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

Transcription

1 University of Evansville Student Launch Enclosed: Preliminary Design Review Submitted by: Rocket Team Project Lead: David Eilken Submission Date: November 04, 2016 Payload: Fragile Material Protection Mentor: Dr. David Unger, NAR 89083SR Level 2 Submitted to: NASA Student Launch Initiative Program Officials Faculty of the UE Mechanical Engineering Program University of Evansville College of Engineering and Computer Science 1800 Lincoln Avenue; Evansville, Indiana i P a g e

2 Table of Contents Table of Contents... ii List of Figures... iv List of Tables... vi PDR Summary... 1 Design Updates from Proposal... 2 Changes Made to Vehicle Criteria... 2 Changes Made to Payload Criteria... 2 Changes Made to Project Plan... 3 Vehicle Criteria... 4 Selection, Design, & Rationale of Launch Vehicle... 4 Mission statement... 4 Mission Success Criteria... 4 System Level Alternatives and Analysis... 6 Component Alternatives Motor Alternatives Recovery Payload Electronic Payload Fragile Material Payload ii P a g e

3 Mission Performance Predictions Safety Overview Final Assembly Checklist Launch Procedures Checklist Personnel Hazard Analysis Failure Modes and Effects Analysis Environmental Considerations General Risk Assessment Project Plan Requirements Compliance Budget Schedule References Appendix A Machine Prints Appendix B OpenRocket Simulation Appendix C Parts List Appendix D Task Breakdown iii P a g e

4 List of Figures Figure 1 - Updated 3D Model of Launch Vehicle... 2 Figure 2 - Rocket System Decomposition... 6 Figure 3 - Weight breakdown (all weights are in lbf)... 7 Figure 4 - Dimensioned drawing of full body (all dimensions in inches)... 8 Figure 5 - Subsection dimensions... 8 Figure 6 - Nosecone mounting diagram... 9 Figure 7 - Exploded View of the Motor Mount Figure 8 - Propulsion Components Labeled Figure 9 - Dimensional Drawing for the Motor Mount Figure 10 - PerfectFlite Stratologger CF Altimeter Figure 11 - Block diagram of major recovery system electrical components Figure 12 - Recovery bay bulkheads and hardware Figure 13 Exploded View; Recovery System Figure 14 - Recovery system layout within airframe Figure 15 Tethering of Rocket Sections Figure 16 - Electronic Payload within Nosecone Figure 17 - Exploded View of Electronic Payload Figure 18 - Exploded Electronic Payload View with Nosecone Figure 19 - Top View, Assembled Electronic Payload Figure 20 - Bottom View, Assembled Electronic Payload Figure 21 - Payload Exploded View Figure 22 - Components of the Main Payload iv P a g e

5 Figure 23 Payload Inner Cylinder Figure 24 - Springs Attached to the Main Payload and Bulkhead Attachment Figure 25 - System Drawing and Force Balance Figure 26 - Free Body Diagram for the Entire Rocket and Force Balance Figure 27 - Free Body Diagram for the Spring Assembly and Force Balance Figure 28 - Free Body Diagram for Cylinder 1 and the Fragile Material and Force Balance 42 Figure 29 Simulink Mathematical Model Figure 30 - Predicted Altitude from OpenRocket Simulation Figure 31 - OpenRocket Flight Simulation Inputs Figure 32 - Predicted Altitude from Rocksim Simulation Figure 33 - Inputs for Rocksim Simulation Figure 34 - Thrust Curve from AeroTech Motor Figure 35 - Thrust Curve for the L850W Motor in OpenRocket Figure 36 - Thrust Curve for the L850W Motor in Rocksim Figure 37 - Center of pressure and gravity Figure 38 - Gantt Chart v P a g e

6 List of Tables Table 1 - Decision Matrix Key Table 2 - Decision Matrix: Body Tube Table 3 - Decision Matrix: Fin and Nosecone Material Table 4 - Decision Matrix: Bulkhead Material Table 5 - Decision Matrix: Fin Shape Table 6 - Decision Matrix: Nosecone Shape Table 7 - Decision Matrix: Motor Mount Design Table 8 - Decision Matrix: Centering Rings Table 9 - Decision Matrix: Recovery Altimeter Table 10 - Decision Matrix: Recovery Harness Material Table 11 - Decision Matrix: Drogue Parachute Table 12 - Decision Matrix: Main Parachute Table 13 Motor Considerations and Specifications Table 14 - Testing Matrix for Fragile Material Table 15 - Force Events for the Simulink Model Table 16 - Final Values for Constants Table 17 - Kinetic energy of each section upon landing Table 18 - Landing site distance from launch site by wind speed Table 19 - Personnel Hazard Analysis Table 20 - Failure Modes and Effects Analysis Table 21 - Environmental Consideration Analysis Table 22 - General Risks Associated with the Project vi P a g e

7 Table 23 - Requirement Compliance Table 24 - Team Requirements Table 25 - Section Level Budget Table 26 - Funding Sources Table 27 - Critical Dates vii P a g e

8 PDR Summary Project ACE plans to field a 111 long, 35-pound carbon fiber and aluminum based rocket. The leading tip of the rocket begins with a G-10 Fiberglass, 22, ogive nosecone. Contained in a waterproof compartment in the nosecone sits the official altimeter as well as a GPS tracking system. Just aft of this compartment are four threaded rods for fastening ballast. A fragile material protection system resides below the nosecone. This payload contains concentric cylinders, connected by an array of springs and wire-rope isolators selected through extensive mathematical modeling. The innermost cylinder, where the fragile material will be contained, will feature variable position cap and fill material to ensure that the fragile material will be contained under sufficient pressure regardless of volume. It is the team s objective to produce a successful payload that provides meaningful vibration and impulse reduction information. Moving down the rocket from the payload is the recovery system. This system features completely redundant separation circuits. At apogee, a 48 drogue chute will eject, followed by a 96 main chute closer to ground level. At the aft end of the rocket sits the propulsion section. A 75-mm L-850W Aerotech motor will propel the rocket for just over four seconds. This motor will be held in place via 6061-T6 Aluminum centering rings and thrust plates. All components will be housed in two carbon fiber body tubes. The fins, which adhere to the centering rings and body tubes, will be made out of G-10 Fiberglass and have a clipped delta design. Each system is covered in much more depth in the Vehicle Criteria section of this report. For specific team information, such as the mentor and mailing address, please see the cover page of this report. For more quick facts on the rocket please reference the associated milestone review flysheet. 1 P a g e

9 Design Updates from Proposal Changes Made to Vehicle Criteria The bow body tube was elongated by 8 to accommodate the design changes made to the main payload. The fin thickness was also decreased to and designed to have a beveled leading edge. This will decrease the drag on the launch vehicle. Lastly, it was determined through manufacturer specifications that the exact length of the nosecone will be The remainder of the vehicle criteria remained unchanged. An updated 3D model of the launch vehicle can be seen in Figure 1. Main Payload Figure 1 - Updated 3D Model of Launch Vehicle Changes Made to Payload Criteria The spring system used to support the payload added 5 base springs after the math model proved that wire rope isolators alone would not be sufficient. To accommodate this design change, the entire previous payload was re-designed to oscillate within the body tube. The spring selections originally planned also changed due to system optimization through a math model. More detail can be found in the payload section. 2 P a g e

10 Changes Made to Project Plan Few changes have occurred to the project plan since the proposal was submitted. NASA SLI officials have indicated that the due date for PDR documentation has been moved to November 4 th, 2016 (originally October 28 th ). Despite this, Project ACE has decided to keep to the schedule of having PDR documents completed by October 26 th. This will enable the team to focus on the build phase of the sub-scale rocket. More on the schedule can be found in the Schedule section of this report. The budget has been decreased by $ Additionally, funding has been allocated in a slightly different fashion than in the proposal. The reason for this is twofold: first, the motor had unforeseen hardware costs associated with it, increasing the funds needed for that section. The travel and lodging portion of the budget decreased substantially, as Project ACE decided not to have the team cover any meal costs. Also, it was determined that advisor expenses would come out of the University of Evansville College of Engineering and Computer Science budget instead of the project budget. A detailed budget breakdown can be found in the Budget section of this report. 3 P a g e

11 Vehicle Criteria Selection, Design, & Rationale of Launch Vehicle Mission statement Project ACE is an interdisciplinary university project with the united goal of constructing and flying a high powered aircraft with a unique experimental payload. Our team intends to perform at a high level at the national competition and pass down the knowledge gained from this experience to current underclassmen and future Project ACE members. Mission Success Criteria 1. Aerodynamics a. The airframe, nose cone, and fins should remain intact for the duration of the flight. b. The airframe, nose cone, and fins should be reusable for any following flights. c. The airframe and nose cone should protect all internal components from damage from external sources. 2. Propulsion a. The vehicle should attain an apogee between 5,125 feet and 5,375 feet. b. The vehicle should remain below Mach 1. c. The motor mount should withstand propulsion forces and remain reusable for any following flights. 3. Recovery a. The drogue parachute and main parachute are ejected at apogee and 1000 feet, respectively. 4 P a g e

12 b. The drogue parachute and main parachute inflate successfully following ejection. c. The maximum kinetic energy of any independent section of the rocket is less than 75 ft-lbf at landing. 4. Electronic Payload a. The data sent from the electronic payload should be able to be received remotely during and after the vehicle s flight. b. The electronic payload should withstand flight forces and remain reusable for any following flights. c. The electronic payload should accurately determine the apogee of the rocket. 5. Main Payload a. The fragile object(s) should remain undamaged. b. The force felt by the payload should be reduced by 50% for each of the areas of interest: takeoff (thrust curve, parachute deployment, and landing.) c. The force felt by the payload should be reduced by 35% for each of the areas of interest: (thrust curve, parachute deployment, and landing.) 5 P a g e

13 System Level Alternatives and Analysis The launch vehicle was designed with 5 interconnected systems: the airframe, electronic payload, main payload, recovery, and propulsion. These systems and relationships can be seen in Figure 2. The airframe is the parent system and houses all the sub-sections. Figure 2 - Rocket System Decomposition The full weight of the launch vehicle is lbf. A weight breakdown of the rocket and the individual subsections can be seen Figure 3. 6 P a g e

14 Figure 3 - Weight breakdown (all weights are in lbf) The purpose of the airframe is to provide a structure for the internal systems and protect them from external stresses. The airframe was designed to be comprised of two carbon fiber body tubes and an ogive fiberglass nosecone. The body tubes will be made of carbon fiber. Both body tubes will have a diameter of 5.5. The aft body tube will have a length of 48. The bow body tube will have a length of 41. The nosecone will be made of fiberglass and will have a 4:1 ogive profile. The total length of the nosecone is A dimensioned drawing of the full body is provided in Figure 4. 7 P a g e

15 Figure 4 - Dimensioned drawing of full body (all dimensions in inches) The airframe will house 4 main systems: electronic payload, main payload, recovery, and propulsion. The allocated space and sizing for the individual subsections can be seen below in Figure 5. Figure 5 - Subsection dimensions The two body tube system was chosen over a single body tube system. This was done in order to incorporate a dual deployment recovery system that would separate between the two body tubes. The retention system for the nosecone is currently designed to be mounted with 3 bolts and 3 adhesive mount nuts. This was chosen over alternatives such as a threaded rod mounted down the length of the nosecone or threads on the interior wall of the bow body tube. The current system was chosen because it allows the bow body tube to remain completely free of permanent mounting hardware. This allows the main payload to be removed and inserted with ease. This design can be seen below in Figure 6. 8 P a g e

16 Figure 6 - Nosecone mounting diagram The purpose of the electronic payload is to provide an official altitude, GPS coordinates for the launch vehicle, and hold ballasts. It will be mounted with a gasket and removable mount combination in the nosecone with the electronics facing towards the bow end of the rocket. This will provide an added measure of security towards water damage. The Atlus TeleMega was chosen against other altimeters because it records flight data in addition to apogee and GPS location. Much of this data can be compared to RockSim. Other altimeters were cheaper, however, the extra data (such as rocket tilt) was determined to be worth the cost difference. The purpose of the main payload is to protect fragile materials. It consists of a concentric cylinder design as well as a series and parallel spring system. The inner cylinder utilizes wire rope isolators to absorb smaller vibrations while larger springs at the base of the cylinder reduce the force of large impulse impacts such as takeoff, landing, and main parachute deployment. 9 P a g e

17 Prior to choosing the main payload design that currently exists, several options were discussed. One option was simply a payload bay with support material and a cap that had built in damping to hold the unknown fragile material object(s) in place and hopefully protect them. The other alternative was the concentric cylinder design with wire rope isolators; however, the math model used to predict the behavior of the system showed this was not sufficient. That is what prompted the additional larger springs that were added in series with the wire rope isolators. The recovery system serves to return the launch vehicle to the ground safely, minimizing the ground impact velocity to preserve the structural components of the rocket as well as the fragile payload. A dual-deployment system utilizing a 36" drogue parachute and a 96" main parachute has been designed to use identical, redundant electrical systems to trigger black powder ejection charges. The electrical systems will be housed in a coupling tube that unites the bow and aft body tubes. The drogue chute will be packed in the bow tube, and the main chute in the aft tube. All sections of the rocket will be tethered together using a tubular nylon recovery harness. Several system-level alternatives were considered for the recovery system. In particular, a gas ejection system was investigated, in which a canister of compressed CO2 is used to pressurize the parachute compartment during a deployment event. While gas ejection systems do not subject the parachute to the high temperatures of a black powder ejection, they tend to be heavier, more complicated, and more expensive than a simple black powder ejection. For these reasons, a gas ejection system was not selected. Additionally, different parachute deployment schemes were considered. In many rockets, the drogue parachute is packed underneath the nose cone and deployed by blowing the nose cone out 10 P a g e

18 of its body tube. This method was not selected because it would require that the recovery electronics be located in close proximity to the transmitting components of the competition altimeter, which could create unwanted interference. Recovering the rocket in multiple components was also considered; for example, the bow and aft body tubes could be tethered together after drogue deployment and split during the main deployment to be descended under separate main parachutes. This setup was not selected due to limitations on body tube space created by the main payload. Lastly, the aft body tube houses the propulsion section. The purpose of the propulsion section is to propel the launch vehicle to a height of 5,280 ft. The propulsion section was designed to house 3 centering rings and an engine block (all made of 0.25 aluminum). The aft body tube will be slotted to allow the fiberglass fins to be attached to the inner tube and centering rings. This adds further support for the fins and centering rings. The inner tube will be made of blue tube and have a 3.1 OD and 20 length. The inner tube will house an Aerotech L850W motor with a max thrust and impulse of 1185 newtons and 3695 newton seconds, respectively. The fins will have a clipped delta design. The propulsion system was designed around a few key criteria. First, it was decided to use 3 centering rings versus 2 centering rings. This decision was made to increase stability of the inner tube. With a 3-centering ring system, two centering rings can support the fin tabs and one centering ring can be used as a thrust plate and serve as a mounting point for the motor retention system. Secondly, two motor retention systems were evaluated. The first system included threaded rods mounted to the engine block. The second system mounts directly to the furthest aft centering ring. The second system was chosen because of the decreased complexity and decreased weight. 11 P a g e

19 Component Alternatives Decision matrices were used to visually and concisely evaluate multiple component-level options. These matrices can be seen throughout the report, and the key that they follow is located in Table 1. Bolded and underlined options indicate design selections. Discussion of the various decision matrices can be found immediately following each matrix. Table 1 - Decision Matrix Key Decision Matrix Criteria О Good Δ OK X No Good Table 2 - Decision Matrix: Body Tube Decision Matrix Body Tube Option Cost Strength Ductility Overview Decision Explanation Carbon Fiber X О О Fiberglass Δ Δ Δ Blue Tube О X X Carbon Fiber provides the highest tensile strength and lowest ductility at the highest cost Fiberglass provides a moderate strength, ductility, and cost relative to Carbon Fiber and Blue Tube. Blue Tube provides the lowest ductility and strength at the lowest cost. Material considerations for the airframe included fiberglass, carbon fiber, and Blue Tube. The team intends to use carbon fiber for the body tubes because it has a higher tensile strength, lower density, and a lower ductility compared to that of fiberglass or Blue Tube. Flexibility in a rocket airframe is an unwanted characteristic so a lower ductility is beneficial. In addition, the 12 P a g e

20 higher tensile strength of carbon fiber will ensure a higher allowable stress and a higher factor of safety than that of fiber glass. Table 3 - Decision Matrix: Fin and Nosecone Material Decision Matrix Fin and Nosecone Material Option Cost Strength Ductility Overview Decision Explanation Carbon Fiber X О О Fiberglass О Δ Δ ULTEM X О О Carbon Fiber provides a high tensile strength and low ductility at a high cost. Fiberglass provides a moderate strength, ductility, and costs significantly less than Carbon Fiber or ULTEM. ULTEM provides a high tensile strength and low ductility at a high cost. The material for the fins and nosecone will be G-10 fiberglass because it is commercially available at a low cost. Carbon fiber and ULTEM plastic are also materials used for fin design; however, these did not provide adequate benefit to mitigate the significantly higher cost. This is because the nosecone and fins are not being required to undergo the same stresses caused by recovery process as the body tubes, so the additional strength of carbon fiber is not sufficient for these components. 13 P a g e

21 Table 4 - Decision Matrix: Bulkhead Material Decision Matrix Bulkhead Material Option Cost Strength Weight Overview Decision Explanation Aluminum offers the highest strength of all materials considered. It comes at an increased cost and weight. Plywood offers the lowest cost and weight at the price of strength. Fiberglass offers a moderate alternative to plywood and fiberglass. Aluminum X О Δ Plywood O X O Fiberglass Δ Δ Δ The bulkheads will be made of aluminum. Aluminum will be used to ensure the recovery and propulsion sections have strong attachment points. Fiberglass and plywood are common choices for bulkheads because they are sturdy, lightweight materials. However, since the design of the rocket is for an L-class motor, weight is not a significant constraint for material selection. This allows the team to choose the material with the highest tensile strength (aluminum) over fiberglass or plywood. Option Stability Table 5 - Decision Matrix: Fin Shape Decision Matrix Fin Shape Ease of Manufacturing Likelihood of Damage Clipped Delta Δ О O Trapezoidal X Δ О Tapered Swept О Δ X Overview Decision Explanation The Clipped Delta is the easiest to manufacture and offers moderate stability and drag. The Trapezoidal offers the lowest drag but the least stability. The Tapered Swept offers the highest drag but the least stability. 14 P a g e

22 The clipped delta design will be used for the fins. This design was chosen over other possible design choices such as a trapezoidal or tapered swept design. The difference between these designs is the sweep angle. This angle affects the center of pressure (CP) and thus affects stability. The clipped delta design was chosen after OpenRocket simulations and research was done on the various design choices. The research and simulations found the benefit of a different sweep angle to be minimal (<0.1 calipers stability increase). Additionally, changing the sweep angle to increase the stability would move the trailing edge of the fins aft of the end of the rocket. This would require the weight of the rocket to sit on the fins and increase the likelihood of damage. Table 6 - Decision Matrix: Nosecone Shape Decision Matrix Nosecone Shape Option Cost Drag Overview Decision Explanation Ogive Δ О The Ogive nosecone is the most difficult to manufacture and thus the most expensive but offers the lowest drag. The Elliptical nosecone can be purchased at a moderate cost for a moderate drag. The Conical nosecone is the easiest to manufacture and thus the least expensive but offers the highest drag. Elliptical O Δ Conical O X Although the Ogive nosecone shape is the most difficult to manufacture, it offers the lowest drag of all nosecone profiles. For this reason, the nosecone will be purchased. With the components of the body for the initial design of rocket chosen, the motor was the next area of the design. The first design of the motor was to use a cluster motor featuring three lower level motors to power the rocket. The other design consideration was using a single large motor to power the rocket. 15 P a g e

23 Table 7 - Decision Matrix: Motor Mount Design Decision Matrix Motor Mount Design Option Cost Safety Cluster Motor Against Regulations Mount O Δ X Single Motor Mount X Δ O Overview Decision Explanation The cluster motor would cost less and reach the optimal altitude with minimal safety concerns. The single motor cost is high and creates concerns about safety and reaching altitude. The motor mount that Project ACE was originally going to use was for a cluster motor configuration. This was due to the low cost of low level motors compared to a single large motor. Also, the cluster motors provide the ability to mix and match motors. The safety and complexity of the cluster motor, however, were concerns. There exists a heightened chance of misfires with use of more than one motor. There is also a chance that one motor does not ignite with the initial light, but could light from the other motors which is a clear safety concern. Table 7 shows the decision matrix for the motor mount design. Originally, the single motor mount was the back-up plan. As previously mentioned cost was a major concern with the single, large motor design. From a first inspection, the cost for a single large motor was five times that of a cluster motor configuration. Additionally, few large motors were suitable to reach the one-mile mark. This, in turn, limited the design of the motor mount due to the lack of motor choices. The forces being produced with a single large motor may also be more concentrated within the mounting configuration, requiring more robust mounting. Table 7 shows the decisions for the motor mount. 16 P a g e

24 Initial designs for the motor mount were considered, before the 2017 handbook was posted for the USLI teams and utilized a cluster motor. Once the new handbook was released, the team learned that cluster motors had been disallowed. Thus, the team decided to go with the single motor and single motor mount for the propulsion for the rocket. The single motor mount design would use a larger motor and thus concerns arose about the shear forces being produced on the centering rings and the bulkhead. These concerns will be mitigated using FEA. Table 8 - Decision Matrix: Centering Rings Decision Matrix Centering Rings Option Cost Strength Weight Overview Decision Explanation Plywood O X О G10 Fiberglass Δ О О Aluminum Δ О X Plywood is great for weight and cost but the strength is a problem for large motors The cost of fiberglass is budgetable because of the high strength and the weight of the material Aluminum has a good cost associated with machining it in house with high strength. Only concern is the weight The structural integrity of centering rings was already under review when the initial motor mount design was decided. This was due to the shear forces that could be expected with high power rocket motors. Due to this, strength was the major criterion that was used to select centering rings material. Table 8 shows the decision matrix for the centering rings. Plywood was the first material considered because of its low cost and low weight. However, the strength of the material (primarily Tensile Strength) was deemed significantly more important than cost or weight. 17 P a g e

25 The next material that was considered was G10 fiberglass. Fiberglass has high strength and low weight. This was appealing as the forces could likely be handled by the material and the low weight aided in raising the rockets altitude. However, the cost of Fiberglass is significantly greater than that of plywood. The last material that was considered was 6061-T6 aluminum. This was researched due to the high strength and machinability of the material. The cost of the material is manageable, especially since all machining would be conducted by the team. The only problem with the aluminum is the weight. Weight was decided to be a minor factor. Thus, the material that was selected was the aluminum. As it turned out, the added weight of the aluminum helped with controlling the altitude and bringing the rocket down to a desirable apogee. Also with the strength of aluminum being so great, the risk of the material shearing is low. Several dual-deployment altimeters were considered for the recovery electronics system; the PerfectFlite Stratologger CF, the AltusMetrum EasyMini, and the Entacore AIM3. To select this component, cost was given priority, as two of the selected altimeter type would need to be purchased to create redundancy within the system. All altimeters considered had similar feature sets which were sufficient for the purposes of the rocket, as more complex data collection and transmitting functions will be handled by the competition altimeter in the nosecone. The PerfectFlite Stratologger CF was selected. The decision matrix for the altimeter can be seen in Table P a g e

26 Table 9 - Decision Matrix: Recovery Altimeter Decision Matrix Recovery Altimeter Option Cost Feature Set Power Draw Overview Decision Explanation PerfectFlite For a low cost, this altimeter Stratologger provides a full set of features with O O Δ a higher power draw. CF AltusMetrum EasyMini Δ Δ О Entacore AIM3 X O О For a medium cost, this altimeter provides a reduced feature set with a low power draw. For a high cost, this altimeter provides a full set of features with a low power draw. Three materials are common when choosing a recovery harness for high-powered rockets: elastic, kevlar, and nylon. As this is a critical component, cost was not considered to be a high priority in the decision-making process. In order to reduce the maximum forces experienced by the rocket, a material with moderate elasticity was sought high elasticity in the recovery harness can cause the tethered components to snap back and collide with one another. The large forces involved with parachute deployment require a material with a high breaking strength. An elastic recovery harness would not be an acceptable selection due to its low strength and high elasticity. While Kevlar is incredibly strong, it has almost no elastic potential, which would do little to reduce the forces experienced by the rocket. Nylon was selected because it maintains a moderate degree of elasticity with a breaking strength well above the maximum force experienced by the rocket. Table 10 shows a decision matrix for the recovery harness material selection. 19 P a g e

27 Table 10 - Decision Matrix: Recovery Harness Material Decision Matrix Recovery Harness Material Option Cost Strength Elasticity Overview Decision Explanation Elastic O X X For a very low cost, elastic provides a low-strength, highelasticity solution. For a high cost, Kevlar provides the greatest strength with a very low elasticity. Tubular For a medium cost, nylon provides acceptable strength at a Nylon Δ Δ О medium elasticity. Tubular Kevlar X O Δ After investigating many parachutes from multiple manufacturers, the field was narrowed to focus on three different diameter Fruity Chutes parachutes for each the drogue and the main. Fruity Chutes was selected as a manufacturer based on a reputation for tough, well-made parachutes, as well as the small packing volume of their parachutes relative to their competitors products. In the selection of both parachutes, cost was deemed to be of minor importance due to the critical nature of the recovery system. Drogue parachute selection focused primarily on ensuring that the initial descent rate is low enough to minimize the force of the main parachute inflation, while keeping the initial descent rate high enough to ensure that the main parachute inflates predictably. Simulations were conducted in OpenRocket for each parachute diameter. A 24 parachute caused the rocket to experience high accelerations during main parachute deployment, which could damage the rocket or the fragile material payload. A 48 parachute resulted in an initial descent rate that may not allow the main parachute to inflate properly. A 36 drogue parachute was selected to ensure 20 P a g e

28 that the main parachute inflates while limiting maximum acceleration. Table 11 shows a decision matrix for the drogue parachute selection. Table 11 - Decision Matrix: Drogue Parachute Option 24 Fruity Chutes Classic Elliptical 36 Fruity Chutes Classic Elliptical 48 Fruity Chutes Classic Elliptical Decision Matrix Drogue Parachute Cost Descent Rate Max Force O Δ X O O Δ Δ Δ О Overview Decision Explanation For a low cost, a relatively quick descent rate can be achieved, but at the cost of a large maximum force at main parachute deployment. For a low cost, a good descent rate can be achieved with an acceptable maximum force at main parachute deployment. For a medium cost, a relatively slow descent rate can be achieved with a low maximum force at main parachute deployment. Main parachute selection focused on minimizing the kinetic energy of the rocket at ground impact, as this event has the greatest potential for causing costly damage to the rocket. Managing main parachute deployment acceleration was also a consideration. Using OpenRocket simulations, a 72 parachute resulted in ground impact kinetic energy greater than the 75 ft-lbf allowed by NASA. A 96 parachute was selected to give a maximum kinetic energy of 29.4 ftlbf for the aft body tube, which is the heaviest section of the rocket. Table 12 shows a decision matrix for the main parachute selection. 21 P a g e

29 Table 12 - Decision Matrix: Main Parachute Option 72 Fruity Chutes Iris Ultra 84 Fruity Chutes Iris Ultra 96 Fruity Chutes Iris Ultra Decision Matrix Main Parachute Cost Ground Impact Velocity Max Force O X O O Δ Δ Δ O Δ Overview Decision Explanation For a low cost, an unacceptable ground impact velocity can be achieved with a good maximum force. For a low cost, an acceptable ground impact velocity can be achieved with an acceptable maximum force. For a medium cost, a good impact velocity can be achieved with an acceptable maximum force. Motor Alternatives The motor was decided to be either a K or L class motor upon running simulations in OpenRocket. With the range of motors narrowed, 54mm motors were selected as that diameter was conducive to the rocket dimensions. The motors were then narrowed further by length. Finally, simulations were run on each motor to see the apogee obtained and the final motors were selected by running multiple simulations. The final three motors that were considered were from AeroTech, Aminal Motor Works, and Cesaroni Technology. The motor data can be found in Table P a g e

30 Table 13 Motor Considerations and Specifications Manufacturer AeroTech Cesaroni Technology Inc Animal Motor Works Make L850W L800 L1080BB-P Total Impulse 3695 Ns 3731 Ns 3686 Ns Weight 8.1 lbs 7.75 lbs 7.92 lbs Weight Empty 3.54 lbs 3.79 lbs 4.13 lbs Length 20.9 in 19.1 in 19.6 in Diameter 2.95 in 2.95 in 2.95 in Type Reloadable Reloadable Reloadable Burn Time 4.24 s 4.63 s 3.31 s Average Thrust 868 N 805 N 1112 N Max Thrust 1185 N 1024 N 1258 N Altitude Reached 5,379 ft 5,460 ft 5, 329 ft The L850W motor from AeroTech was ultimately selected. Cesaroni was not producing motors at the time of selection and a strict time schedule needed to be kept for the project. The L1080BB-P motor from Animal Motor Works was not chosen because of its relatively high empty weight. Using the L850W motor, the rocket has achieved a thrust to weight ratio of 5.61:1. The velocity that the rocket experiences (max) is 592 ft/s and an acceleration of 208 ft/s 2. The mach number for the rocket is Additionally, the rail exit velocity is 69.2 ft/s. With the motor selected and the materials decided, the propulsion system (housing) was designed. The bulkhead had a 5.38 diameter, 0.25 thick aluminum plate placed in front of a 3.1 outer diameter inner tube to accommodate the L850W motor. There will be two centering rings located along the inner tube with a inner diameter and a 5.38 outer diameter. These rings will be 0.25 thick 6061-T6 Aluminum. The thrust plate had the same dimensions as the centering rings and was located 0.25 from the end of the inner tube to allow for a retention system to be attached to the rocket. An exploded view of the motor mount can be found in Figure 23 P a g e

31 7. Figure 8 shows labeling of the components for the propulsion system. For a dimensional drawing, Figure 9. Figure 7 - Exploded View of the Motor Mount 24 P a g e

32 Bulkhead Centering Rings Thrust Plate Inner Tube Motor and Case Figure 8 - Propulsion Components Labeled Figure 9 - Dimensional Drawing for the Motor Mount 25 P a g e

33 The motor mount will be long, including the bulkhead. This will leave enough room for the recovery system to reach the desired pressure needed for the system. The total estimated weight of the propulsion system with propellant is lbf and the weight with no propellant is lbf. All motor mount drawings can be found in Appendix A. Recovery The launch vehicle will utilize a dual-deployment recovery system with redundant altimeters to ensure that the vehicle lands safely at a reasonable distance from the launch site. A coupling tube will house the recovery electronic systems and serve to unite the two carbon-fiber body tubes. At apogee, a black powder ejection charge will pressurize the volume above the coupling tube, separating the rocket into two sections and deploying a ripstop nylon drogue parachute. When the rocket has descended to an altitude of 1000 feet, a second black powder ejection charge will pressurize the volume below the coupling tube, separating the rocket again and deploying the main parachute, which will also be made from ripstop nylon. All three sections of the rocket will be tethered together using tubular nylon cord, which shall be protected from the ejection charges by flameproof fabric and attached to aluminum bulkheads using U-bolts. At the heart of the recovery system are two PerfectFlite StratoLogger CF altimeters, shown in Figure 10. This particular model was chosen for its simplicity and cost-effectiveness; while the StratoLogger CF has a relatively limited set of functions, the alternatives considered were generally much more expensive and provided unnecessary features for the purposes of a simple dual-deployment operation. 26 P a g e

34 Figure 10 - PerfectFlite Stratologger CF Altimeter The altimeters will be powered independently of each other using two 9-volt batteries, and armed independently using two rotary locking switches accessible externally via two small holes in the airframe. These holes also serve to expose the altimeters to the external air pressure to allow accurate determination of the launch vehicle s altitude. To preserve the redundancy of the system, each altimeter will operate on completely separate circuits, including separate igniters for each altimeter. Lead wires will connect the altimeter outputs to terminal blocks mounted to the outside of the coupler bulkhead. The terminal blocks allow for quick replacement of igniter wires. A block diagram showing the redundant recovery electrical system is shown in Figure P a g e

35 Figure 11 - Block diagram of major recovery system electrical components The altimeters, batteries, and arming switches will be mounted to a plywood sled inside the recovery bay. This plywood sled will be located on two threaded rods that are secured at each end to aluminum bulkheads, as shown in Figure 12. The bulkheads will mount flush to the coupling tube to isolate the altimeters from the pressure bursts associated with the black powder ejection charges. 5/16 steel U-bolts with steel backing plates will serve as attachment points for the 1 tubular nylon recovery harness. 28 P a g e

36 Figure 12 - Recovery bay bulkheads and hardware A very important consideration in the development of a recovery system for a high-powered rocket is the parachute configuration. The launch vehicle will utilize a system that houses the drogue and main parachutes in separate compartments on opposite sides of the recovery bay, as shown in Figure 13 and Figure 14. These compartments are bounded by aluminum bulkheads that are epoxied to the body tube and have identical U-bolts that serve as mounting points for the recovery harness. This configuration keeps all three sections of the rocket (nose, recovery bay, booster) tethered together after parachute deployment via the tubular nylon recovery harness, as shown in Figure 13. As per the decision matrices in the Component Alternatives section, a P a g e

37 Fruity Chutes Classic Elliptical parachute will serve as the drogue parachute, and a 96 Fruity Chutes Iris Ultra parachute will serve as the main parachute. A 25 length of tubular nylon will be used for the drogue parachute tether, and a 35 length will be used for the main parachute tether. Figure 13 Exploded View; Recovery System Figure 14 - Recovery system layout within airframe 30 P a g e

38 25 35 Figure 15 Tethering of Rocket Sections 31 P a g e

39 Payload Electronic Payload The electronic payload is located in the nosecone of the rocket. It contains a Atlus Telemega, which will record and transmit all flight data and a battery. The entire payload will be water proof. The location of the payload with respect to the nosecone can be seen in Figure 16. An annotated exploded view can be seen in Figure 17. Figure 16 - Electronic Payload within Nosecone Figure 17 - Exploded View of Electronic Payload 32 P a g e

40 The assembly of the electronic payload with respect to the nosecone is shown in Figure 18. clearly shown. Figure 18 - Exploded Electronic Payload View with Nosecone In Figure 19, the assembled payload can be seen and in Figure 20 the mounting studs are Figure 19 - Top View, Assembled Electronic Payload Figure 20 - Bottom View, Assembled Electronic Payload 33 P a g e

41 Fragile Material Payload The main objective of the fragile material housing payload is to protect an unknown object(s) throughout the duration of the flight. To do this, many designs were brainstormed and down selected. One main idea developed was to have a supporting material within a cylinder to house the object and keep it in place within the rocket. The main alternative was a spring damper system to reduce the force of the rocket felt by the payload entirely. Both ideas were combined, resulting in the current design. Project ACE s design consists of two concentric cylinders, one with supplemental material inside to hold the fragile material in place. The entire system consists of two different springs in series and parallel meant to absorb both large and small vibratory impacts. Concentric cylinders within the rocket tube to allow payload oscillation. Figure 21 shows allof the components of the payload spring system in an exploded view. 34 P a g e

42 Figure 21 - Payload Exploded View a) Cylinder 2 b) Cylinder 1 c) 12 CR1-400 Wire Rope Isolators d) Baseplates e) Hardware used to assembly the system f) Main springs g) Coupling baseplate h) Outer most baseplate i) U bolt holding assembly together Project ACE plans to use 12 CR1-400 Enidine wire rope isolators. These will allow oscillation of Cylinder 1 to reduce forces transmitted to the fragile material, small vibrations, and overall acceleration. The concentric cylinders and wire rope isolators can be seen in Figure P a g e

43 Figure 22 - Components of the Main Payload a) Cylinder 1 b) Cylinder 2 c) Wire Rope Isolators In Figure 22, Cylinder 2 will be concentric with the rocket s main body tube as well as Cylinder 1. Cylinder 2 will be made of aluminum while cylinder 1 will be made of ABS plastic and will be 3D printed then sealed. c shows the location of 3 of the 12 wire rope isolators. A dimensioned drawing of Cylinder 2 can be found in Appendix A. Cylinder 1, shown in Figure 23 is designed to have inside dimensions of 3.5 diameter and 9 long. The dimensioned drawing can be seen in Appendix A. The maximum envelope given to teams in the project requirements is 6 long, however, we designed the cylinder to have 3 extra 36 P a g e

44 of thread for the cap to screw down at variable lengths. The reason for this is that Cylinder 1 will contain support material (material to be determined through testing) and when the unknown object(s) are placed within the container, the support material will be displaced the same volume as the object(s). To be sure the support material firmly holds the object(s), the lid will screw to variable distances to compress the material and object(s) regardless of their size. Figure 23 Payload Inner Cylinder Attached to Cylinder 2 are 5 base springs - designed to absorb most of the large impact forces such as the initial takeoff, parachute deployment, and landing. Prior to completing the first mathematical model utilizing Simulink, these springs were not included. However, the forces induced on Cylinder 1 and thus the fragile material were too large so a series and parallel spring system was created by introducing the 5 base springs. These springs can be seen in Figure 24 and are labeled a. 37 P a g e

45 Figure 24 - Springs Attached to the Main Payload and Bulkhead Attachment Then entire system consisting of Cylinder 1, Cylinder 2, 12 wire rope isolators, and 5 base springs, oscillates within the body tube of the rocket and is mounted to the bulkhead separating the payload bay and the recovery bay. This bulkhead can be seen in Figure 24 is labeled b. The walls of the payload bay, as well as the outside of Cylinder 2, will be lubricated to ensure smooth translation during oscillation with graphite powder. Again, the exploded view of the entire system can be seen in Figure P a g e

46 Testing on the payload will not only decide the support material but will also test the validity of the math model s ability to select springs. Testing will be performed on the entire spring and concentric cylinder system with the matrix seen below in Table 14. Table 14 - Testing Matrix for Fragile Material Testing Materials Weight # To be Tested Egg 1.75 oz 2 Glass Stir Rod.2 oz 1 Glass Sheet N/A N/A Light Bulb 1.1 oz 3 Small Ceramic/Porcelain China N/A N/A Contact Support Materials (within Cylinder 1) Weight per cubic ft. Density Grain Size Liquid/Solid Viscosity Aerogel N/A N/A N/A N/A N/A Packing Peanuts.2 lb N/A Varies Solid N/A Styrofoam Pellets.2 lb N/A Varies Solid N/A Non-newtonian Fluid N/A N/A N/A Both Varies High Density Foam (cubes/sheets) Varies.93 g/cm 3 As needed Solid N/A Spray in High Density Foam (injection system) Varies 3 lb/ft 3 N/A Solid N/A Testing is a primary part of this section as it will not only give validation to the design but will also show shortfalls and areas of interest going into the demonstration. Testing for the payload as a system will be done with drop tests at various heights associated with desired impulse forces. The three main phases of flight to be tested will be the impact force, the main parachute deployment force, and the force caused by the motor. The two impulse forces, the parachute, and impact, will be estimated and then tested with drop tests. Additionally, the team will be modifying the Charpy Impact Tester to give desired impulses. The engine thrust will be tested by selecting points of interest from the thrust curve given by the manufacturer and mimicking those forces at that point in time again with a drop test. Each test will be repeated with the top filler material choices from the material and testing object matrix found in Table P a g e

47 The math model for the payload started as an analysis of the system in the form of free body diagrams. The system drawing can be seen below in Figure 25. Figure 25 - System Drawing and Force Balance Figure 25 is the system diagram for the entire rocket (M1), the concentric cylinder and spring assembly (M2), and Cylinder 1 including the unknown object(s) (M3). This system was then derived into free body diagrams and accompanying force equilibrium equations seen below in Figure 26- Figure P a g e

48 Figure 26 - Free Body Diagram for the Entire Rocket and Force Balance Figure 27 - Free Body Diagram for the Spring Assembly and Force Balance 41 P a g e

49 Figure 28 - Free Body Diagram for Cylinder 1 and the Fragile Material and Force Balance The first round of Simulink models did not work since the model was under constrained. For this reason, the team redesigned the Simulink model as a base excitation vibration model. The main change this induced was that the only input in Simulink for M1 was the external force for the given situation. Three different models were made simulating three force events. A table showing these values is given below in Table 15. Table 15 - Force Events for the Simulink Model Force Thrust Main Parachute Deployment Landing Input Thrust curve 400 ft. lb 75 ft. lb The way the math model helped us to select the needed springs was by selecting one of the inputs from Table 15 above and then iteratively selecting springs until one was found that fit our 42 P a g e

50 application. Different k (spring constant) and c (damping coefficient) values were inserted in the model for k1 and 2 and c1 and 2. This model can be seen below in Figure P a g e

51 Figure 29 Simulink Mathematical Model 44 P a g e

52 The mathematical model shows the Simulink model used to select springs for the system. The approach taken was entering in estimated or known forces as the input for the base excitation model for mass 1 and then three graphical outputs were created, position, velocity, and acceleration. The acceleration graphs were then used to determine the overall force on the payload and springs were optimized by plugging in various k and c values to determine the best reduction in force and acceleration on the payload or mass 3. To determine the best spring selection, it was decided to perform iterations using said k and c values since those could be easily selected from standard commercial parts and then the system could be solved to find the resultant forces and accelerations desired. We decided the best spring was selected when the maximum displacement of the spring was reached without bottoming out and the smallest force and accelerations were transmitted to the payload. The damping coefficients (c) present in the model were calculated from the manufacturer specifications that stated the damping was 5 percent of the spring constant. The final values for the constants as listed in Table 16. Table 16 - Final Values for Constants Final values (N/m) k v c v 7.6 k s c s Mission Performance Predictions The main source of flight simulator data used for flight predictions was OpenRocket. This software s flight simulation is based off of an atmospheric model that estimates variable 45 P a g e

53 conditions with changing altitude. This model assumes ideal gas for the air. This model also considered a wind model, importing the Kaimal spectrum equation and the assumption that the wind speed is uniaxial. Another assumption the program makes is that the earth is flat, which negates Coriolis effects. Additionally, turbulence intensities are based on wind farm load design standards, which may or may not translate to higher altitudes. With these models taken into consideration, the program runs a 4 th order Runge-Kutta integration method with the following steps: 1. Initialize the rocket in a known position and orientation when time is equal to zero 2. Compute the local wind velocity and other atmospheric conditions 3. Compute the current wind speeds, angle of attack, and other flight parameters 4. Compute the aerodynamic forces and moments on the rocket 5. Compute the motor thrust and center of gravity 6. Compute the mass and moment of inertia of the rocket from linear and rotational acceleration of the rocket 7. Numerically integrate acceleration to the rocket s position and orientation during the time step t and update the time. (Niskanen, 2009) The program computes steps 2-6 until the rocket has reached its end time which is normally reaching the ground (Niskanen, 2009). This open source software is similar to commercially available software such as Rocksim. OpenRocket originated at Helsinki University of Technology as a Master s Thesis by Sampo Niskanen (Niskanen, 2009). Experiments working to prove that OpenRocket is accurate found that during one test on a B size motor that the program over estimated the altitude by about 16%, and for a C size motor altitude was over estimated by 7%. For another experiment, a larger motor was used and the 46 P a g e

54 Altitude (ft) program under estimated altitude by 16%. However, the program was also compared to commercially available software and it was found to be as accurate as Rocksim. In the same experiment, Rocksim s uncertainty was B motor 24%, C motor- 19% and Larger motor- 12% (Niskanen, 2009). For Project ACE s rocket, the plan is to add around 50% of the allowable ballast to lower the projected altitude to exactly 5,280 ft. Figure 30 shows the predicted altitude results from OpenRocket. The inputs for the OpenRocket simulation can be found in Appendix B Time (s) Figure 30 - Predicted Altitude from OpenRocket Simulation The predicted altitude from the OpenRocket software is 5,379 ft. The inputs for the simulation were 4 mph for the average wind speed with a standard deviation of 0.4 mph. The inputs for the OpenRocket simulation can be found in Figure P a g e

55 Figure 31 - OpenRocket Flight Simulation Inputs The predicted altitude from the flight simulation of OpenRocket was compared to the predicted altitude using the same inputs and rocket design in Rocksim. This was to show that the altitude was predicted in multiple ways. The altitude that was predicted for the Rocksim model was 5,368 ft which was close to the predicted altitude from OpenRocket. Figure 32 for altitude predictions from the Rocksim simulation. This altitude showed a percent difference of 0.20% between the Rocksim simulation and the OpenRocket simulation. The OpenRocket value is used as the base because OpenRocket is the original program used to calculate the altitude of the rocket. Inputs for the Rocksim simulation are provided in Figure 33. Equation 1 showed how the percent difference was calculated and Equation 2 showed the calculated values for the percent difference. 48 P a g e

56 Altitude (ft) % Difference = OpenRocket Value Rocksim Value OpenRocket Value Eq. 1 % Difference = 5379ft 5368ft x 100 = 0.20% Eq ft Time (s) Figure 32 - Predicted Altitude from Rocksim Simulation 49 P a g e

57 Figure 33 - Inputs for Rocksim Simulation The thrust curves produced by the simulations show the same thrust for the L850W motor. The thrust curve produced by Aerotech is shown in Figure 7. The thrust curve from the OpenRocket simulation can be found in Figure 35 and the thrust produced in Rocksim simulation can be found in Figure 36. The components that were used in the simulations can be found in Appendix B, along with weights of each component. Figure 34 - Thrust Curve from AeroTech Motor 50 P a g e

58 Thrust (N) Thrust (N) Time (s) Figure 35 - Thrust Curve for the L850W Motor in OpenRocket Time (s) Figure 36 - Thrust Curve for the L850W Motor in Rocksim 51 P a g e

59 The Center of Gravity (CG) is in. from the tip of the nosecone. The Center of Pressure (CP) is in. from the tip of the nosecone. This produces a stability of 3.69 calipers. This was determined using OpenRocket. CG CP Figure 37 - Center of pressure and gravity Using the average atmospheric and weather conditions for an April day in Huntsville, Alabama, an OpenRocket simulation was conducted to predict the performance of the recovery system. The drogue parachute provides a safe initial descent rate of 50.2 ft/s, which is suitable for keeping the landing site within walking distance of the launch site while also ensuring that the main parachute does not open under excessive speed. The rocket will impact the ground with a speed of 14.8 ft/s, giving each section of the rocket a kinetic energy under the maximum allowable 75 ft-lbf as shown in Table 17 below. Table 17 - Kinetic energy of each section upon landing Section Mass (lb) Kinetic Energy (ft-lbf) Nose Cone & Payload Recovery Bay Booster Apart from these average atmospheric conditions, drift distances were simulated in OpenRocket for different wind speeds as shown in Table 18. These distances assume a perfectly 52 P a g e

60 vertical launch angle with medium atmospheric turbulence. As the simulated drift distance for 20 mph winds is over the allowable distance of 2500 ft, the main parachute deployment altitude will be lowered using the altimeter s built-in adjustment features in the event of excessive wind speeds on launch day. Table 18 - Landing site distance from launch site by wind speed Wind Speed (mph) Lateral Distance (ft) Safety Overview The University of Evansville, in conjunction with Project ACE and all team members, is dedicated to a successful launch, and, most importantly, safe operation of the rocket throughout all phases of the project. Led by Safety Officer, Bryan Bauer, the team members will be saturated with information regarding proper safety protocols for each stage of the project. In addition to this, all team members will be briefed on the hazards that are specific to the materials they will come in direct contact with so that accidents and injury can be prevented. Furthermore, material data sheets (MSDS) will be available to all students in the working area, so that potential hazards can be identified before construction begins. 53 P a g e

61 During the construction and fabrication phase of the project, students will work in groups of no less than two, to ensure that at least one team member would be able to provide immediate assistance and call for help in the event that an accident occurs. Additionally, the team safety officer will monitor use of personal protection equipment (PPE), such as glasses and gloves amongst other things, during construction to ensure all team member are safe. The team safety officer will also ensure that the energy systems lab is equipped with working smoke detectors and fire extinguishers as well as first aid kits. During the sub-scale and full-scale testing of the rocket, all team members will wear safety glasses and will maintain a safe distance from the launch pad. Due to the risks associated with various facets of the rocket, checklists will be developed and reviewed before final assembly and launch to guarantee safety of all team members and spectators. Additionally, the team will work together to construct a hazard analysis which will be used to identify risks, their causes, and proposed mitigations in order to minimize the chance of accident and injury, and ensure safe operation. This focus on safety and education of all team members will create optimal working conditions, which ultimately will keep the project on schedule and allow for safe and successful launch. 54 P a g e

62 Final Assembly Checklist Initials Check-Off Points Check rocket tube for cracks, bumps, abrasions or any other imperfections that could have been acquired during construction or transport that could adversely affect the flight of the rocket. Check parachute for any inadequacies or tears that could alter deployment and safe landing. Ensure that the parachute is packaged properly inside the rocket tube. Check payload for any cracks or chips that could have been acquired during transport. Check motor and casing to ensure it is not wet or containing any visible imperfections that would cause a misfire or deviation from the ideal flight path. Ensure recovery harness is properly attached for flight readiness. Check motor mount for structural integrity. Check primary fins for cracking or bowing. Check thrust plate and couplers for solid attachment and structural integrity to ensure proper flight. 55 P a g e

63 Check avionics bay for proper functioning to ensure noting was broken or altered during transport Check nosecone for structural integrity and secure attachment to the rest of the rocket. Insert motor into casing and check for secure fit Ensure all connections of the rocket are solid and cohesive UE SLI Safety Officer Signature UE SLI Team Lead Signature UE SLI Adult Educator Signature 56 P a g e

64 Launch Procedures Checklist Launch Procedures Checklist Ensure a safe working area before unloading the rocket and bringing it to the launch pad. Check the safety and readiness of team members by ensuring all team members have on safety glasses and other proper PPE for the part of the rocket they will be handling Visually inspect the rocket for proper connections between all sections before placing on the launch pad. Test electronics (i.e. camera, altimeter, etc.) to ensure they are fully functional and turned on before launch Check launch pad and guide rails for readiness Place rocket on launch pad Have non-essential team members move away from the launch pad to the safe viewing distance Arm the rocket motor for ignition Disarm all safeties on the rocket Have remaining team members move to safe viewing distance to watch the launch 57 P a g e

65 Check with Range Safety Officer (RSO) to ensure all codes and rules are met and the rocket is clear for launch. Initiate rocket ignition. Check for proper ignition UE SLI Safety Officer Signature UE SLI Team Lead Signature UE SLI Adult Educator Signature *Note: The launch procedures checklist will be edited during the course of the project to include more detail as the team learns more about standard launch procedures and the setup of the rocket. 58 P a g e

66 Personnel Hazard Analysis A preliminary personnel hazard analysis was conducted to identify hazards, causes and resulting effects. This analysis was created make team members aware of potential hazards, and lists mitigations to reduce the chance of risk or injury during the course of the project. This analysis is summarized in Table 21. Table 19 - Personnel Hazard Analysis Risk/Hazard Effect/Severity Severity Likelihood Mitigation and Control Epoxy Inhalation of toxic fumes, accidental ingestion, or contact with skin leading to irritation or rash Minor High Work in well ventilated spaces Dust Particles Inhalation of dust particles from sanding or machining operations resulting in breathing problems Minor High Wear mask when sanding to avoid inhaling dust particles Heavy Tools and Machinery in Lab Improper handling of shop tools or machining operations leading to personal injury or destruction of equipment Significant Medium Ensure proper training for all team members working with any tool or machinery in shop Rocket Propellant Exposure to rocket fuel in contact with skin leading to irritation and burns Major Medium Properly transport motor from offsite location to launch site Black Powder Gases may be toxic if exposed in areas with inadequate ventilation. Also keep away from open flame, sparks, and heat Major Low Store in portable fireproof case to keep away from fire and high temperatures Craft and Exacto Knives Cuts leading to injury as a result of precision cutting operations on fins or other pieces of the rocket body Minor Medium Ensure at least one teammate is working alongside the person doing the cutting. Practice safe cutting procedures by cutting away from body. 59 P a g e

67 Fire Burns, significant and/or fatal injury, or damage to school from fire as a result of faulty wiring, or improper handling of the motor and black powder Major Low Store a fire extinguisher in the room where the rocket will be constructed. If an object starts to overheat, let it cool and have the fire extinguisher ready Handheld Tools Bruises, cuts or scrapes from mishandling of basic handheld shop tools such as hammer or saw Significant High Be aware of surroundings when operating the handheld tools and ensure proper training before any construction is undertaken. Failure Modes and Effects Analysis A preliminary Failure Modes and Effects Analysis of the proposed design of the rocket, payload, payload integration, launch support equipment, and launch operations, which can be seen in Table 22, was completed to identify hazards, effects and proposed mitigations. Table 20 - Failure Modes and Effects Analysis Risk/Hazard Effect Severity Likelihood Motor Handling/Accidental Ignition Launch Failure Main Parachute Deployment Failure Drogue Parachute Deployment Failure Improper handling or storage of motor resulting in accidental or unexpected ignition Failure of motor to ignite and launch rocket properly Failure of the secondary parachute to deploy leading to freefall or unstable flight of rocket back to the ground Failure of the initial parachute to deploy leading to freefall or unstable flight of rocket back to the ground Major Significant Major Significant Low Low Low Low Proposed Mitigation Properly transport motor from offsite location to launch site. Ensure proper connections before launch Maintain safe distance from launch pad. Have team mentor/safety officer inspect rocket on launch pad Maintain safe distance from launch pad Maintain safe distance from launch pad 60 P a g e

68 Instability During Flight Altimeter or Other Electronics in Avionics Bay Malfunction/Fall Off Coupler Excessively Tight Payload Not Secured Properly Failure of the rocket to maintain its projected flight path due to unforeseen design flaw or in flight malfunction Potential short circuiting or harm to spectators below Failure of parachute to deploy leading to damage to rocket Inability to return materials without breaking Major Minor Major Minor Low Medium Low Medium Maintain safe distance from launch pad Verify all electronics work properly before launch and are firmly attached to the rocket Run multiple tests to ensure proper amounts of black powder is used to allow rocket to separate Take caution when inputting payload into rocket before launch and ensure all items are properly sealed and secured before launch Environmental Considerations Additionally, when considering the safety and impact of the rocket, considerations must be given to how the vehicle will impact the environment, and how the environment will impact the vehicle. This analysis is shown below in Table 23. Table 21 - Environmental Consideration Analysis Risk/Hazard Effect and Impact Severity Likelihood Mitigation and Control Vehicle Effects on Environment Epoxy Fumes When epoxying various pieces of the rocket together, harmful fumes are released into the atmosphere Minor High Work in well ventilated spaces and dispose of waste properly 61 P a g e

69 Dust Particles Small dust particles from sanding or machining operations are released into the environment which can result in breathing problems Minor High Wear mask when sanding to avoid inhaling dust particles and try to contain dust when sanding opposed to freely releasing it into surrounding air. Rocket Motor Ignition Upon ignition, the motor reaches high temperatures and hot exhaust is released, which could potentially burn the areas where the rocket is launched or lands Major Low Place flame resistant material beneath the launch pad to avoid burning the immediate surroundings or starting a fire Debris from Rocket If various pieces of the rocket do not stay intact during decent, or the parachutes do not operate properly, pieces of the rocket could break off during flight or upon impact and be irretrievable, leading to minor environmental harm. Significant Low Ensure fully functioning parachutes before launch via pre-launch checklist and check that all components of the rocket and payload are accounted for upon return. Environmental Effects on Vehicle Water Precipitation and moisture within the rocket could affect the structural integrity of the rocket, or could lead to malfunctions of the electronics housed in the avionics bay Significant Low Avoid launching rocket in wet conditions and ensure a dry area for storage and transport Wind Strong wind or unpredictable wind gusts can cause the rocket to deviate from its ideal flight path and can lead to damage to the rocket and potential harm to spectators Significant Medium Avoid launching rocket on days where high speed winds or unpredictable, strong wind gusts are present 62 P a g e

70 Humidity Humidity can lead to moisture in the body of the rocket which can lead to corrosion and weakening of various materials used to construct the rocket. It can also negatively impact onboard electronics Minor Low Store rocket in a dry area to avoid moisture entering the rocket over time via humid air General Risk Assessment Finally, a general risk assessment was conducted in order to account for various extraneous risks not accounted for in previous sections, such as time, resources, the budget, scope, and functionality. Seen in Table 24. Table 22 - General Risks Associated with the Project Risk/Hazard Effect Impact Value Likelihood Proposed Mitigation Limited Resources Due to the new nature of the project to this team specifically, if the team is unable to find valuable insight from external sources, the design and performance of the rocket could suffer High Medium The team will work with faculty members as well as local rocketry club members in order to gain a better understanding of rocketry and develop a functional rocket. Tight or Minimal Budget Lack of flexibility in the budget could lead to the team being forced to use parts that are not optimal, or being unable to replace parts of the rocket that are broken during testing High High The team and its adult educators will apply for grants and fundraise to provide the team with a flexible budget beyond the normal amount of money allotted to the project by the school 63 P a g e

71 Mismanagement of Time Inability of the team to keep up with the initial schedule set forth in the task breakdown could lead to major delays, poor quality of work, or the rocket not being completed by competition Medium Low Team members will fill out weekly time cards and log their hours in the task breakdown in order to ensure everyone remains on schedule Underestimation of Scope of Work Increase in Safety Regulations Failure to properly account for the work needed to complete the project could lead to the project running behind schedule and various facets of the rocket not being completed in a quality manner Adding material to the rocket in order to increase safety will result in an increase in expenses Medium Low Low Medium There will be constant communication amongst all team members and with NASA to ensure the scope of work is clear The team will design and downselect with safety in mind, and will clearly identify all safety measures before construction so that additional, last-minute safety measures do not have to be taken that will inflate the budget. Project Plan Requirements Compliance Table 23 - Requirement Compliance Handbook Number Summarized Requirement NASA Requirements Verification Description of Verification Plan Method(s) 1.1 The vehicle shall deliver the science or engineering payload to an apogee altitude of 5,280 feet above ground level (AGL). Test Analysis The rocket team will utilize OpenRocket, RockSim, CFD, & test flight data to achieve an accurate prediction of altitude. 64 P a g e

72 Handbook Number Summarized Requirement The vehicle shall carry one commercially available, barometric altimeter for recording the official altitude used in determining the altitude award winner. All recovery electronics shall be powered by commercially available batteries. The launch vehicle shall be designed to be recoverable and reusable. Reusable is defined as being able to launch again on the same day without repairs or modifications. The launch vehicle shall have a maximum of four (4) independent sections. The launch vehicle shall be limited to a single stage. The launch vehicle shall be capable of being prepared for flight at the launch site within 4 hours. NASA Requirements Verification Method(s) Inspection Inspection Test Inspection Inspection Inspection Demonstration Inspection Demonstration Description of Verification Plan The rocket will house a Atlus Metrum TeleMega altimeter in the nosecone to record the official altitude used in determining the altitude award winner. Batteries & altimeter will be purchased from online rocketry sources. The rocket is reusable in design because our team is using a motor that has refuels that can be reloaded into the motor under supervision. The launch vehicle will have 3 independent sections: the aft body tube, the bow body tube and nosecone, and the coupler. The launch vehicle shall be a single stage. The launch vehicle will be designed with an efficient and quick to construct design that requires fewer than 4 hours to prepare. 65 P a g e

73 Handbook Number Summarized Requirement The launch vehicle shall be capable of remaining in launchready configuration at the pad for a minimum of 1 hour without losing the functionality of any critical on-board component. The launch vehicle shall be capable of being launched by a standard 12-volt direct current firing system. The launch vehicle shall require no external circuitry or special ground support equipment to initiate launch (other than what is provided by Range Services). NASA Requirements Verification Method(s) Test Inspection Test Inspection Description of Verification Plan The launch vehicle design will ensure all components have a life of greater than 1 hour without loss of functionality. The ignition system will be using a 12 volt direct current firing system. There will be no external circuity for the ignition system because it will be a ground based ignition system being placed underneath the rocket before launch with 300 ft of cord between the igniter and the controller. 66 P a g e

74 Handbook Number Summarized Requirement The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR), Tripoli Rocketry Association (TRA), and/or the Canadian Association of Rocketry (CAR). Pressure vessels on the vehicle shall be approved by the RSO. The total impulse provided by a University launch vehicle shall not exceed 5,120 Newtonseconds (L-class). The launch vehicle shall have a minimum static stability margin of 2.0 at the point of rail exit. The launch vehicle shall accelerate to a minimum velocity of 52 fps at rail exit. NASA Requirements Verification Method(s) Inspection Inspection Test Analysis Test Analysis Test Analysis Description of Verification Plan The motor being used is a solid fuel motor from AeroTech. The motor is the L850W. No pressure vessels will be used. The motor will produce an impulse of 3695 N-s which is below the specified total impulse that is allowed. The launch vehicle will have a static stability margin of The rocket team will utilize OpenRocket, RockSim, CFD, & test flight data to achieve an accurate prediction of minimum velocity at rail exit. The current value is 67.2 ft/s. 67 P a g e

75 Handbook Number Summarized Requirement All teams shall successfully launch and recover a subscale model of their rocket prior to CDR. All teams shall successfully launch and recover their fullscale rocket prior to FRR in its final flight con- figuration. Any structural protuberance on the rocket shall be located aft of the burnout center of gravity Vehicle Prohibitions Vehicle must deploy a drogue parachute at apogee, followed by a main parachute at a much lower altitude. A successful ground ejection test for both parachutes must be conducted prior to sub- and full-scale launches. NASA Requirements Verification Method(s) Test Analysis Test Analysis Test Analysis Inspection Test Analysis Demonstration Inspection Test Description of Verification Plan A subscale model with comparable weights, lengths, and masses will be launched prior to the CDR. The project schedule will ensure a fullscale rocket launch occurs before the FRR. The rocket will have 3 bolts holding the nosecone to the bow body tube and shear pins holding the coupler to the bow and aft body tubes. These structural protuberances are all located aft of the burnout center of gravity The launch vehicle will follow all prohibitions laid out in section 1.19 of the 2017 SL NASA Student Handbook. Dual-deployment altimeters will be programmed to fire ejection charges at apogee and at ~1000 feet. Multiple ejection tests will be conducted prior to sub- and full-scale launches. 68 P a g e

76 Handbook Number Summarized Requirement No part of the launch vehicle may have a kinetic energy of greater than 75 ft-lbf at landing. Recovery electrical circuits must be independent of payload circuits. Recovery system must include redundant, commercial altimeters. Motor ejection cannot be used for primary or secondary deployment. Each altimeter must be armed by a dedicated switch accessible from the rocket exterior. Each altimeter must have a dedicated power supply. Each arming switch must be lockable to the ON position. Removable shear pins must be used to seal the parachute compartments. NASA Requirements Verification Method(s) Analysis Demonstration Inspection Inspection Demonstration Inspection Inspection Inspection Inspection Inspection Description of Verification Plan Parachute sizes will be optimized to minimize kinetic energy at ground impact. Recovery electronics will be located in a separate, shielded coupler. Two PerfectFlite Stratologger CF altimeters will be used. Black powder ejection charges will be used to eject parachutes. Locking rotary switches and LED indicators will be used to confirm the state of the recovery electronics. Separate 9-Volt batteries will be used to power the altimeters. Locking rotary switches will be used to arm each altimeter. Threaded nylon shear pins will be used to seal the parachute compartments. 69 P a g e

77 Handbook Number Summarized Requirement Tracking device(s) must transmit the position of any parts of the launch vehicle to a ground receiver. Recovery system electronics must not be adversely affected by any other on-board electronics. Design container capable of protecting an unknown object of unknown size and shape. Object must survive duration of flight Once the object is obtained, it must be sealed in its housing until after the launch and no excess material may be added after receiving the object. Each team shall use a launch and safety checklist NASA Requirements Verification Method(s) Test Demonstration Inspection Test Inspection Testing Testing Demonstration Inspection Description of Verification Plan All parts of the launch vehicle will be tethered together; position will be transmitted via a flight computer in the nosecone. Recovery electronics will be located in a separate, shielded coupler. Math model is used to develop spring system in conjunction with a concentric cylinder model to provide sufficient vibration dampening and force reduction. The spring and concentric cylinder design will be tested with a matrix of different support materials as well as testing materials to assure the unknown object(s) can survive the flight during demonstration. Support material within cylinder 1 that allows object to be inserted and not spill any material such as a high viscosity fluid or malleable solid. Final assembly and pre-launch checklists will be created and reviewed at the appropriate time to ensure safe launch of the rocket and all members involved in the launch 70 P a g e

78 Handbook Number Summarized Requirement Each team shall identify a student safety officer who shall be responsible for the safety of the team and ensure all proper rules and guidelines are followed The team safety officer shall monitor team activities with an emphasis on safety throughout the design, construction, and testing of the rocket by maintaining MSDS sheets and hazard analyses Each team shall appoint a mentor who has certification and is in good standing with the NRA. This member will be designated as the individual owner of the rocket and assumes liability During test flights, teams shall abide by the rules and guidance of the local rocketry club's RSO NASA Requirements Verification Method(s) Inspection Inspection Inspection Demonstration Description of Verification Plan The team has appointed a safety officer to monitor the safety of the team throughout the project and ensure all federal rules and laws are met. The team safety officer will monitor the progress of the project emphasizing the proper safety procedures for the current stage of the project. The team has assigned an school faculty member to mentor the project to provide valuable insight on the rocket design and construction as well as assume full liability of the rocket. Team will converse with RSO at local rocketry club to ensure all of their chapter s rules and regulations are abided by. 71 P a g e

79 Handbook Number Summarized Requirement Teams shall abide by all rules set forth by the FAA Students shall do 100% of the project excluding motor / black powder handling. A detailed project plan shall be maintained. Foreign National members shall be identified by the PDR. All team members attending launch week shall be identified by the CDR. The educational engagement requirement shall be met by the FRR. The team shall develop and host a website for documentation. The team shall post & make available for download all deliverables by the specified date. All deliverables must be in PDF format. NASA Requirements Verification Method(s) Demonstration Demonstration Inspection Demonstration Inspection Inspection Inspection Test Inspection Inspection Description of Verification Plan Team will converse with NASA lead safety officer and thoroughly research all rules and regulations set forth by the FAA to ensure all rules and regulations are abided by. The team will continuously demonstrate an independently managed and executed project. The team lead will routinely monitor this quality. Documents for scheduling, budget tracking, outreach, and safety will be continuously updated and reported. The team lead will ensure that any Foreign National members are clearly indicated in the PDR. It will be checked that a list of team members, with indications of those attending launch week, will be included in the CDR. The Educational Engagement lead shall confirm that all documentation has been received and approved by NASA prior to the FRR. Team members will periodically confirm that the website is functioning as intended by opening each posted document. The team lead shall confirm that all documents are posted prior to the specified date. The team lead shall confirm that all documents posted are in PDF format. 72 P a g e

80 Handbook Number Summarized Requirement A table of contents must be included in all reports. Page numbers shall be provided in each report. The team shall provide videoconference equipment needed for reviews. All teams shall use launch pads provided by the SLS provider. The team must implement the EIT accessibility standards. NASA Requirements Verification Method(s) Inspection Test Demonstration Test Demonstration Demonstration Description of Verification Plan The team lead shall ensure that a table of contents is located at the start of each report. Page numbers shall be checked to the table of contents to ensure continuity throughout the report. Videoconference rooms will be reserved and trialed immediately prior to each design review. The team shall design the rocket to utilize launch rail. If software or applications are created (not planned) the team will abide by 36 CFR Part Otherwise, all components containing software will be checked to ensure compliance. Team requirements have been developed in addition to the NASA requirements. These can be seen in Table 24. Table 24 - Team Requirements Team Requirements Number Requirement Verification Method Description of Verification Method 1 All reports shall be compiled at least three days prior to NASA due dates. Demonstration Reports shall be completed, according to team schedule, prior to NASA due dates to allow for revision time and mitigate risk of late submissions. 73 P a g e

81 Team Requirements Number Requirement Verification Method Description of Verification Method Each member of the team shall have a working knowledge of each subsystem. Safety shall be made the team s first priority. Altimeters shall be in good working order. The tracking system shall be in good working order. A solid output signal must be given from triggered altimeters. All circuits shall be checked prior to use. Impulse for the parachute deployment shall be determined experimentally. A spring constant for parachute cords shall be determined experimentally. Payload must reduce force felt by object(s) by 50 % Inspection Test Test Test Test Analysis Demonstration Test Analysis Test Analysis Testing At each team meeting, every sub-section lead will review the status of their section with the entire team. The team leader will confirm that the information presented is sufficient. The safety officer will periodically ask team members what the most important aspect of the project is. All altimeters shall be flown on sub-scale and full scale flight tests. Altitude readings will be compared to confirm consistency. The tracking system shall be flown on the sub-scale and full scale flight tests. This will be used to find the rockets thus confirming its operation. All altimeters will be triggered while voltage is read on the output. This output will be read to confirm it is acceptable. All circuits will be confirmed at each node to ensure connections. The main parachute shall have an apparatus (strain gauge) attached to it that enables a force to be read as it opens at high speed. This will cut down in the large ambiguity that exists in estimating an impulse value. The spring constant shall be determined using forces related to what is experienced with parachute opening. This helps when estimating energy absorption by the cord when the chute opens. From the mathematical model, appropriate springs will be selected to induce oscillation and reduce force. 74 P a g e

82 Team Requirements Number Requirement Verification Method Description of Verification Method 11 Payload must reduce acceleration of object(s) by 35 % Testing From the mathematical model, appropriate springs will be selected to insure acceleration graphs show 35 percent reduction from inputs. Budget The budget was able to be based on a detailed parts list due to much preliminary work by the Project ACE team. This list can be seen in Appendix C. To create the budget, the team first broke down the rocket into a number of sections (i.e. recovery, aerodynamics, etc.) Then the aforementioned parts list was created for each section. The total cost of each section then had a contingency budget implemented based on the risk of that section. The aerodynamic section can be taken as an example. The parts list calls for $1, in components. $ was added to this amount to mitigate component failure risk (a new nosecone can be quickly purchased if necessary, for example.) The sum of these for all sections of the rocket is shown in the Forecasted Amount column of Table 25. Propulsion and travel were the only major budgetary change from the proposal. Propulsion increased by nearly $1,000 due to unforeseen motor costs while travel costs decreased by nearly the same amount due to the University of Evansville Department of Engineering agreeing to cover advisor (professor) travel costs. 75 P a g e

83 Table 25 - Section Level Budget Item Forecasted Amount Percent of Total Operating $ % Travel / Lodging $2, % Launch Pad $ % Aerodynamics (Body) $1, % Propulsion $2, % Main Payload $ % Electronic Payload $ % Recovery $1, % Scale Model $1, % Educational $ % Engagement Total $10, % Project ACE s funding plan has had a slight re-allocation of funding since the proposal. Less funding will be received through the student government association and more funding will be received through the college of engineering. The breakdown of project funding is shown in Table 26. Table 26 - Funding Sources Funding Amount Remaining NASA Grant $5, $5, SGA $2, $2, U.E. ENGR $2, Total $10, Schedule The team has broken up the project in numerous tasks. The full extent of these tasks and associated schedule can be found in Appendix D. To be concise, the team has combined many of these tasks into activities and developed a Gantt chart (Figure 38). For each of these activities, Project ACE is currently on schedule or ahead of schedule. In the Gantt chart, the 76 P a g e

84 yellow column represents the current week. The vertical green line indicates where the team is at for each task. For example, the team is three weeks ahead of schedule for the Rocksim model. Figure 38 - Gantt Chart 77 P a g e

85 In addition to the project tasks/activities the team has compiled a list of critical dates. These dates are crucial to the success of the project and are listed in Table 27. Table 27 - Critical Dates Due Date Activity NASA U.E. Team Project Kickoff Aug. 15, General Motor Selection/Data Sept. 30, Sept. 16, 2016 Informal Design Sketches - Sept. 21, 2016 Sept. 19, 2016 Proposal Sept. 30, 2016 Oct. 3, 2016 Sept. 27, 2016 Motor Selection/ Data Oct. 31, 2016 Oct. 7, 2016 Proposal Presentation - Oct. 24, 2016 Oct. 22, 2016 PDR Report Nov. 04, Oct. 26, 2016 PDR Flysheet Nov. 04, Oct. 26, 2016 PDR Presentation Nov. 04, Oct. 28, 2016 Sub-Scale Launch Motor Selection - - Nov. 30, 2016 Sub-Scale Launch - - Dec. 11, 2016 Design Report - Dec. 2, 2016 Nov. 29, 2016 Motor Mount Design/ FEA Jan. 13, Nov. 30, 2016 All Structural elements FEA Jan. 13, Nov. 30, 2016 CDR Report Jan. 13, Dec. 9, 2016 CDR Flysheet Jan. 13, Dec. 9, 2016 CDR Presentation Jan. 13, Jan. 11, 2017 Full Scale Launch - - Feb. 12, 2017 FRR Report Mar. 6, Mar. 1, 2017 FRR Flysheet Mar. 6, Mar. 1, 2017 FRR Presentation Mar. 6, Mar. 3, 2017 Competition Apr. 5, Apr. 5, 2017 LRR Report Apr. 6, Apr. 3, 2017 UE Final Report - Apr. 17, 2017 Apr. 12, 2017 UE Final Presentation - Apr. 20, 2017 Apr. 17, 2017 PLAR Report Apr. 24, Apr. 21, P a g e

86 References 1. Center, G. C. (2016, 08 10) NASA's Student Launch. Retrieved 08 11, 2016, from NASA: 2. Niskanen, S. (2009). Development of an Open Source model rocket simulation software. OpenRocket. Helsinki: HELSINKI UNIVERSITY OF TECHNOLOGY. 3. Ring, C. (2016, 9 27). Launch Crue. Retrieved from LaunchCrue.org: 79 P a g e

87 Appendix A Machine Prints Dimensioned Drawings 80 P a g e

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