty of G force on anything small enough to ride in it. Then, the lecture will switch to microgravity, or 0 G forces. This is why astronauts are weightless in space. After the 8 minute ascent to orbit, astronauts get to enjoy flying around inside the vehicle. However, despite what many people think, there are also many health risks involved in microgravity space flight. Next, the students will watch a video from NASA about what happens to the human heart and circulatory system when subjected to long periods of microgravity. The video will also give a little insight to how astronauts mitigate those negative health risks. Real World: Heart Rate and Blood Pressure s Conclude lesson and hand out the homework worksheet. Transition the class to an AURA team member to begin model rocket assembly with the class. 5.4 Model Rocket Construction th Grade Launch Day 112

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120 Section 6: Outreach Feedback Inventory 6.1 Feedback Instructions Will be included in next draft 114

121 Section 7: Model Rocket Safety Code 7.1 NAR Model Rocket Safety Code From the National Association of Rocketry, the following safety code is to be implemented in all aspects of the Auburn rocketry outreach program. Effective August Materials. I will use only lightweight, non metal parts for the nose, body, and fins of my rocket. 2. Motors. I will use only certified, commercially made model rocket motors, and will not tamper with these motors or use them for any purposes except those recommended by the manufacturer. 3. Ignition System. I will launch my rockets with an electrical launch system and electrical motor igniters. My launch system will have a safety interlock in series with the launch switch, and will use a launch switch that returns to the "off" position when released. 4. Misfires. If my rocket does not launch when I press the button of my electrical launch system, I will remove the launcher's safety interlock or disconnect its battery, and will wait 60 seconds after the last launch attempt before allowing anyone to approach the rocket. 5. Launch Safety. I will use a countdown before launch, and will ensure that everyone is paying attention and is a safe distance of at least 15 feet away when I launch rockets with D motors or smaller, and 30 feet when I launch larger rockets. If I am uncertain about the safety or stability of an untested rocket, I will check the stability before flight and will fly it only after warning spectators and clearing them away to a safe distance. When conducting a simultaneous launch of more than ten rockets I will observe a safe distance of 1.5 times the maximum expected altitude of any launched rocket. 6. Launcher. I will launch my rocket from a launch rod, tower, or rail that is pointed to within 30 degrees of the vertical to ensure that the rocket flies nearly straight up, and I will use a blast deflector to prevent the motor's exhaust from hitting the ground. To prevent accidental eye injury, I will place launchers so that the end of the launch rod is above eye level or will cap the end of the rod when it is not in use. 115

122 7. Size. My model rocket will not weigh more than 1,500 grams (53 ounces) at liftoff and will not contain more than 125 grams (4.4 ounces) of propellant or 320 N sec (71.9 pound seconds) of total impulse. 8. Flight Safety. I will not launch my rocket at targets, into clouds, or near airplanes, and will not put any flammable or explosive payload in my rocket. 9. Launch Site. I will launch my rocket outdoors, in an open area at least as large as shown in the accompanying table, and in safe weather conditions with wind speeds no greater than 20 miles per hour. I will ensure that there is no dry grass close to the launch pad, and that the launch site does not present risk of grass fires. 10. Recovery System. I will use a recovery system such as a streamer or parachute in my rocket so that it returns safely and undamaged and can be flown again, and I will use only flame resistant or fireproof recovery system wadding in my rocket. 11. Recovery Safety. I will not attempt to recover my rocket from power lines, tall trees, or other dangerous places. LAUNCH SITE DIMENSIONS Installed Total Impulse (Nsec) Equivalent Motor Type Minimum Site Dimensions (ft.) /4A, 1/2A A B C D E 1, F 1,

123 G 1, Two Gs 1, Additional Safety Points 1. No students will be allowed to handle rocket motors. All motors will be stored in a safe place until launch day. Only AURA team members will install the rocket motors. 2. No students will be allowed to use cutting knives needed for model rocket assembly. All steps involving the cutting knives may only be done by AURA team members or teachers. 117

124 Section 8: Forms and Documents facilities Auburn in order University to design has and agreed build our to let rocket. AURA One use such several facility of the is engineering The Davis th Grade Lesson Plan Materials The following two pages feature the transparency needed for the lesson and the vocabulary worksheet for the students to complete. 118

125 FLIGHT SEQUENCE OF A MODEL ROCKET 119

126 Name WORDS FOR ROCKETEERS Directions: As you learn these words during each session about rockets, you can fill in the definition. If you need more information, you can also use a dictionary. ROCKETS IN MOTION 1. ACCELERATE 2. APOGEE 3. DECELERATE 4. DELAY ELEMENT 5. DRAG 6. EJECTION CHARGE 7. GRAVITY 8. IGNITER 9. LAUNCH 10. PROPELLANT 11. RECOVERY SYSTEM 120

127 12. THRUST th Grade Lesson Plan Materials 8.3 School Agreement Letter Will be included in print version of handbook. 8.4 Teacher Agreement Letter Will be included in print version of handbook. 8.5 Safety Procedures Acknowledgement Form Will be included in print version of handbook. 8.6 Feedback Questionnaire Will be included in print version of handbook. 8.7 Alpha Rocket Assembly Manual Will be included in print version of handbook. 121

128 Appendix B: Sampling of MSDS 122

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