2015­2016 NASA USLI Flight Readiness Review (FRR)

Rensselaer Rocket Society (RRS)

Rensselaer Polytechnic Institute 1999 Burdett Avenue

Troy, NY 12180

Project Name: Red Gemini Task 3.1.1 – Atmospheric Measurements

Task 3.1.6 – Aerodynamic Analysis

14 March 2015

1. Table of Contents

Table of Contents 2. Executive Summary

2.1 Team Summary 2.2 Launch Vehicle Summary 2.3 Payload Summary

3. Changes Made Since Critical Design Review 3.1 Vehicle Changes 3.2 Payload Changes 3.3 Project Plan Changes

4. Vehicle Criteria 4.1 Design and Construction of Launch Vehicle

4.1.1 Design and Construction of Structural Elements 4.1.2 Design and Construction of Electrical Elements 4.1.3 Mission Reliability and Confidence 4.1.4 Test Data and Analysis 4.1.5 Workmanship 4.1.6 Safety and Failure Modes 4.1.7 Mass Report 4.1.8 Full­Scale Flight Results

4.2 Recovery Subsystem 4.2.1 Structural Elements 4.2.2 Avionics 4.2.3 Redundancy Features 4.2.4 Transmitters and Sensitivity 4.2.5 Safety and Failure Analysis

4.3 Mission Performance Predictions 4.3.1 Mission Performance Criteria 4.3.2 Flight Simulations and Predictions 4.3.3 Thoroughness and Validity of Analysis 4.3.4 Kinetic Energy Analysis 4.3.5 Drift Calculations

4.4 Launch Vehicle Verification 4.5 Safety and Environment (Vehicle) 4.6 Payload Integration

5. Payload Criteria 5.1 Experiment Concept 5.2 Science Value 5.3 Design of Payload Equipment

5.3.1 Drag Flap Subsystem

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5.3.2 Atmospheric Measurement Subsystem 5.4 Payload Verification 5.5 Safety and Environment (Payload)

6. Launch Operations Procedures 6.1 Checklist

6.1.1 Recovery Preparation 6.1.2 Motor Preparation 6.1.3 Launcher Setup 6.1.4 Igniter Installation 6.1.5 Troubleshooting 6.1.6 Post­Flight Inspection

7. Safety Plan 7.1 Safety Officer and General Safety Actions Hazard recognition and accident avoidance Pre­launch briefing Purchasing and handling of rocket motors Transportation of rocket to huntsville Safety Agreement 7.2 Preliminary Checklists 7.3 Failure Modes 7.4 Hazard Analysis and Environmental Concerns

8. Project Plan 8.1 Budget Plan *Note that the above numbers for income, spending, and difference are assuming that we reach our $3000 fundraising goal. 8.2 Funding Plan 8.3 Timeline 8.4 Educational Engagement 8.4.1 National Manufacturing Day Event

8.4.2 STEM Engagement with the Boys and Girls Club 8.4.3 STEM Engagement with Troy School 2

Appendix Appendix A: Scientific Payload Schematics Appendix B: Structural Design Assembly Drawings Appendix C: Code to Rotate Camera Images Appendix D: Motor Thrust Curve

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2. Executive Summary 2.1 Team Summary This report was produced by the Rensselaer Rocket Society (RRS) of Rensselaer Polytechnic Institute (RPI). The RRS operates out of the Ricketts Building at RPI. The mailing address for the RRS will be the same as that for RPI: 1999 Burdett Ave, Troy, NY 12180. The Community Mentor for the RRS is John Sicker, who is certified to Level 2 with the National Association of Rocketry (NAR) and Level 3 with the Tripoli Rocketry Association (TRA). His NAR member number is 49422, and his TRA member number is 01017. 2.2 Launch Vehicle Summary The launch vehicle design originated with the Liberty 4 from Giant Leap Rocketry, but has undergone significant changes since. The rocket will be approximately 97 inches in height and will have a body 4.02 inches in diameter. The rocket will have a mass of about 15.81 pounds mass. The launch vehicle will be propelled by an Aerotech K1103X motor for a 54mm motor mount. The recovery system will consist of a main and a drogue parachute to be deployed via electronic deployment. This deployment will be controlled by one of two altimeters to be used for redundancy: a StratoLogger SL100 altimeter and a Raven3 altimeter. The altimeters will only be connected to their independent power supplies and the deployment ejection charges. 2.3 Payload Summary The rocket payload, the Red Hawk, will be composed of two, physically separate systems. The Drag Flap Analysis Module will focus on accomplishing Task 3.1.6, “An Aerodynamic Analysis of Structural Protuberances.” The Environmental Analysis Module will focus on accomplishing Task 3.1.1, “Atmospheric Measurements.” The Drag Flap Analysis Module will chart the movement of the launch vehicle and compare the computer projected path of the rocket to the actual movement of the rocket. After the motor has completed its burn, the on­board computer will determine the angle at which the flaps should be deployed in order to apply an appropriate drag force to hit the target apogee. After deployment of the four drag flaps, barometric pressure sensors below the flaps will take readings of the pressure behind these structural protuberances and compare this data to readings forward of the drag flaps. The Environmental Analysis Module will take readings from several sensors to measure ambient conditions during the rocket’s descent. There will also be a gravity­guided camera system to take the required pictures during descent and after landing through a

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plexiglass section of body tube. The camera will be oriented correctly through a weight that allows the camera to rotate around the center axis.

3. Changes Made Since Critical Design Review 3.1 Vehicle Changes The vehicle design underwent minor changes and came to full maturity and fruition. Due to workability issues, the Lexan polycarbonate camera chamber was trimmed on the exterior to act as a coupler section itself, instead of trimming the interior to integrate standard couplers. This shaved 8” off of the rocket, reducing the length from 104” to just over 96”. Other final details of the assembly developed, such as the placement of structural screw and shear pin holes. Following a structural failure during a test launch, considerable changes will be made to the vehicle design, which will be outlined in section 4.1.8. The changes to be made to the vehicle will reflect the experience from the test launch and ensure that similar failure modes do not occur in the future. 3.2 Payload Changes

Several changes to the payload have been made to address concerns brought up in the CDR. The bottom of the 3D Printed Drag Flap system has been extended from 4 inches to 6 inches in order create a longer coupler section. The infill percentage of some of the 3D printed parts was lowered from 100% to 40% after more testing was done on the parts revealed a shear plane issue. It was found that prints with 100% infill more easily sheared along the plane of the print when subjected to large shear forces. By setting the infill to 40%, the shear stresses flowed through and around the support structure, as well as reduced the weight of the parts. The stepper motor used to drive the drag flap system was changed in order to employ a motor with a higher torque value. The drap flap control algorithm was also changed to a proportional algorithm as opposed to a PID algorithm in order to free up system resources on the microcontroller.

3D Printing Infill % Changed due shear plane issue Bottom of Drag Flap System has been extended as per recommendation made

during CDR New Stepper Motor to provide a higher torque Changed Flap Algorithm to proportional instead of PID to conserve processing

power 3.3 Project Plan Changes The project plan has been updated to reflect the construction and testing plan and actual events between the CDR and FRR. This includes significant construction and testing events, such as glassing of the fin can and attachment of recovery mounts. In

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addition, significant changes were made to the plan between the FRR and competition week after the events of the full­scale test launch.

4. Vehicle Criteria 4.1 Design and Construction of Launch Vehicle

4.1.1 Design and Construction of Structural Elements The construction of the vehicle systems aimed to closely follow the design of the project outlined in the PDR and CDR. As suggested by the designs in those reports, the vehicle construction was split between lower airframe and upper airframe assembly. For clarity, each of these sections will be described independently. Within each section, procedures will be provided in roughly chronological order. 4.1.1.1 Lower Airframe Construction

The majority of the effort in constructing the lower airframe went to creating the fin canister/motor mount piece, preparing and slotting the phenolic resin body tube, and final assembly. Upon arrival, parts were inspected for any sign of damage or wear from the manufacturer or shipping processes. After passing this inspection, work began on the construction of the motor tube. The 54 mm diameter phenolic resin tubing to be used for the motor tube came in as a 36” segment, which needed to be trimmed to a length of 20” to meet our design. Since phenolic resin is slightly brittle, special care had to be taken when cutting the tube so as to avoid cracking the tube or separating its layers and thus wasting material. To ensure cuts about the circumference were as accurate as possible, carefully measured and templated lines were drawn where the cut was to be made. Then an X­acto knife was used to carve an initial grove around the tube. Once deep enough, the grove was then slowly deepened using a utility knife. This process created a clean cut without breaking either end of the material. The 20” portion of the tube was then squared off using a belt sander. The tube was then marked at 1” from the bottom, 8.1675” from the bottom, and 1” from the top. These lines marked where the centering rings would be placed, leaving a 7” gap between the first and second rings for the fins. One at a time, the 3/16” thick plywood centering rings were tacked onto the motor tube using a thin layer of high­strength, aerospace­grade JB­weld. The rings were checked for alignment, then left to fully set. At this point, preparation of the lower airframe body tube began. The lower airframe required a 44” long section of 4” outer diameter (OD), 3.9” inner diameter (ID) phenolic resin tubing with 1/8th” fin slots extending 8.1675” from the bottom of the tube. Public Missiles Ltd (PML) had a 48” un­slotted tube available, which was used for this project. The fin slots needed to be as precise

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as possible to work in conjunction with the fin alignment guide (which will be described shortly). To this end, a weighty piece of paper was used to mark exactly one circumference on both ends of the paper. This length was then subdivided into four sections on both sides, so that perpendicular lines could be marked where each slot was to be made. The paper was then wrapped around the tube again, and an aluminum angle was used to mark one side of the slots as straight as possible. The paper guide was then shifted an eighth of an inch and the other side of each slot was marked. In order to cut perpendicular to the circumference, some sort of backing was needed inside the tube to ensure that the cut did not deform the tube in anyway. To accomplish this, a long wooden rod about an inch thick was clamped to a table, and the body tube rested on this rod. The same procedure as above was then followed, making an initial cut with an X­acto knife and then transitioning to a utility knife. The slots were intentionally narrower than needed to be so that edges could be sanded provide as precise of a fit as possible for the fins. Then the extra 4" were removed from the top side of the body tube in the same fashion as cutting the motor tube. When this piece was not in use, a piece of coupler section was inserted and a piece of paper wrapped tightly around the exterior to prevent any warping from occurring. The fins then needed to be prepared and tacked onto the motor tube. The fins for the rocket were custom ordered from PML, and fit all of the design specifications upon arrival. In preparation for tacking onto the motor tube, the edges about the leading and parallel sides of the fins were lightly rounded using 100 grit sand paper, the interior edge was sanded to create an adhesive surface, and the tang was roughed up to allow bonding during the application of fiberglass. In respect to the rounding of the leading edges, extreme care was taken to ensure that the fins were as identical as possible to avoid introducing any asymmetry that would promote spin during launch. Additionally, the fins needed to be attached as perfectly orthogonal to the motor mount as possible. To this end, we created a fin alignment guide out of machine cut plywood boards framed in wooden planks. In the center of each board, a 4” circle was cut with four equally spaced 1/8” slots to align the fins. Then two boards in separate frames were clamped down to a flat surface such that their slots were perfectly aligned. The motor mount then had a very thin layer of 30 minute set epoxy applied where each fin would be tacked. The motor mount was then placed inside the body tube at the slotted end, and each fin into a slot. The entire assembly was then inserted into the alignment guide and secured in place and left to cure for several hours. Then the airframe was removed from the guide, and the fin can with all four fins tacked onto the motor tube was removed from the body tube and placed on a support to avoid placing any weight on the fins. A simple tack using 30 minute epoxy would provide severely insufficient structural integrity of the lower airframe. The first reinforcement added was a

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thick fillet in each groove between the motor tube and the fin. Each of these eight fillets was made using the high­strength, aerospace­grade epoxy. Once all of the fillets were made and set, strips of fiberglass weave were cut to fit each of the four sections spanning from the top of one fin’s tang, across the motor tube, and to the top of the opposing tang. To ensure neatness and prevent epoxy from falling on the fins, tape was applied above each tang. Then a layer of West Systems 105 Epoxy Resin mixed with 206 Slow Hardener was applied to the area, and the strip of fiberglass carefully laid down. A thick bristled brush was then used to press down the fiberglass and ensure that it was completely coated in the epoxy. A full 12 hour set was allowed between each application of fiberglass. Once all four sections were successfully glassed, a dremel tool and sandpaper were used to clean up the edges of each application. Reinforcing fillets were then applied around each of the centering rings. Then small wooden blocks were secured above the top and bottom centering rings to be used as backings for the rail buttons. A forged eye­bolt was drilled into the top centering ring and secured with epoxy, and the motor retainer was secured to the end of the motor tube. Finally, the proper length of shock cord was measured and tied to the eye­bolt, and safely placed inside of the motor tube. With reinforcements for the fin can complete, the lower airframe was ready for assembly. To prepare the lower airframe for assembly, the interior of the body tube was sanded where the centering rings would rest. The fit of the fin can in the body tube was tested and the position of the rail buttons marked on the body tube. After verification of the fit, high­strength epoxy was applied to the centering rings and pre­sanded areas, and the fin canister was permanently fixed in the body tube and left to fully set. Then small fillets were applied along the edge between each fin and body tube section, while also sealing the back end of the fin slots. Once all eight fillets were set, the rail buttons were installed 2.5” and 18” from the bottom using careful alignment and a drill attachment that ensured a hole that was perfectly normal to the exterior of the rocket. Three shear pin holes for the separation point to the upper airframe were drilled 1.7” from the top of the tube, and placed symmetrically about the circumference of the rocket at that point.

4.1.1.2 Upper Airframe Construction One of the major challenges for the upper airframe revolved around the integration of the Lexan polycarbonate section. No suppliers were found that had polycarbonate tubing that matched the dimensions of the phenolic resin tubing sections. The closest available match was a 12” section of polycarbonate tubing with an ID of 3.75” and an OD of 4”. Due to the thicker walls, standard coupler sections would not be compatible without modification. A horizontal bandsaw was used to remove 3” of excess material from the length of the tube. The piece’s ends were squared with a belt sander. Initially, the plan was to bore

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the interior of the polycarbonate using a boring bar on a lathe with the piece supported by a steady rest. This method was infeasible due to size limitations of available steady rests. The design changed to turn the tube’s exterior to 3.9” for 4” on each end of tube, leaving a 1” band in the middle of polycarbonate for the camera’s viewing window. This allows the polycarbonate to couple the two sections of body tube together, while leaving the camera 1” of clear section to take pictures through. The polycarbonate was held by a three jaw chuck with a 3D printed plug inserted inside to prevent deformation caused by the pressure of the chuck jaws. A bull nose center was used to support the piece on the other end, to help prevent workpiece deformation during cutting. Since the polycarbonate chamber for the camera subsystem lay in the middle of the upper airframe, two sections of phenolic resin body tubing had to be cut. From a 36” section of tubing, two sections of 12” and 19” respectively were cut and squared off using the same procedure as above. Next, holes needed to be drilled in several locations for assembly and payload integration. On the 19” section, three shear pin holes were drilled into the body tube and nose cone simultaneously, 1.38” down from the top of the body tube, spaced symmetrically. Then the bottom of the 19” body tube section had two sets of three structural screw holes drilled, both symmetrically spaced. One set was placed 1.5” from the bottom of the tube, and the other 60 degrees offset and 3” from the bottom of the tube. These structural holes aligned with corresponding tapped holes in the polycarbonate section. The 12” section had two sets of three holes drilled in the top to line up with corresponding tapped holes in the polycarbonate section, symmetrically spaced and offset from one another by 1.5”, starting 1.5” from the top of the section of body tube.. As in the 19” section, the sets were offset from one another by 60 degrees. On the bottom of the 12” section two more sets of three holes were drilled, to line up with holes in the top of the payload shell where the drag flaps assembly interfaces with the rest of the main airframe. These holes were placed in a similar manner to the other holes on the body tubes, with the sets offset from each other along the length of the rocket and along the rocket’s axis by 60 degrees. In between these two sets of screw holes, six more symmetrically spaced holes were drilled; three for access ports to avionics screw switches and three for static pressure holes, all of which aligned with their corresponding holes in the drag flap shell.

4.1.2 Design and Construction of Electrical Elements The electrical components of the vehicle include the altimeters associated with the recovery system, the scientific payload (which itself includes a number of different sensors and processors), the control system for the drag flaps, and the processors for the camera system. Since each of these elements is a very complicated system in and of itself, this section will simply refer to other sections of the report in which each one of these electrical subsystems can be found.

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Table 1: In depth electronics sections

Recovery System Electronics 4.2.2

Scientific Payload 5.3.2

Drag Flap Subsystem 5.3.1

Camera Subsystem 5.3.2

4.1.3 Mission Reliability and Confidence In order to allow for adequate testing and use of the launch vehicle, it is important to design and construct the rocket with reliability engineering in mind. This means that emphasis is placed on engineering knowledge and construction techniques to reduce the chance of failure of any part of the rocket. In the case of failure of a part on the launch vehicle, reliability engineering also emphasizes a reduced cost of repair for that part, including mission­critical and redundant parts of the vehicle. Repair cost must be analyzed in terms of time and money. Reliability engineering is a similar unit with safety engineering, as it is important that failure of any part does not occur. One of the major steps that was taken in order to reduce the likelihood of structure failure was the use of basic RRS construction procedures and workmanship with relation to each of the steps of construction. These procedures are covered on in Section 4.5 in relation to ensuring adequate mitigation of safety failure modes. Main emphasis is placed in this procedure on construction of the “fin can”. As emphasized in section 4.1.1.1, the RRS uses a “fin can” to install the fins of the rocket to the motor tube and lower airframe. The procedure for installing these fins is to install the centering rings to the motor tube with high­strength and high temperature epoxy, tack the fins to the motor tube using a specially designed, carefully manufactured guide frame, secure the fins to the motor tube with a high­strength and high temperature epoxy, and then strengthen this attachment with a layer of fiberglass and high­strength West Systems epoxy mixture. The fin can is then inserted in the rear of the motor tube and solidified in the lower airframe with high­strength West Systems epoxy. Solid workmanship is emphasized at each of these steps. Another major step that is taken to stop at each step of the construction procedure or at the beginning of each build session to carefully plan what each member is going to do. This planning session emphasizes the individual steps that need to be taken to ensure that the goal of each step is accomplished. These goals usually center around ensuring the strength and stability of the parts and ensuring accurate placement of parts for each sub­structure. For example, when attaching the centering rings to the motor tube, lines are drawn on the motor tube to mirror the location of each centering ring. These lines

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serve as a reference for where the centering ring needs to be placed. The lines are carefully measured and marked all of the way around the tube. Then, a thick layer of tape is adhered just above one of the centering ring lines to assist as a stop when sliding the centering ring on the tube. After carefully placing the motor tube in a position to dry that will not interfere with the drying of the ring in place, this tape is removed. Reliability engineering also emphasizes the strength and durability of each part used in the structure. While each part was carefully analyzed during the design phase of this project for strength and potential failure, it is important to ensure that these parts are indeed as strong as the theoretical analysis stated it would be. When any part arrived at the RRS work area, it was inspected for damage or irregularities. During the design phase of this project, the RRS attempted to use quality, local dealers as much as possible. This allowed the RRS to directly contact these companies to track each purchase and ensure the quality of the product received. The RRS also purchased at least one extra version of every part in case one of the parts was damaged upon arrival or during the construction process. This idea was particularly important in this particular build with the centering rings. Upon arrival, each centering ring was inspected for damage and irregularities. It was discovered that one of the four centering rings ordered was extremely warped. The warp was discovered to be over a 0.125” difference between one side of the warp and the other. This centering ring was marked and separated from the others. Special care was taken to ensure that no further damage occurred that might compromise the integrity of the remaining three rings. Remaining inspections of parts showed no warping greater than 0.05”, no chips to material that were judged to be potentially compromising to the structural integrity of the vehicle, and no other notable irregularities. When designing the launch vehicle, parts were made to be easily accessible if they needed to be replaced for any reason. Any portion of the rocket that could not be easily accessed or repaired had special emphasis placed on its strength during construction. Large safety margins were used when building each critical structural member. Recovery system points are the most accessible and repairable parts of the launch vehicle, such as the parachutes and shock cord as well as the electronics. Despite the close attention to construction procedures that emphasize strength of all structural members of this rocket, some concerns still remain in regards to a successful test launch. The upper airframe contains payload components that serve a dual purpose as structural reinforcement, which paired with its placement far from the motor makes that section the least likely to fail. However, the overall length of the vehicle combined with the lack of internal reinforcement of the drogue chute bay make the lower airframe much more susceptible to failure, especially at the coupling point to the upper airframe. To mitigate this chance, the RRS has lengthened the 3D printed coupler section from 1 body width to about 1.3 body widths to reduce any looseness in this critical connection point. Though the phenolic resin material selected as the primary

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structural material possesses great strength as outlined in section 4.1.1, some research and testing has suggested that any failure of this material will most likely be critical. Due to the nature phenolic resin, any break will have a high chance of major fracturing, and small damage will greatly increase the chance of future failures. Thorough inspection of parts as they arrive and go through the construction process will reduce unnecessary risk associated with this material chance, though the risk cannot be completely removed. The initial test flight with the full­power motor shredded the rocket at the end of the motor burn. This was the point at which the rocket reached the maximum velocity, and the point in flight where failure was most likely to occur. This failure is in the process of being mitigated by the ongoing failure analysis of the launch and the measures discussed in depth later in the report on the construction fixes to the rocket. The test flight is discussed further in Section 4.1.8. 4.1.4 Test Data and Analysis The majority of testing for the vehicle can be split into simulation results, component testing, and powered testing. The simulations have been completed, and with detailed results found in Section 4.3. Component testing was included throughout the construction and assembly process. When subsystems such as the fin canister in the lower airframe are assembled, several evaluations will be performed to ensure proper alignment and appropriate structural reinforcement. Details of the full­scale vehicle launches and results can be found in Section 4.1.8. Additionally, the completed design will undergo several more test flight mirroring the full­power of the rocket as closely as possible. This further testing is to ensure that all systems are functional and secure in preparation for the competition launch. Component testing and substructure evaluations were conducted at each point in the construction of the launch vehicle. Inspection of construction materials was conducted as soon as each part arrived at the RRS work area. After materials were cut to size, the leftover material was used to confirm the strength of the part. After each substructure was completed, an evaluation was conducted to ensure that the substructure was adequately assembled. Table 2 describes the damage, warping, and other irregularities that were found in the initial inspections of the materials.

Table 2: Material Damages, Warping, and Irregularities

Part/Material Damage? Warping? Other Irregularities? Notes

Centering Rings 0/4 4/4 0/4

Only significant (> 0.05") warping on 1/4

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Motor Tube No No No

48" Airframe Yes No No Minor Chip at end of tube. That section cut off

36" Airframe No No No

Fins No No No

Nosecone No No No

Eye Bolts No No No

Main Parachute No No No

Drogue Parachute No No No

Shock Cord No No No

Bulkheads No No No

Motor Retainer No No No

Screws No No No

Shear Pins No No No

Lexan Tube No No No

Material strength confirmation testing was conducted on the motor tube material and airframe material. No material strength confirmation testing was able to occur on the fins, bulkheads, and centering rings due to the nature of the shear forces that will be acting on those members. Static load testing was conducted on a 4 inch section of airframe tubing and on a 3 inch section of motor tube. This testing revealed that the motor tube section failed at about 620 lbf, and the airframe tubing failed at about 650 lbf. This is a factor of safety of over 1.3 for each material. As each of these parts are also being heavily reinforced, this factor of safety is well within a reasonable level for this mission. Evaluations were also conducted on the fin can substructure, drag flaps structural members, and the lexan tube. Each of these substructures was inspected by the RRS members and club mentor to be extremely strong for its purpose in the rocket.

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4.1.5 Workmanship The workmanship of any construction project directly correlates to mission success. Workmanship includes the use of safe construction practices in order to produce a quality project using the best standard industry operating procedures With this in mind, the RRS has made it a key goal to maintain the highest levels of workmanship in the construction of its launch vehicle. In particular, it is critical for the RRS to maintain solid workmanship throughout the launch vehicle construction because of the high impulse motor that has been selected for use in this project. Some of the ways that the RRS will strive towards good workmanship include emphasizing the use of personal protective equipment (PPE), using epoxies and tools in the proper manner, and an emphasis on constructing the launch vehicle according to the predetermined design with quality construction methods based on the RRS mentor’s significant experience. The team’s use of workmanship with respect to PPE and use of epoxies and tools in a safe manner is discussed in depth in later sections. Some of the techniques that the RRS will use to facilitate quality construction and workmanship of the launch vehicle include proper epoxy use, a strengthened installation of the fins and motor tube, and ample testing of all critical components of the launch vehicle. Proper epoxy use consists of selecting the correct epoxy to bond two materials for their particular launch environment, sanding each surface to provide ample quality bonding surface, and vigorously mixing both parts of the epoxy in correct ratios for a enough time to ensure that adequate cementing of the epoxy can occur. Testing of each component depends on the use of that individual component. Some examples of this testing includes load testing of samples of body tube not to be used in construction, electronics testing of all circuits and altimeters to be used in the recovery system, and ensuring all components are not significantly warped or damaged based on visual inspection before their installation. The RRS will also rely on a common rocketry construction process for installing the fins and motor tube that has been used by the RRS numerous times and highly recommended by the club mentor. This technique involves epoxying the centering rings on the motor tube, then attaching the fins to this assembly with an accurately constructed fin alignment guide made with high­precision power tools from sheets of plywood. This “fin can” is then secured with additional epoxy or fiberglass sheets before insertion into the lower body tube with epoxy. The techniques used by the RRS and recommended by the club mentor will be followed closely in the construction of the launch vehicle with quality workmanship. 4.1.6 Safety and Failure Modes

The RRS analyzed the safety and failure modes of the launch vehicle to a high level. The majority of this analysis can be seen in Section 4.5. However, the RRS also analyzed the entire launch vehicle structure for safety and failure levels with a high level of engineering calculations. The RRS looked at the strength of the rocket with respect to

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stress­strain, material analysis, and buckling. These calculations and the relating analysis is shown below.

Considering the 4.0” diameter of the rocket and 0.1” thickness of the body tube at its thinnest level, the cross sectional area is 0.75 square inches. The maximum thrust that the rocket will experience is 364.2 lbf. This leads to a maximum axial stress on the rocket of 485.6 psi. Based on a conservative estimate for the body tube’s Young’s modulus of 300,000 psi and an assumption that Hooke’s Law applies to this situation, the maximum axial strain of the rocket could only reach ­0.002”. Since no stresses occur in the transverse direction, there is no natural strain in that direction.

In a report by Branko Sekulic of the Polytechnic University of Catalonia, the compressive yield strength of cardboard is about 13 MPa, or 1885.5 psi. When compared to the expected stress that the launch vehicle will face of 485.6 psi, there is a factor of safety over 3.8.

A complex buckling analysis was performed on the launch vehicle due to the uncommon boundary conditions. The launch vehicle was assumed to be approximately a beam of length 104 inches that was made of purely the Phenolic body tube. The analysis shown in Figure 1 is the derivation of the governing equation for the buckling of the rocket. This equation provides the amount of deformation in the transverse direction due to the buckling of the rocket. Initial solutions of this equation provide approximately zero deformation due to buckling. The RRS worked with an RPI professor to derive this equation and is currently working with this professor to solve the equation with more advanced methods to verify the initial conclusions.

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Figure 2: Buckling Analysis of Launch Vehicle

4.1.7 Mass Report The following tables give detailed mass contribution for each part in the upper and lower portions of the airframe.

Table 3: Mass of the Lower Airframe

Part Mass (oz) Quantity Subtotal (oz)

Assembled Airframe

60.8 1 60.8

Drogue Chute 6 1 6

Motor (Pre­Launch) 51.5 1 51.5

Subtotal ­ ­ 7.393 (lbm)

Table 4: Mass of the Main Airframe

Part Mass (oz) Quantity Subtotal (oz)

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Nosecone 10 1 10

Assembled Airframe

18.22 1 18.22

Scientific Payload 8.2 1 10.2

Drag Flap System 27.5 1 27.5

Shock Cord 6.72 1 6.72

Main Chute 13.3 1 13.3

Avionics 8.3 1 15.3

Camera System 4.2 1 10.2

Electronics Mounts

23.3 1 23.3

Subtotal ­ ­ 8.421 (lbm)

Total Mass: 15.81 lbm

As illustrated in tables 3 and 4, the mass of the rocket was obtained by careful measurement of the components of the rocket. Smaller parts such as payload subsystems were weighed with a scale accurate to a hundredth of an ounce, while larger sections such as the assembled lower airframe were weighed on a scale accurate to a tenth of a pound. The masses provided now include the majority of construction elements, such as bulkheads, epoxy, and other reinforcing elements. A large portion of the mass increase since the CDR is attributed to the heavy reinforcements made to the lower airframe around the fin canister. The process of glassing the motor tube involved a large sum of epoxy and fiberglass strips which was not included in previous mass statements.

4.1.8 Full­Scale Flight Results

The full­scale flight was launched at the Maryland Rocketry Club launch site at their third regularly scheduled launch. The launch was conducted as planned, with the launch vehicle design outlined so far in the project and the Aerotech K1103X motor identified as the full­power motor for the launch. The launch was set as a test flight for the absolute altitude ceiling of the rocket without the use of the drag flap system. The test would be

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modeled against base predictions for the maximum velocity, acceleration, and altitude of the launch vehicle, as well as testing of the structural and recovery systems of the rocket. The flight was modeled on the same Open Rocket software used in the subscale launch and all other planned launches to be conducted later in the testing phase of the project. This modeling can be found in Section 4.3.

The flight ended in mission failure. Near the end of the motor burn, the rocket shredded due to the extreme forces experienced at that critical point in the launch. Figures 3 and 4 show the moment of failure of the rocket, and the full video of the launch will soon be uploaded to the RRS website.

Figure 3: Moment of Launch Failure

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Figure 4: Moment Right After Launch Failure

All parts of the launch vehicle were successfully recovered. Photographs of these parts were taken, and will soon be available to see on the RRS website. The pieces of the launch vehicle, the video and pictures taken during the launch, and the recovered flight data was used in the mission failure analysis.

The launch vehicle was recovered in 3 main independent sections: the upper airframe, the drogue parachute and drogue parachute shock cord, and the lower airframe. The upper airframe contained the nosecone, the payload and avionics bay, the main parachute, and the main parachute shock cord. The lower airframe was recovered as the bottom 18 inches of the rocket, including the fin can, spent motor, and motor casing. The eyebolts connecting the drogue parachute to the lower airframe and payload bay were sheared off from each section, and all other airframe tubing was shattered. After recovery, each part of the launch vehicle was carefully analyzed under 2X magnification and bright lighting. No structural damage was found on the payload and avionics section, nosecone, shock cord, or either parachute and, aside from the top centering ring where the lower airframe split from the remainder of the launch vehicle, no damage was found on the fin canister. The break at the fin canister appears to be a result of the eye bolt

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being ripped from the top centering ring, followed by a brittle­ductile shearing of the airframe just above that section, Figure 5.

Figure 5: Failure at Top of Motor Canister

The break of the rocket just below the payload section was a brittle­ductile shearing of the airframe and a ductile shearing failure of the forged­steel eye­bolt just below the nut holding the bolt to the bulkhead attachment, Figure 6. The break of the rocket just above the payload section was a brittle failure, with no damage to the eye bolt connection or payload bay below it. The break at the nosecone was a jagged edge from a brittle failure, with two of the shear pins bent and pulled out of the connection hole and the third fully intact in the connection hole. All shear pins were still in place, but two of the shear pins for the upper separation point had been bent slightly. The upper electric match was still intact, excluding the black powder charge, but the lower electric match had been ripped from its terminal block, leaving only small copper leads in the terminal blocks.

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Figure 6: Failure at Bottom of Payload Section

The screen­captures of individual frames of the video just before the moment of failure indicate a bending of the rocket before it snapped at the midsection, Figure 7. This bending and snap occurred while the motor was still in the process of burning, Figure 3.

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Figure 7: Rocket Bending Before Failure

Due to the minimal damage to the payload and avionics portion of the upper airframe, both the primary recovery Stratologger and secondary Raven3 altimeters were successfully recovered. With the Raven3’s built in USB interface, the RRS was able to recover data that helped to solidify potential modes of failure. The data is shown in Figure 8.

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Figure 8: Initial Raven3 Data with Axial Acceleration (Red Curve) and Barometric Altitude (Blue Curve)

As clearly visible in Figure 8, the failure occurred at 1.78 seconds into the flight. Accounting for the 0.18 second delay at the beginning of data collection, the failure occurred exactly at the end of motor burnout; 1.6 seconds after ignition. After this point, both the barometric sensor and accelerometer experienced extreme forces which caused large oscillatory spikes in data collection. Though it is possible that the force of the failure measured upwards of 64 g’s of acceleration, it is not possible that a spike up to 2000 feet AGL occurred at this point. Under the assumption that maximum height occurred shortly after failure event, it is safe to assume that apogee just above 900 feet AGL. Several seconds later, the Stratologger successfully detected that the altitude above ground level was below the preprogrammed main deployment height of exactly 900 feet.

From this evidence, the RRS determined that the potential modes of failure were: premature firing of the ejection charges due to electrical failure of the recovery system, brittle failure of the phenolic tubing from pure compression loading, gaseous emissions from the motor leaking to the lower airframe section to burst the airframe section, or buckling of the launch vehicle at the coupling section to the payload bay. Electrical failure of the ejection charge system can be eliminated from the list of potential failure modes because one of the electric matches had been completely ripped from its terminal blocks, while the other was intact, excluding the actual black powder charge. Pure compression loading failure can be rejected because the phenolic tubing section between the Lexan coupler and the 3D printed coupler in the payload bay showed no signs of being close to failure, while the two other sections failed in close succession during the launch. The failure mode from gaseous emissions from the motor bursting the lower airframe can be eliminated because no burn marks were observed in the motor tube or airframe sections above the motor casing, and the motor casing is perfectly intact. Failure of the launch vehicle from buckling at the coupling between the payload bay and lower airframe is the most likely cause of mission failure.

Failure of the launch vehicle due to buckling at the coupling between the payload bay and lower airframe is confirmed by all forms of evidence collected. Individual frames can be identified that clearly show extreme bending of the rocket around, and immediately following these frames, the rocket is seen splitting in half at that location on the rocket, Figures 2 and 3. This evidence suggests a buckling failure. Based on these images, the initial break occurred somewhere on the lower airframe. From the pieces recovered, this break most likely occurred at the connection of the phenolic tubing to the payload bay, as the phenolic tubing split suggests a shear failure in that location, and the eye bolt at that location sheared in one, neat plane from the bulkhead. Buckling theory also supports this conclusion. Buckling theoretical analysis conducted before the launch confirmed that a pure beam of phenolic tubing similar to this rocket would not fail. However, the assumption that this rocket acts as a beam would not hold at a loose

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coupling section. A strong, tight­fitting coupler would allow this assumption to hold, but a loose coupling section would introduce a weak constraint on the beam at that location. The coupling section at the location of the failure was slightly loose at launch due to tolerance limitations of the 3D printer, however, it was critically judged to be close to effective tightness. However, the break must have been violent, as the airframe below that section shattered and pulled the eye bolt out of the top centering ring, as well, Figure 9.

Figure 9: Shattered Centering Ring Section with Eye Bolt

The upper airframe failed because of a compounding series of events. The initial failure at the lower airframe buckling occurred towards the end of the burn. The separation of the upper airframe at this point caused a large amount of forces to act on this section. The flight data confirms that 64 g's of acceleration acted on this section of the rocket after initial failure. This causes a huge amount of stress on the rocket airframe. In addition, after the shredding of the rocket, the acceleration vector and low altitude of the rocket would cause the altimeters to fire both recovery events. The black powder charge of the rocket would compound with the effects of the large magnitude of the acceleration on the upper section of the launch vehicle to cause an additional break in the rocket. This was the second break of the rocket because less damage occurred; the eye bolts in this section are intact, compared to sheared eye bolts in the lower airframe. Physical evidence also indicates that two of the shear pins were bent and pulled out of their connection holes, while the third was intact. This section most likely failed in the tubing because the stress in the tubing from the initial failure in the launch vehicle caused the the tubing to be significantly weaker than the shear pins. In addition, the shear pins did not have a proper shear plane to break over. With the bending in the tubing from buckling, the shear forces on the shear pins were distributed across a larger distance, causing bending rather than a clean shear break, as designed.

In an optimistic light, this failure was a fortunate one in that it showed serious design flaws and in that it left the most expensive and labor intensive portions of the rocket undamaged and salvageable. Moving forward from this failure, the RRS will repair and improve the launch vehicle with significant design changes to counteract the failures seen at this launch. With the exception of the salvaged portion of the lower airframe, all

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sections of phenolic resin body tubing will be replaced or modified with fiberglassed phenolic resin. Further discussion will decide whether the RRS will glass the tubes, or if excess funds will be used to purchase pre­glassed tubing from Public Missiles Ltd. Each approach has advantages and disadvantages, and as such this decision requires careful consideration. The lowermost 18 inches of the lower airframe will be salvaged by cutting off the section of the tube that fractured, including the destroyed forward centering ring. The top 25 inches of the lower airframe will be replaced and permanently fixed through screws and other structural reinforcement to the current separation point below the 3D printed assembly. The new separating point for the drogue chute will then be at the bottom of the airframe. To ensure that this segmentation point has absolutely no slack to allow for the same type of movement that lead to the test launch failure, a Locktite coupler will be used. To mitigate the secondary failure of the shear pins on the nose cone separation point, the main chute bay will be lengthened slightly, and the nose cone will be tightly secured in such a way as to not disturb their shear plane. Unfortunately, making these adjustments, specifically switching to glassed phenolic resin tubing will add considerable mass to the rocket. Due to this, if the launch vehicle is to have a chance at reaching the target apogee the RRS will have to investigate a motor with much higher total impulse.

4.2 Recovery Subsystem

4.2.1 Structural Elements 4.2.1.1 Parachutes The parachutes chosen for the dual­deploy recovery system are the SkyAngle Classic II 52” main parachute from B2 Rocketry and the Ballistic Mach II 2 ft drogue parachute from Rocketman. The main parachute is constructed of zero­porosity 1.9 oz. balloon cloth with ⅜” tubular nylon suspension lines sewn around the canopy, rated for 950 lbf. The suspension lines meet at a heavy duty 12/0 nickel­plated swivel joint rated for 1500 lbf. Data gathered from OpenRocket simulations indicates that the maximum force on the main parachute and associated hardware will be approximately 90 lbf. This equates to the parachute and suspension lines having a factor of safety of 10. The factor of safety is excessive, but to achieve the desired descent rates and kinetic energy, a parachute of this size must be chosen. The drogue parachute is constructed from ballistic­grade rip­stop nylon. Calculations estimate that the maximum force on the parachute and associated hardware will be 13.75 lbf. The drogue parachute was chosen for it’s size and manufacturer reputation. The strength of the parachute is not a concern because of RRS's experience flying Rocketman parachutes and their trusted quality.

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4.2.1.2 Harnesses and Attachment Hardware The shock cord that tethers the parachutes to the airframe will be ½” tubular nylon cord. The cord for the main parachute will be connected to a 1” forged steel eyebolt. The drogue parachute will be tethered to the body in a similar manner, connecting to a 1” forged steel eyebolt on the aft bulkhead on the main airframe, and to a 1” forged steel eyebolt on the forwardmost centering ring of the motor housing. For each parachute there will be 300 inches of shock cord to allow for adequate separation of vehicle components during descent. The ½” tubular nylon cord is rated to 1000 lbf. The forged steel eyebolts are rated for 600 lbf. Using the above estimate of the maximum force experienced on the hardware of 90 lbf, there is no concern of attachment hardware failure during parachute deployment. 4.2.1.3 Bulkheads The eyebolts that serve as the parachute’s attachment hardware will be mounted to bulkheads at both ends of the main airframe. The bulkheads will be made of ¼” birch plywood. A schematic of each bulkhead can be seen below in figure 10.

Figure 10: Forward (left) and aft (right) bulkheads (not to scale)

At the top of the main airframe will be the forward bulkhead, which contains the eyebolt for the main parachute shock cord tether and the blast cups the parachute deployment. A drawing showing the arrangement is below in figure 11.

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Figure 11: Bulkhead hardware arrangement

Using the above estimate of 90 lbf in the eyebolt, a strength analysis of the bulkhead can take place. The load will be transferred to the bulkhead from the nuts that secure the eyebolt in place, creating bearing stress on the bulkhead. Using the area of the nuts, the stress is calculated to be 3.80 MPa. Applying a factor of safety of 1.7, the NASA factor of safety for critical components, the stress becomes 6.47 MPa. A simplified finite element model was created to analyze the bulkhead, with a uniform pressure of magnitude 6.47 MPa applied to the bulkhead over the area simulating the nut’s bearing load. The boundary conditions in the model reflect how the bulkhead will be supported in the rocket; fully constrained on the outer edge, and with fixed rotation, and fixed X and Z displacement in the central hole. Figure 12 illustrates the stresses in the bulkhead with the load applied.

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Figure 12: Von Mises stresses in forward bulkhead

The finite element analysis (FEA) indicates that the maximum von Mises stress in the bulkhead is 31.9 MPa. The ultimate stress of plywood can be difficult explicitly measure, and is nearly impossible to find standards for due to the nature of being a composite and the manufacturing processes used to create the composite will greatly impact quality. Typical values for birch plywood range between 31.0 MPa and 41.0 MPa however. Assuming an ultimate strength of 31.0 MPa, the design’s factor of safety is slightly less than 1.7, but is still a safe design. The second bulkhead is the one that houses the eyebolt for the drogue parachute. It is located at the lowest point of the main airframe. When the drogue parachute deploys, there will be a load applied to the eyebolt just as there is for the forward bulkhead when the main parachute deploys. However, from above, the maximum load that the aft bulkhead will experience from the drogue parachute is 13.75 lbf. Due to the fact that the load is considerably less than the load on the forward bulkhead, a strength analysis of the aft bulkhead is not necessary. Being made from the same material, in a nearly identical arrangement with a considerably smaller load indicates that the design is as safe or safer than the forward bulkhead. 4.2.1.4 Ejection Custom charge holders created from disposable gloves are placed on both the forward and the aft bulkheads of the main airframe. These charge holders are filled with black powder and ignited by electric matches connected to the altimeters. Approximately 0.6 g of black powder is required for the main parachute deployment, and 1.1 g of black powder for the drogue parachute deployment. At each ejection event, the altimeter’s

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connection to the electric matches becomes active and the current ignites the electric­matches, igniting the black powder charges. The charges separate the airframe and allow the unfurling of the parachutes. Both parachutes are protected from the heat effects of the charge firings with kevlar parachute protectors. Drogue parachute deployment occurs at apogee and main parachute deployment occurs at 800 ft AGL. 4.2.2 Avionics 4.2.2.1 Flight Computers The main electrical components of the recovery system consist of a Perfectflite Stratologger SL100 (Stratologger) barometric altimeter and a Featherweight Raven3 (Raven) barometric and accelerometer altimeter. The Stratologger acts as the primary recovery altimeter, set to deploy the the drogue parachute at apogee and the main parachute at 800 ft AGL during descent. The Raven acts as a the backup altimeter, and is set to deploy the drogue parachute at apogee and the main parachute at 700 ft AGL during descent. A 100 ft difference in the firing of the ejection charges for the main parachute ensures that if both altimeters have worked correctly, the ejection charges would fire into an empty chamber, posing no risk to structural failure; if both ejection charges fired at the same time there is a risk of overpressurizing the body and causing a structural failure of the body tube. Two altimeters provide system redundancy in the event that one flight computer fails. Table 5 below details each altimeter’s technical specifications.

Table 5: Hardware Specifications

Operating Voltage

Dimensions Sample Rate Altitude Accuracy Maximum Altitude

Stratologger SF100

4­16 V 0.90” W 2.75” L 0.50” T

20 Hz +/­ 0.1% 100,000 ft

Raven3 3.8­16 V 0.80” W 1.80” L 0.55” T

20 Hz +/­ 0.3% 100,000 ft

The Stratologger measures the rocket’s altitude by sampling the atmospheric pressure and comparing the readings with the ground level pressure. After launch, data is sampled at a frequency of 20 Hz. Altitude readings are stored in nonvolatile onboard memory and can be downloaded to a computer after launch using a standard FTDI interface. The Stratologger has two channels for parachute deployment, one for the main and one for the drogue. Main parachute deployment is adjustable from 100 ft to 9,999 ft in one foot increments. Figure 13 shows the Stratologger altimeter.

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Figure 13: The Perfectflite Stratologger SL100

The Raven measures the rocket’s altitude by sampling the atmospheric pressure, the same as the Stratologger, but also by measuring the acceleration of the rocket using an onboard accelerometer. After launch, data is sampled at a frequency of 20 Hz. Flight data is stored in nonvolatile onboard memory for download after launch using a USB interface. The Raven has four channels for parachute deployment, one for the main, one for the drogue, and two additional channels that will not be used. All channels are adjustable in one foot increments. Figure 14 shows the Raven altimeter

Figure 14: The Featherweight Raven3

4.2.2.2 Mounting and Electrical Layout The altimeters and accompanying hardware will be mounted to a two sided sled in the avionics bay. The altimeters will be mounted to one side of the sled, and the 9 volt batteries that power the altimeters will be on the other side, securely held in place by 9 volt battery holders.. The bay will have static pressure ports sized so the pressure in the bay is equivalent to the atmospheric pressure, allowing accurate altitude readings from the barometric sensors of the altimeters. The avionics bay will be hermetically sealed from the rest of the rocket by the tightening of bulkheads onto rings of silicon rubber acting on each side of the bay. The wires that pass through the bulkheads will do so through terminal blocks mounted on each side of the bulkhead with connections running through the bulkhead. Any gaps in the bulkhead will then be filled with epoxy to ensure a seal. An exploded view of the avionics section with tracelines indicating how the components come together can be seen below in Figure 14. Figure 15 is a block diagram detailing the recovery system circuitry.

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Figure 14: Exploded View of avionics assembly

Figure 15: Recovery System Electronics Schematic

4.2.2.3 Arming The recovery system electronics are armed through external, independent switches. Two screw switches externally accessible on the body tube allow each altimeter to be armed on the pad before launch, the prevent premature firing of ejection charges while the rocket is being handled and set up on the launch pad.

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4.2.3 Redundancy Features The recovery system features several redundancy features related to the altimeters and the ejection charges. The first system level redundancy feature is the use of two independent dual deploy altimeters. The altimeters share no electrical connections and are powered and armed by their own independent batteries and screw switches. The independent batteries minimize the risk of power loss to both altimeters; in the event that one battery becomes disconnected only one altimeter will lose power. In addition to the redundancy accomplished by having two batteries, having two altimeters provides additional redundancy, in the event that one altimeter loses power, or malfunctions and doesn’t trigger the ejection charges to deploy, there will be a second altimeter to fire the ejection charges and have the rocket successfully separate. The last redundancy feature is the use of two electric matches at each separation section, one for each altimeter. This provides redundancy in the case that one of the electric matches malfunctions and fails to ignite the black powder, there will be a backup.

4.2.4 Transmitters and Sensitivity The scientific payload portion of the payload subsystem consists of one transmitter, a commercially available Xbee Pro 900HP. The Xbee transmits data at a frequency configurable between 902 and 928 MHz. This frequency will be configured between the payload and the ground station prior to launch. The Xbee transmits with power configurable up to 250 mW. Experience with this Xbee module indicates that a 5 to 6 mile operating range is expectable. During descent and after landing the Xbee will transmit the GPS data from the on board GPS module to the ground station for recovery. The transmitters on the rocket pose no concern to the functionality of the recovery system electronics. A back of the envelop calculation involving the inverse square law and the transmitter’s power shows that it is not nearly enough to affect the pressure transducers inside of the altimeters. 4.2.5 Safety and Failure Analysis A table listing the failure modes, causes, effects, and risk mitigations for the recovery system can be found in the complete safety and failure analysis table listed in section 7.The table shows both pre and post mitigation RAC ranks.

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4.3 Mission Performance Predictions 4.3.1 Mission Performance Criteria The rocket will launch with a course that would lead to an apogee between 5,280 ft and 5,600 ft if left unaltered. This requirement will be successful if the vehicle travels between 5,000 ft and 5,600 ft, leaving room for error resulting from error in the payload. To facilitate the measurement and reading of this requirement, a barometric altimeter with the audible beep delivery system will be used for scoring purposes.

The launch vehicle will be reusable with fewer than four (4) independent, separable sections. These requirements will be successful if the rocket is able to be prepared for a re­launch immediately after landing, and if the rocket design includes fewer than four independent sections.

The launch vehicle will be powered by a single stage. The single stage will 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) and Tripoli Rocketry Association (TRA). The motor will be less than an L­class motor and will be launched by a standard 12 volt direct current (DC) firing system. This requirement will be met by the design parameters of the rocket if it is powered by one stage, if the motor used is approved by the NAR and TRA, if the motor is less than an L­class motor, and if the motor is capable of being launched on a standard 12­volt DC firing system. The recovery system of the launch vehicle will be electronic dual­deploy with a drogue parachute deployed at apogee and a main parachute deployed at a much lower height afterwards. The stages will be held together by removable shear pins, and, at landing, each independent section of the launch vehicle will have a kinetic energy of less than 75 ft • lbf. The requirements will be met if the launch vehicle successfully deploys its drogue parachute at apogee and the main parachute much later, after breaking the shear pins that hold each parachute in place. The kinetic energy requirement will be met in the design parameters of the parachutes and if the launch vehicle is able to sustain the landing forces associated with its kinetic energy at landing. The recovery system electrical circuits will consist of redundant altimeters that are physically and electronically separate from any payload electronics and power supply. Each altimeter will have a designated power supply and arming switch. These requirements will be met by the launch vehicle design.

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4.3.2 Flight Simulations and Predictions Using OpenRocket with a motor configuration for the Aerotech K1103X, a simulated model of the rocket was produced as shown in Figure 15.

Figure 15: Simulated Model for Flight Simulation

On this model, there is a stability margin of 5.15 cal with the motor included, and 7.91 cal with no motor. The Center of Gravity (Blue dot in Figure 15) is approximately 46.8” from the top of the nosecone, and the Center of Pressure (Red dot in Figure 15) is approximately 67.4” from the top. These placements have a variance of ±0.3”. Using this model and the selected motor, a flight simulation was created which is shown in Figure 16.

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Figure 16: Flight Simulation from OpenRocket

The simulated flight data in Figure 16 does not include the effect of the drag flap system. With this uncorrected flight, projected apogee is 5659 feet in 17.6 seconds, with a maximum acceleration of 702 feet/second2, maximum vertical velocity of 781 feet/second, and a total flight time of 138 seconds. The velocity of the vehicle off an 8 foot tall rail is exactly 100 feet/second 4.3.3 Thoroughness and Validity of Analysis The OpenRocket analysis is a starting point from which to gain an understanding of the order of magnitude of quantities such as altitude, velocity, and acceleration. As evidenced by the subscale launch, however, the results from OpenRocket are not always perfect.The goal of the payload section to is abate any discrepancies in the OpenRocket simulation by utilizing a control system which takes in real time data. Consequently, a certain amount of error in the analysis is permissible.

4.3.4 Kinetic Energy Analysis At the deployment of the main parachute, the main airframe will have a kinetic energy of 389.7 ft • lbf. The lower airframe will have a kinetic energy of 253.0 ft • lbf. When the rocket touches down, the main airframe will have a kinetic energy of 68.1 ft • lbf, and the lower airframe will have a kinetic energy of 44.3 ft • lbf. All of these values were determined by setting the gravitational force on the rocket equal to the drag force on the parachutes, which allows us to solve for the descent velocity of the rocket. In combination with the mass of the rocket this allows us to compute the kinetic energy of each section. 4.3.5 Drift Calculations Table 7 lists the drift distances with respect to the wind speeds for launch. The calculations were made assuming a vertical launch.

Table 7: Drift Analysis at Various Wind Speeds and Vertical Launch

Wind Speed Drift Value

0 mph 0 ft

5 mph 592 ft

10 mph 1158 ft

15 mph 1737 ft

20 mph 2317 ft

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The calculations show that for any launch below 20 mph, the rocket will land safely in the one square mile field to be easily recovered. These numbers were calculated using a simple python program that approximates the position of the rocket after ascent, and then uses the wind speed and decent rates to calculate drift during descent. 4.4 Launch Vehicle Verification

Table 8: Verification of System Level Requirements

Requirement Design Feature

Verification that Launch Vehicle Meets Requirements

Apogee between 5,280 ft and 5,600 ft if left unaltered

Rocket Mass, Rocket Motor, Design

Updated Unaltered Projected Apogee just over 5,600 ft. DOES NOT MEET REQUIREMENTS.

Reusable

Body Strength (fins,

airframe, parachutes,

etc.)

A strength, materials, and buckling analysis of vulnerable components was conducted to determine the rocket strength with respect to launch and landing forces. These analyses used a factor of safety of at least 1.4

in critical components During construction, the use high­strength epoxies has been followed. Components inspected upon order arrival show no signs of damage or

structural integrity issues.

Four or fewer Independent Sections

General Design

Rocket was constructed in three independent sections based on the original design (nosecone section, payload/avionics section, lower body)

Single Stage

General Design, Motor Design

Rocket designed with only a single stage motor. Rocket was constructed with a design that follows this restriction.

Commercially available solid

motor propulsion

system under Class L

Motor

OpenRocket Simulations based on a given 6 lbm payload mass with the selected rocket design allow for a motor well under Class L motor

approved by the TRA and NAR An Aerotech K1103X motor was selected.

Capable of Launch by 12 V

DC firing system

Motor, motor retainer

An Aeropack motor retainer that allows for access to motor was selected, and the selected motor is fully capable of being launched with 12 V DC

firing system.

Electronic Dual Deploy

General Recovery Design

(parachutes, shock cord, ejection

The general design is dual deployment, and it uses a drogue and main parachute with design for multiple separation points.

Initial calculations and simulations are run with these design parameters. The final design launch uses this recovery system.

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charges, etc.)

Drogue Deploys at

Apogee, Main Parachute Deploys at much lower altitude

Parachutes, ejection charges, General Rocket Design, Altimeters

The general design has the drogue parachute at first separation point and the main parachute at another. OpenRocket simulations were run with

these deployment locations. The altimeters used in the final construction were programmed to follow

these requirements.

Shear Pins hold rocket sections together until Parachute Deployment

Shear Pins, Ejection Charges

In the overall launch vehicle design, the shear pin strength was accurately calculated and inserted into the recovery design. The ejection charge

strengths were also accurately calculated for the full scale rocket design.

Independent Sections have less than 75 ft • lbf of KE at Landing

General Rocket Design (Mass),

Parachutes

The kinetic energy of each independent section was analyzed accurately. A major focus of the final construction has been to stay as close as possible to original mass estimates for the full­scale rocket build.

Parachutes and Recovery system components show no signs of damage upon order arrival.

Redundant, Safe Altimeters

Altimeters, Supporting Recovery Electronics

The full­scale launch vehicle uses 2 commercially available altimeters, the Raven3 and Stratologger CF altimeters. The electronics design has an

independent power supply to each altimeter. These altimeters are being stored safely, along with all other supporting

electronics components. During the final construction phase or the project, the team stayed as close to the original plans to meet the project requirements. The only requirement that is not met is the “unaltered altitude requirement”. Due to changes in payload mass, the projected altitude rose to 5659 feet, which is 59 feet about the maximum unaltered altitude. Every other requirement was met by the team, however.

4.5 Safety and Environment (Vehicle) 4.5.1 Safety Officer See Section 7 4.5.2 Preliminary Checklists See Section 7 4.5.3 Failure Modes See Section 7 4.5.4 Hazard Analysis and Environmental Concerns See Section 7

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4.6 Payload Integration 4.6.1 General Description The payload will be integrated into the overall vehicle using several compartments in the upper airframe, separated by bulkheads. This will consist sections for drag flap electronics, avionic devices, scientific instruments, and the camera subsystem. The drag flap module features threaded rods throughout the assembly along outside of the interior for structural integrity which will also be used to secure bulkheads and electronics. The lower coupler section of the 3D printed assembly will contain the electronics mounted to sleds along these rods, with the stepper motor will be secured to an upper bulk plate. A lower bulkhead on this section will protect the drag flap electronics and provide an anchor point for the drogue chute shock cord. Shear pins conjoining the couple section to the lower airframe will ensure that the flaps can clear the lower airframe and properly deploy during launch. Forward of the drag flaps themselves, another section will be created for the avionics bay, where the altimeters and associated electronics will be mounted on sleds, similarly to the drag flap electronics. Scientific instruments and electronics will be mounted on sleds above the avionics bay. The camera subsystem will be another isolated section, with slip rings allowing wires to pass through without concern for the rotating parts. The upper bulkhead above the camera subsystem will be air tight and serve as the attachment point for the main chute shock cord. The uppermost and lowermost bulkheads will have shock cord and parachute attachment points in the form of forged eyebolts and quick links to facilitate easy access to the payload and electronics, as well as custom charge holders to hold the black powder of the ejection charge. There will also be connection points for the electric matches that will trigger these ejection charges. Electronics arming and disarming will take place using screw switches accessible from outside the rocket. Shear pins will connect the upper airframe bay to the lower airframe and the nose cone to the upper airframe. These will sever at the first charge event to deploy the drogue, and at the second charge event to deploy the main chute, respectively. A more detailed discussion of the components of the recovery subsystem can be found in section 4.2. Data from the electronics section will be transmitted back to the ground station via an antenna. Actual physical modifications to the payload will require removal of the electronics sled as only the arming switches allow an external action to influence the payload. 4.6.2 Element Compatibility and Integration The vehicle’s payload can be split into two distinct categories; the drag flap subsystem, and the scientific payload. The drag flap subsystem acts as a structural member for the rocket, serving as the lowest portion of the main airframe. The drag flap subsystem is fully compatible with the vehicle because they were each designed with integration in mind. The top of the drag flaps shell slides into the 12” section of the main airframe and is secured by six 6­32 machine screws with clearance holes drilled into the phenolic resin, and tapped holes in the printed shell of the drag flaps shell.

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The scientific payload is integrated into the vehicle by sliding on the two threaded rods that run through the main airframe. The roads are a structural member a structural member of the vehicle that help to transfer stress through the body of the rocket in addition to the stress carried by the body tube, but also act as a convenient way to mount circuit boards to. The scientific payload consists of two circuit boards that are mounted back to back on the threaded rods. They are secured by nuts and washers on either end of the board, preventing translation and rotation in and about all axes. A Figure 17 is shows an exploded view of the main airframe including tracelines indicating assembled positions of each component.

Figure 17: Payload integration

4.6.3 Payload Housing Integrity All 3D printed parts are printed at a 40% infill with 3 shells/perimeters. Since the cross sectional area per print layer is fairly small, the selected infill percentage coupled with the 3 printed perimeters provides a highly durable housing for the drag flap housing. The infill pattern for the 3D printed sections was chosen to be a "honeycomb" pattern because it provides a great amount of rigidity and strength while keeping the amount of material needed low. It also lowers the total print time per part, because the honeycomb pattern is a pattern that the 3D printer can print relatively fast compared to other infill patterns. The other 3D printed parts in the rocket, mainly the different payload sleds and battery holders were printed with 3 perimeters also, but at 20% infill. Since these parts are mounted on threaded rods which will take most of the force produced during launch, and do not act as critical structural elements, they do not need to be printed at such a high infill percentage.

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Figure 18: Slicer View of 3D Printed Payload Pieces

5. Payload Criteria 5.1 Experiment Concept The experimental concept of the payload is that of a modular, easy to assemble design which can accomplish multiple payload requirements. The table below shows the requirements being fulfilled, and which elements of the payload fulfill each requirement.

Table 14: List of payload requirements

Payload Requirement Payload element(s) which fulfills requirement

3.2.1, Gather data for studying the atmosphere during and after descent

Avionics Bay, Camera Subsystem

3.2.6, Aerodynamic analysis of structural protuberances

Drag Flap subsystem

1.1, Vehicle shall deliver a payload to an apogee altitude of 5280 feet AGL

Drag Flap subsystem

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Though each of these elements will be explored in further detail, the concepts behind the functioning of the payload are as follows. The Avionics bay is a sealed section of the payload which measures a variety of atmospheric data as well as tracks the altitude, acceleration, and position of the vehicle in flight. The camera subsystem is a camera which can take correctly oriented pictures of the atmosphere outside the vehicle body during flight and after landing. The drag flap subsystem is a system of mechanically deployed flaps which slow the rocket down in order to reach the required apogee height.

5.2 Science Value There are a number of objectives that were established for this project. Many of these objectives are related to data collection. Data will be collected during this launch from a camera, mounted within the rocket, aerodynamic sensors, located beneath the drag flaps, sensors for measuring humidity, temperature, barometric pressure, solar irradiance, UV light levels, a GPS module, a wireless transmitter capable of sending data to the ground team, and an accelerometer to determine payload orientation during flight and recovery. The camera is going to be mounted and oriented in such a fashion that it is able to include the horizon and the ground within the context of the same photo. This camera will be able to store its data within an SD card that is also mounted on the rocket for extraction after recovery. The aerodynamic sensors will be able to analyze the turbulent flow behind the drag flaps to determine the feasibility of the design for accurate adjustment of the apogee of a rocket mid­flight. It will also be used to collect data so that an on­board microprocessor can determine the drag coefficient of the protuberances. The other sensors will be set up to simultaneously record data on the SD card while transmitting the data to the ground team. The GPS module and the altimeter will be used to determine the height of the rocket. This, in turn, will be used to determine the control of the drag flaps to adjust the apogee. It will also send data to the ground team via the transmitter. The payload success criteria for the payload is that the payload can successfully transmit at least two images during the descent of the rocket, and at least three images after landing (where the images have the horizon in frame with the sky at the top of the frame and the ground in the bottom of the frame), that the payload can successfully gather the required sensor data (pressure, temperature, relative humidity, solar irradiance, and UV radiation) every 60 seconds after landing and transmit/store the data for up to 10 minutes after landing, and finally, that the effects of the structural protuberances (the drag flap system) can be accurately measured and transmitted back to the ground station. For most of the scientific payload, sensors will be passively collecting and transmitting data, so the logic, approach, and method of investigation are fairly straight­forward. The drag flap system however, involves actively measuring pressure in front of and underneath the flaps in order to analyze the suction effect.

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According to several early CFD models that have been tested, there should be an increase in pressure underneath the flaps that increased as the flaps are deployed. In particular, the Fluent software that was run since the PDR has shown that the greatest pressure differential can be found at the outermost radial edge of the flaps. In order to investigate this, barometers can be placed beneath the flaps to measure the increase in pressure and barometers placed in front of the drag flaps. A comparison can be made between the readings from both data sets in order to see if there is truly a suction effect being created by the drag flaps. In this test, the angle in which the drag flaps are deployed at is one of the largest variables, since according to the CFD tests, the larger the angle they are deployed at, the greater the suction force. For a control, the barometer in front of the drag flaps will be used, since it should not be affected by the structural protuberances. Since a suction effect will be accompanied by a change in pressure, the collecting the pressure data from beneath the drag flaps and then comparing it to pressure data taken in front of the drag flaps should show a difference in pressures, mainly that the pressure beneath the drag flaps is higher than the ambient pressure taken by the sensors in front of the drag flaps. The barometric sensors chosen have a resolution of 0.03hPa / 0.24m. This resolution allows for a fairly accurate reading of pressure in the different areas of the rocket. The temperature gathering sensor on the sensor also has a +/­ 2C accuracy, which might affect the resolution of data gathered. Sources of error include the accuracy of the sensors which could lead to inaccurate readings. The preliminary process procedure is as follows. Before launch, pressure reading will be taken from both sets of barometers and logged. After motor burn out when the drag flaps begin to deploy, measurements from both sets of barometers can be collected again and a difference between the two can then be calculated and then logged with a timestamp in order to compare pressure the effects of the angle of the drag flaps and the pressure differences. 5.3 Design of Payload Equipment 5.3.1 Drag Flap Subsystem The drag flap subsystem of the rocket will be used to accomplish multiple requirements simultaneously. These are Vehicle Requirement 1.1, and Payload Requirement 3.1.6. Vehicle Requirement 1.1 states that the rocket must deliver the payload to an altitude of 5,280 feet. The rocket motor has been selected such that if the rocket were to fly without the drag flaps, the final altitude would be slightly above this height. The team will ballast the rocket to ensure a maximum altitude of less than 5600 feet in the event of a total drag flap system failure. The drag flaps will be dynamically controlled in­flight to decrease the vertical velocity of the rocket and adjust the maximum height of the rocket to exactly 5,280 feet at apogee.

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The drag flap subsystem will also fulfill Payload Requirement 3.1.6, “Aerodynamic Analysis of Structural Protuberances.” Barometric pressure sensors will be placed above and below the flaps, and will be gathering data throughout the flight. Additionally, the performance of the drag flaps will be simulated and studied using ANSYS Fluent CFD software. This analysis will help to verify the robustness and performance of the drag flap subsystem.

Figure 19: Diagram of Drag Flap Subsystem Assembly

The Drag Flap Subsystem consists of four quarter­cylinder shaped elements which fan out from the body of the rocket to create a drag force, which in turn decreases the rocket’s vertical velocity. The quarter­cylinder shaped elements, or flaps, are designed to move simultaneously with one another. Their motion is controlled by a single DC stepper motor. The motor will be attached to a threaded rod, which will in turn be attached to the middle plastic collar via the threads of a nut glued onto the plastic slider. The threaded

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rod drive shaft is connected to the motor shaft with a small cylindrical coupler made of aluminum. The coupler will grip the drive shaft with a length of threads, and it will connect to the motor with epoxy and a small set screw that grips the flat of the motor shaft. As this slider moves up and down, it causes the flaps to rotate in and out. The flaps are designed to extend to a maximum angle of 38.8 degrees. This maximum deflection results in an increase of 27.37 square inches of additional vertical projected area. One stage of the verification of the drag flap design was performed using ANSYS Fluent CFD software. The goal of this analysis was twofold. The first goal was to characterize the flow around the drag flaps at different deployment angles to ensure that the stability of the vehicle would not be compromised during flight (this part of the analysis was further verified by the results of the full scale launch). The second goal of the CFD analysis was to determine a function which would relate the deployment angle of the flaps to a resultant drag force on the flaps. This analysis would aid in the creation of a control algorithm to drive the flaps. It would also provide a basis for which to perform a more detailed finite element analysis on the flaps themselves to further verify their ability to withstand the drag forces seen in flight. Again, this FEA analysis would supplement the static load testing which was already performed by the team. For all of the CFD analysis, the flaps were deployed at their maximum deployment angle of 38.8 degrees. Furthermore, the simulation was performed at the maximum airspeed the vehicle would experience, as determined by OpenRocket simulation. The figure below shows the pressure distribution for the flaps:

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Figure 20: Pressure distribution around fully deployed flaps This analysis shows the maximum and minimum pressures seen by the drag flap system. The minimum pressure is approximately 0.5 atmospheres, and the maximum pressure is approximately 1.7 atmospheres. Another way to preliminarily assess the stability of the rocket using CFD is to look at the velocity streamlines around the flaps. The figure below shows these streamlines:

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Figure 21: Velocity streamlines around fully deployed flaps

It can be seen that the air flowing around the flaps forms vortices which curl around the flaps. This effect will be further characterized by pressure sensors placed around the flaps. Additionally, the CFD was used to develop a function to describe the increase in drag force on the vehicle as a function of deployment angle. The graph of that function is shown below:

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Figure 22: Function relating deployment angle to force and drag coefficient

This polynomial function can be used to develop a control algorithm to drive the motion of the flaps. The control system for the drag flaps will be discussed later. Additionally, the function can be used to verify the strength of the drag flap design using FEA techniques. Since the maximum force on the flaps is known, this value can be applied to the flaps in a finite element analysis environment to simulate the stresses on the flaps. The maximum force on the total flap system is 357 N at full deployment. This corresponds to a force of 89.25 N on each flap. This scenario was modeled using NX Nastran FEA software. The flap was loaded with 89.25 N of force on its outer face, and a fixed constraint was imposed on both of the two hinge points of the flap to hold it in place. The results of the simulation are shown below:

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Figure 23: Stress distribution in loaded drag flap

The largest stresses seen by the flap occur at the corner which connects the lower arm to the flap body. The magnitude of the stress at this location is 11.651 MPa. Since the yield strength of PLA (the plastic out of which the flaps are made) is about 62 MPa, this provides a factor of safety of 5.32. Since this analysis represents the “worst case” that the drag flaps will ever see, this analysis is a very reassuring tool to verify the design robustness of the flaps. The final piece of analysis to verify the structural robustness of the drag flap subsystem is an analysis of the forces put on the stepper motor. Firstly, The stall torque of the stepper motor being used is 27 N*cm. Since we are operating the motor at half power to conserve battery, this means the stall torque of our system is 13.5 N*cm. Figuring out how much torque is needed to move the drag flaps is difficult, since the amount of force on the flaps is a function of angle, but the amount of moment that the drive system can impart on the flaps is also a function of angle. To understand the basis for the following analysis, it is helpful to consider two points in the motion of the flaps during which the calculations become trivial. The first point is the condition in which the

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flaps are fully opened. It can be see in the drawing of the system that in this state, the rods which connect to the flaps and push on them are vertical.

Figure 24: When flaps are fully closed, pusher rods are vertical

It is important to note that in this state, the force on the flaps which would resist their motion is 0 newtons, because they are flush with the body of the rocket. Thus, a motor with any stall torque could move them while they were in this position. However, as soon as they move out of this position, a moment about the flap’s axis of rotation is introduced. Let us next consider the case in which the flaps are fully opened. This case occurs when the pusher rods are completely horizontal, and thus they can’t push the flaps out any farther.

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Figure 25: When flaps are fully opened, pusher rods are horizontal

In this configuration, there is significant force on the flaps. This creates a moment which tries to push the flaps back inward. There is an inward force on the pusher rods from the flaps. However, it is not hard to visualize that in this position, it would be easy for the motor to move up and down, since the direction of the force from the flaps onto the pusher rods in completely perpendicular to the direction of motion of the sled on which the rods are attached. Thus, this state results in a similar case to the first state, in which a motor with any stall torque is capable of moving the flaps. In the first state, all of the vertical motion of the rod sled is transferred through the pusher rods to the flaps, because the pusher rods are vertical. In the seconds state, none of the vertical motion of the rod sled is transferred through the pusher rods to the flaps, because the pusher rods are horizontal. It can be assumed that at any state between these two, a fraction of the vertical motion would be transferred, and that that fraction would be a function of the angle of the pusher rods. Thus, a “scaling equation” can be constructed to model the amount of contribution that the rod sled’s motion has on the drag flaps. This equation would have a value of 1 when the flap angle was 0 (fully closed) and a value of 0 when the flap angle was 38.8 (fully opened). A simple linear equation connecting those two points is sufficient for this analysis. Thus, the scaling equation becomes:

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In order to obtain the “effective force” imparted on the rod sled, the scaling equation needs to be multiplied by the force equation obtained from the CFD analysis. First, however, it should be noted that the equation obtained in the CFD analysis is an expression of the total drag force of the rocket. To consider just the drag force on the flaps, the last term of that equation must be dropped, resulting in:

Multiplying the two together yields:

This equation is an approximation of the effective force on the rod sled as a function of flap deployment angle. The equation takes into account the forces on the drag flaps as a function of angle, and the geometry and dynamics of the drag flap system itself. Not surprisingly, this equation has a local maximum. When plotted from 0 to 35 degrees, the local maximum can be found:

Figure 26: Plot of force on drag flaps (red) and “effective force” on rod sled (green)

The maximum occurs when the flaps are opened 26.113 degrees, and it has a value of 47.044 newtons. From this, the torque required in order to impart 47.044 newtons of force on to the rod sled can be calculated. The mechanism which moves the rod sled is a lead screw. The screw has a diameter of 0.288 inches, and a thread count of 20 threads per inch. With these two quantities, a sort of “mechanical advantage” value can be calculated based on the angle of the threads. For now we can neglect friction, but it will be addressed later. With these dimensions, the mechanical advantage of the lead screw is 9.12. This means that for a given radial force on the screw (torque divided by the screw radius), a force can be imparted on the rod sled which is 9.12 times greater. Thus, we can work backwards and see that the radial

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force on the the screw required to impart 47.044 N onto the rod sled is 5.165 N. Since torque is simply force times distance, and the radius of the screw is 0.28956 cm, this means that the torque required to impart 47.044 N onto the rod sled is 1.493 N*cm. Since friction was neglected in this analysis, we must consider that these results are for a system with 100% efficiency. Typical lead screw efficiencies for acme type lead screws range from 30% to 50%. Being conservative, let us assume that our lead screw has an efficiency of 30%. This means that the required moment is about 5 N*cm. This is less than half of the stall torque of our motor, and this value is the “worst case.” Thus, we can be sure that the motor will be able to spin the drag flaps at every point during flight. The motor which actuates the drag flaps will be controlled using an Arduino Zero clone (Neutrino). The Neutrino will be connected to a motor driver, and to one of the accelerometers. The design of the control system is in progress. A simulation software has been developed to characterize controller performance under an array of varying flight conditions, including body mass, ground air density, launch rod angle, and coefficient of drag. The software is sensor­aware and simulates measurement noise to facilitate design for robustness. Using simulation results from OpenRocket, the simulator's accuracy has been verified. For a simulated flight without controls applied and perfect standard conditions, the observed apogee is 1658 meters (~5440 feet). The state history of the simulated flight is seen below in Figure 27.

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Figure 27: State History of First Simulated Flight

A demonstration of the sensor fuzzing capabilities is seen below:

Figure 28: Demonstration of Added Sensor Noise

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The altitude is measured using a barometric altimeter, and the acceleration is measured using an accelerometer. Noise is added to each reading according to expectations of their accuracy. In this case, velocity is derived directly from the altitude data. This signal is visibly noisy, and it will later be shown that it must be processed further for the controller to contribute positively to the competition apogee score. The velocity can also be directly derived from the acceleration data, but it is inaccurate under high accelerations and prone to drift in both velocity and altitude estimates with the latter varying as much as 35 meters (~110 feet). It will be necessary to fuse both readings (possibly with a Kalman filter) in order to extract a state measurement that is accurate enough. An initial glimpse of the sensitivity of the final apogee to velocity noise can also be seen in the previous figure. The red lines represent the minimum and maximum predicted apogees with no control and with maximum control. It can be seen that, using the derived velocity, the final apogee varies up to 50m (~160 ft). A visualization tool developed to aid in the design of the control system is seen in the following graph:

Figure 29: Altitude and Velocity Graph of Possible Control Programs

The state domain (altitude and velocity) is plotted in Figure 29. The greyed out regions represent states for which there is no possible control program that can be employed to guide the rocket to exactly 5280 feet. These regions are found by propagating rockets backwards through time from the target height with control values of 0.0 and 1.0. These values represent the range that the flaps are extended, with 0.0 and 1.0 (fully closed, fully open) being the extreme values. The lower region represents states for which there must be negative control (impossible: u < 0.0); these states do not carry enough energy

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to reach 5280 feet even with no control for the flight. The upper region represents states for which there must be control higher than possible (u > 1.0); these states carry more energy than can be bled off in time even with maximal control throughout the flight. The white region therefore represents the states for which there is a valid control program that can be employed to guide the rocket to exactly 5280 feet in ideal conditions. It is termed the region of feasibility, and is unique to the rocket's geometry and weather conditions. The region of feasibility is parameterized in the drag flap drag coefficient, which is assumed to be 1.28 in this case. An example launch simulation is plotted on the graph. The boost phase expectedly propels the rocket past the minimal control region into the region of feasibility. The region of feasibility is seen to narrow significantly; this happens because the control effectivity decreases with dynamic pressure/velocity, which can be seen in the trend along the x­axis. An important point on the graph is the intersection of the flight data with the x­axis (off chart), which represents the final altitude. Both regions of infeasibility converge on the desired altitude of 5280 feet. The example launch overshoots the target, which the competition rocket is designed to do. The reason that the velocity signal derived from the barometric altimeter reading cannot be used effectively without further processing is visible in the previous graph. Note that there is a slim margin between the simulated flight path of the rocket and the lower bound that averages 20 m/s (~60 ft/s). A control system based on the velocity signal may easily shed too much energy and fall into the infeasibility region due to the high randomness of the signal. This is observed in simulations with an active controller as seen in the following figures:

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Figure 30: Simulation with Control System Based on Velocity Signal

Figure 31: Graph of Control System Based on Velocity Signal with Added Noise

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Figure 32: Graph of Possible Control Programs Based on Velocity Signal

Figure 30 depicts the control program that the rocket creates on­the­fly. The rocket attempts to find an unchanging control value that will guide the rocket to its final apogee based on the measured state. In this case, the drag flap motor is not modelled, and so there is extreme chatter in the program that is caused by wildly varying velocity measurements. In Figure 31, it can be seen in the last graph that the expected altitude (blue line) based on measured state varies considerably. The mostly constant segment after 4 seconds represents the time after which the control system is activated. It is able to achieve the target altitude most of the time, however, as time progresses, the control authority is seen to drop. This is expected because the controller is overcompensating due to overestimated velocity points and bringing the energy down, but the undercompensation cannot bring it up at a similar rate. This also suggests that unbounded controllers like PID controllers may cause issues if they are not properly tuned. The rocket undershoots the target altitude. In Figure 32, this failing is visualized on the feasibility graph. The flight path of the rocket crosses the lower bound prematurely, signifying that there is no longer a possible control to reach the target altitude. The intersection point of the flight path on the x­axis coincides with the final apogee of 1588 meters. In the following figures, an example simulation is carried out with perfect sensor readings that targets the ideal 1609 meter apogee perfectly:

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Figure 33: State History of Third Simulated Flight

Figure 34: Graph of Third Simulated Flight with No Bleed

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Figure 35: Graph of Possible Control Programs Based on Third Control System Type

In Figure 33, it is seen that the controller is applying the expected control. It does not considerably change during the flight. There is no chatter due to perfect sensor readings. In Figure 34, there is no bleed in the control authority. In Figure 35, the flight path correctly straddles the region of feasibility and converges along with the upper and lower bounds on the target apogee. These results and tools lay the groundwork for the evaluation of potential control systems. The next step is designing a filter for the measurements so that the state can be estimated accurately. The team has identified PID controllers as a candidate for control, due to their ubiquity, which will soon be designed and evaluated. The motor which will be used to actuate the drag flaps is a bipolar stepper motor. The motor specifications are shown below in Table 15.

Table 15: drag flap stepper motor specifications

Model number QSH­4218­35­10­027

Steps per Revolution 200

Supply Voltage 5.3 V

Current per Phase 1.0 A

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Holding Torque 27 N*cm

The drag flap subsystem is made up of ten separate plastic pieces, all of which will be 3D­printed using a fifth generation Makerbot. All prints are done using Makerbot’s PLA (Polylactic Acid) plastic. The part’s layer thickness and infill percentage vary based on where the part is located and what types of stresses the part will encounter. The drag flap doors, for instance, are printed at 200 micron layer thickness to allow for a smoother exterior surface and at a 40% infill. All parts were printed at temperatures between 215­225 degrees centigrade to ensure proper layer adhesion. A few manufacturing issues were encountered while printing the parts. One of these issues was a total failure of a print. Print failure was often caused by filament jams or feed failures. Normally, the print operator can notice an impending failure and correct it; however, due to the long print times of some of the parts (upwards of 15 hours), the print could not always be monitored by the print operator. Another manufacturing and assembly issue was considered is the thermal effects of the printing process. Due to thermal enlargement that is a natural byproduct of the 3D printing process, all holes in the system needed to be enlarged after initial prints. One challenge in the analysis comes from the anisotropic nature of 3D­printed parts. This challenge has been analyzed using mechanical properties testing. Once it came to the attention of the project team that one or more load­bearing components of the final rocket design were to be 3D­printed, it was decided that it was necessary to conduct tests on a sample of 3D­printed material to determine its mechanical properties, particularly its Young’s modulus and yield strength. First, a dogbone sample was 3D­printed such that the grain of the sample was perpendicular to the direction in which it would be loaded. This was predicted to be the weakest possible orientation. The dimensions of this sample are tabulated below:

Table 16: Test sample dimensions

Gauge Length: Width: Thickness:

20 mm 6.40 mm 3.37 mm

Using an Instron machine from RPI’s Mechanical Systems Laboratory, under the supervision of a T.A., a tensile test was performed on this dogbone sample. It is noteworthy that, due to the nature of the testing machine, the sample was pre­loaded in compression by roughly 50 newtons. This effect was accounted for when analyzing the data, and the following graphs of load vs elongation and engineering stress vs engineering strain were created:

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Figure 36: Adjusted Load vs Elongation of 3D­Printed Material

Figure 37: Engineering Stress vs Strain of 3D Printed Material

It is noteworthy that the material is extremely brittle, and experienced no measurable plastic deformation at failure. Important specifications such as Young’s modulus and tensile strength were also gathered, and it was decided that 3D­printed material will suffice, even in its weakest orientation.

Table 17: Material properties of PLA plastic

Young’s Modulus:

Yield Strength: Crosshead Displacement Speed:

Extension at failure:

Maximum Load:

334703.3 MPa 33.907 MPa 5 mm/min 0.7916 mm 783.9 N

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5.3.2 Atmospheric Measurement Subsystem

Figure 38: Electrical Schematic of Scientific Payload

The proposed scientific payload design complies with NASA’s requirements outlined in the 2016 USLI Handbook section 3.2.1. The circuit will contain an accelerometer, a magnetometer, a gyroscope, three barometers and thermometers, a GPS, a wireless transmission module, a UV sensor, a solar cell, and a MicroSD module. The scientific subsystem will be constructed on two parallel circuit boards arranged back to back to make efficient use of the space inside the rocket. The external components will connect to the main boards using screw terminals to provide a firm but removable connection. Two Arduino Zero clones (called Neutrinos) working in parallel will serve as the central point for data collection, processing, and transmission. One will be mounted on each board. The first Neutrino will control the two BMP180s (barometric pressure/ temperature sensor), the HIH­4030 (humidity sensor), the UV sensor, and the camera. The second Neutrino will control the GPS, the XBee, the solar cell, and the combination sensor module. Both boards will share control over the MicroSD module, although the first Neutrino will grant card access to the second.

This setup will balance the processing loads across the two microcontrollers which will allow the setup to achieve a high data rate. Devices will be read asynchronously, meaning that wait times will be used to read other devices. For example, the BMP180 needs 47 ms to read while the humidity sensor takes under 300 µs. This setup will call

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the BMP180 and will spend the next 47ms doing other tasks such as reading the humidity sensor before coming back to the BMP180 to finish the read. Data will be buffered to the microprocessor and will be fed into the MicroSD card every time all modules have been updated. Periodically, the second Neutrino will pipe the updated data from the MicroSD card to the XBee. Modules will be throttled to the frequency of the slowest module to ensure that the Neutrinos have ample time to process the data stream. This approach will result in a connection with a cycle time of ~50­75 ms. This time may be throttled further depending upon the load on the XBee, GPS, and camera, whose UART connections only have a limited buffer. The RioRand­LM2596 is a variable switching voltage regulator based on the LM2596 chip. It provides the supplementary components pre­installed. This was found to be easier and safer than building the same circuit with distinct parts. The regulator will be tuned to output 5 V with a max output of 1 A. This supply was tested using an arbitrary waveform generator. The output voltage was found to remain constant even with harsh input waveforms, such as square waves. The UV sensor and humidity sensor are both analog devices requiring an input of 5 V. Since the Neutrinos operate at 3.3 V, the output voltages of these two sensors will be regulated down using a basic voltage divider circuit with 100k Ohm and 200 k Ohm resistors. The other modules feature internal 3.3 V linear regulators. Depending upon the power requirements of the chip, it will be fed from the 5 V source (high power) or from the 3.3 V regulated output of the Neutrinos (low power). Two BMP180 I2C barometers will provide pressure data between the top and bottom of the drag flaps. Due to their communication protocol, their SDA lines will be wired to one of the spare slots on the multiplexer chip. This is because they share the same I2C address, which will cause the protocol to fail if the Neutrino attempts to access one chip, as both will respond. A combination sensor board will provide the orientation, acceleration, compass, pressure, and temperature readings for the scientific cavity. This board will communicate via I2C and will not experience the same problem as the two BMP180s as it will be wired to the other Neutrino. Solar irradiance will be calculated by wiring a small solar cell to a resistor. By measuring the voltage drop across the resistor, the power produced by the solar cell can be calculated. That power can then be scaled with the value of solar irradiance on the ground as the spectrum of light will remain relatively constant for a low altitude rocket given good weather.

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The XBee will be wired to the hardware serial on the second Neutrino. GPS data will be sent in flight. After landing, the primary Neutrino will perform some brief processing before sending out the data logged on the MicroSD card. The GPS module will be wired to the second set of hardware serial ports on the second Neutrino. Data will be saved and transmitted live across the XBee in order to locate the rocket during flight and descent. The camera module will be mounted on a collar that will allow it to rotate freely in a horizontal plane. A diagram of the camera system can be seen in Figure 39 below. A weight equal to or slightly greater than the mass of the camera will also be attached to the collar that will allow the camera to orient itself into the correct position upon landing to capture both the ground and sky. There will be a threaded rod attached radially to the rotating collar. A weight disc will be able to slide along this rod, and will be secured by two nuts. By moving the weight along the threaded rod, the moment arm of the weight will change, allowing for the weight to be finely tuned so that the camera is horizontal when the system is at rest. A diagram of this concept can be seen below in Figures 40 and 41. The camera will be oriented such that the photos taken during descent will be oriented correctly in order to meet the requirement of having the sky toward the top of the photo and the ground toward the bottom. Images taken once the vehicle is landed will need to be rotated digitally by 90 degrees in order to orient the sky toward the top of the photo and the ground toward the bottom. The initial code for rotating the images can be found in Appendix D. This design assumes the rocket body will land on its side. A polycarbonate ring integrated into the body will allow light to pass through to the camera.

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Note # Note

1 Camera

2 Weight Attachment Point

3 Axis of Rotation

4 Slip Ring

Figure 39 ­ Camera subsystem

Figure 40 ­ Weight concept

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Figure 41 ­ Weight concept

The electrical schematic can be found in Appendix A. One aspect not displayed in Appendix A is the presence of an two electrical slip rings in the camera system. These slip rings, which can be seen in Figure 39, shall be used to transmit power and signal to and from the camera while the system is rotating. 5.4 Payload Verification

The drag flap subsystem obtains its data from the altimeter. It then reacts based off the altitude measurements. In order to verify the function of the drag flaps, another Arduino imitated the altimeter on the same serial port. This Arduino then fed the drag flap system with altitude values that imitated a launch that would undershoot and overshoot the goal apogee. The drag flaps were then observed to ensure that they did actuate 2s after launch and retracted after apogee was reached. The scientific payload will be verified through individual testing of the components with all the software running. For example, the GPS will be tested by moving the payload, and the result will be confirmed by seeing the change from the XBee and the base station. The thermometers and humidity sensor will be tested by placing a finger near or over the sensor, which will cause the values to change. UV and solar irradiance will be tested by moving from inside to outside, or from a sunlit area to a shaded one. Barometers will be tested by riding an elevator up and observing the change in air pressure. Acceleration and rotation will be measured by moving and tilting the payload in all dimensions. Since the scientific payload is effectively a linear system, e.g. the humidity sensor’s measurement does not affect that of the barometer, the confirmation of function of each individual component will assure the correct function of the entire system (assuming it is coded properly). To confirm the overall system (apogee

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detection etc) fake values for sensors will be fed to another Arduino, which will confirm the transitions between different modes of the scientific payload, such as “waiting for launch” and “hit the ground”. The structural integrity of the drag flap system has been verified through a combination of CFD, FEA, and physical sample testing. All of this verification has been discussed in previous sections.

Table 18: Payload verification

Requirement Design Feature Verification

Sky toward the top of picture, ground toward bottom Camera, collar system

Manipulate the structure in various orientations (parallel to the ground, right side up, and upside down) and

ensure the sky is at the top of the picture and the ground

is at the bottom

Camera rotates to be parallel to the ground Collar system

Manipulate the camera subsystem in various

orientations like above, and verify that the camera is parallel to the ground

The image stored on the SD card and the image

transmitted to the ground station are the same

Camera, SD card, transmission system

Take pictures, save to SD card and transmit to ground

station. Take hash of resultant images and compare for similarity

Drag Flaps read data from altimeter

Altimeter/Arduino control system

Arduino fed with simulated altitude values to test response of drag flaps

Atmospheric Measurement system records accurate

data

GPS, thermometer, humidity sensor, UV sensor, solar cell, barometer, gyro,

accelerometer

Each sensor is independently tested to verify proper functionality

Drag Flap system is strong enough to withstand forces

of flight Drag Flap system Combination of FEA, CFD,

and Physical Sample testing

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5.5 Safety and Environment (Payload)

5.5.1 Safety Officer See Section 7 5.5.2 Preliminary Checklists See Section 7 5.5.3 Failure Modes See Section 7 5.5.4 Hazard Analysis and Environmental Concerns See Section 7

6. Launch Operations Procedures 6.1 Checklist 6.1.1 Recovery Preparation

1. Pack both parachutes. Ensure that parachutes can easily slide in or out of rocket. a. Physically insert and pull parachute out of rocket main body. If it gives any

undue resistance to being pulled out Then it must be repacked tighter. 2. Attach quick links.

a. When a quick link is attached a team member should watch it’s attachment and write down which link was attached, so that we are certain of which links are hooked.

3. Mark the Rocket’s center of pressure. a. This should be read off of Openrocket Sims.

4. One officer should complete all necessary registration at this time a. Preferably the acting leader.

5. Assemble motors a. Refer to instructions that come with rocket. Have two team members do this,

one that can check the first team member is doing everything right. Go to section 6.1..2 for more details.

6. Insert and lock motor. 7. Mark the rocket’s center of gravity.

a. This can be done by measuring the rocket;s balance. 8. Prepare materials for four electric matches / charges

a. Carefully count out the amount of power that is needed and out it next to each wire.

9. Slide Nomex shock cord sleeve onto bottom section of shock cord. a. Nomex should not bunch, but should be evenly dispersed along coard.

10. Run matches down to top center ring.

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11. Prepare two loads of insulation (dog barf) 12. Insert charges. 13. Pack insulation above charges.

a. A conservative amount should be applied. 14. Z fold shock cord into body section.

a. This should be done by a team member who has packed a rocket parachute before.

15. Refold nomex over drogue parachute. 16. Attach matches to drouge block.

a. Bend wires to ensure firm connection. 17. Attache drogue parachute to quick link, Insert drogue to rocket main frame. 18. Close bottom section 19. Cover main parachute and shock cord with Nomex. 20. Place uncovered shock cord within main parachute fold. 21. Place matches on upper bulkhead

a. The ejection charges are to be properly filled and sealed. 22. Insert insulation and main parachute.

a. Insulation will be placed within the rocket body so that the parachute is not damaged by the ejection charge. The parachute will be folded so that it may unfurl without tangling. Close upper section

b. Seal with shear pins. 23. Finalize flight card 24. Double check all electronics (ensure right number of beeps).

a. The charge trigger altimeter is to be tested on the ground before a launch to verify that it is programmed and wired correctly.

25. Finalize any flight card details and goto RSO. a. Both Safety Officer and acting team leader should approve flight card.

26. Slide onto rail. Do not force rocket! 27. Turn rail up.

a. Lock ground pad pin so that rail stays in place. 28. Insert igniters. Short wire.

a. Wire should remain shorted at all times until final stage. 29. Turn on altimeter and drag flaps.

a. Check again for right number of beeps. Electronics leader must verify. 30. If necessary adjust angle of rod. 31. Unshort ignitor wire.

a. Rocket is now live. If any more work needs to be done wire should be shorted again.

32. Give flight card to LCO/RSO. Rocket is now ready to fly a. Flight card details should be run past acting president before handing in card.

33. Launch. 34. Recover.

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a. Approach downed rocket with caution. Ensure that motor is dead (no flames can be seen) and that all charges have detonated. Safety officer must give approval before team can move in and get rocket. Take multiple photographs before moving rocket for analysis. Immediately document any breaks or tears that become apparent.

35. Clean motor a. Alcohol or regular soap can be used for this.

6.1.2 Motor Preparation The appropriate motor delay will be selected based on previous simulations. The motor will be mounted within the casing and the integrity of the interface with the rocket body will be checked. The retention cap is screwed on. The igniter charge is installed securely before launch. 6.1.3 Launcher Setup The rocket will be slid onto the launch rod horizontally and then raised to a vertical position. 6.1.4 Igniter Installation The igniters will be installed shortly before launch, after the rocket is situated on the launch rod. The igniters will be carefully handled and shorted such that static discharges do not trigger the igniters. The igniter will be fully inserted and stopped off so that it is firmly held in place. The igniter will then be connected to the launch console. 6.1.5 Troubleshooting A static test of the flight computer telemetry will be conducted on the pad to ensure that the radio link between the RRS team and the rocket is operational. The flight computer will be checked for full functionality and sensor readings will be checked to ensure that they are connected/functioning properly. 6.1.6 Post­Flight Inspection The rocket body will be checked for visible structural damage. The flight computer will be tested for functionality and the wiring checked. The interfaces between separating fuselage components will be checked for damage, and the parachutes and shock cords will be checked for tears and burns. The drag flap system will be checked to ensure that the actuation system has not become damaged under the force of landing. The competition altimeter will be read by a competition official.

7. Safety Plan

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7.1 Safety Officer and General Safety Actions RRS has identified Philip Hoddinott as the acting safety officer. His responsibilities include ensuring shop safety and hazardous material procedures, which is partly accomplished through safety quizzes administered by the RPI School of Engineering. He will oversee the safe construction and launch of the pertinent rocket vehicles through supervision and inspections.

Hazard recognition and accident avoidance All team members must have taken and passed a safety quiz on hazard recognition

and avoidance. Additionally they have all attended a hazard briefing at the beginning of the school year.

Pre­launch briefing The safety Officer will give a pre­launch briefing upon arrival to the launch site. All

teammates will be reminded of the rules and regulations present at launch sites. At all times teammates must comply with reasonable orders given by any officer. At any time an officer may call for a pause while

Purchasing and handling of rocket motors All Rocket motors will be purchased on site. This will lower dangers involved with

transporting flammable materials across state lines.This will also allow the team to not have to worry about state and local laws involving the transport of hazardous materials.

Transportation of rocket to huntsville The rocket will be packed and driven to huntsville, AL. Care will be taken to ensure that

the rocket is not damaged in the drive. The rocket will be packed in such a way that it will not provided a hazard to vehicle occupants (it must be firmly secured) and it will not in any way interfere with the driver’s ability to see.

Safety Agreement All RPI Rocket Team members have signed a safety agreement before becoming part

of the team and working in any lab space or attending any rocket launches.

7.2 Preliminary Checklists The safety officer and team mentor will oversee the final assembly of the rocket and its subsystems prior to flight, as well as the launch pad preparation operations. Two preliminary checklists have been developed for each phase. The pre­launch operations checklist broadly covers the assembly of the major structural components prior to the rocket arriving at the launchpad. The launch pad operations checklist covers the procedures to be completed prior to launch. The more detailed checklist can bee seen in the above section. The checklists are detailed in Tables 19 & 20

Table 19: Pre­Launch Operations Checklist

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Category Description Required Actions

Parachute Deployment

Dog barf will be inserted between the ejection charges and the parachute to prevent burning, and the parachute will be properly folded and then inserted.

Dog barf, parachute folded, parachute inserted.

Avionics Avionics equipment will be properly connected and inspected. Batteries will be tested for charge and inserted into their bay. Avionics will be tested and then slid into bay and secured.

Wiring is firm, batteries are charged, batteries are firmly strapped into bay, avionics respond to communication handshake, avionics slid into bay in right orientation, avionics sled secured.

Payload Payload is inserted into payload tube and fixed. Shear pins are inserted.

Payload inserted in right orientation, payload bay secured, shear pins inserted.

Drag Flaps Drag flaps will be checked for unobstructed movement.

Verify that drag flaps can extend and retract fully by running control system test.

Table 20: Launchpad Operations Checklist

Category Description Required Actions

Rocket Assembly Rocket assembly is complete.

RSO Safety Check RSO will check final rocket before launch.

RSO Approval. RRS SO Approval. RRS Mentor Approval.

Avionics Avionics will be checked for proper communication and data relay.

Data link is clear.

Avionics Avionics charge and ignition switch will be turned off to prevent premature ignition.

External switch is OFF. Confirm status light is also OFF.

Ejection Charges Ejection charges will be installed.

Ejection charges installed

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Motor Ignition Motor igniter will be installed. Delay charge will be installed. Leads will be connected during handling to prevent static electricity discharges.

Motor igniter leads are connected during installation. Motor igniter is installed. Motor igniter is properly wired. Motor igniter leads are separated. Delay charge is installed.

Launch Configuration Rocket will be placed on the launch rod and given the desired angle.

Rocket rail buttons are on launch rod. Selected angle is confirmed by RSO.

Avionics Avionics charge and ignition switch is turned on to arm the system.

External switch is ON. Confirm status light is also ON.

Ejection Charges Ejection charges will be armed.

Ejection charges are connected and will not short.

Final Safety Check Final safety check will be conducted by RRS SO, RRS Mentor.

RRS SO Approval. RRS Mentor Approval.

Data Retrieval Final altitude will be retrieved.

Altitude has been retrieved by SL official.

Post­Flight Inspection Inspection will be performed on rocket following flight.

RRS SO Approval. RRS Mentor Approval.

7.3 Failure Modes The following table summarizes the launch and flight failure modes that have been identified. They are addressed with a mitigation plan and a risk likelihood/impact.They are ranked using the USLI Risk Assessment Codes (RAC)

Table 21: Failure Modes

System Failure Effect Pre ­ RAC

Mitigation Post ­ RAC

Motor Dislodgement from housing

Dislodged motor could damage equipment or seriously hurt onlooker.

1A Employ retainer structure with three bulkhead centering rings. Have experienced officer Follow rocket motor instructions to ensure mount is successful.

1C

Computer/Control

Microcontroller failure

Altimeters could prematurely deploy

1B All wiring will be finalized and tested before launch to prevent electrical

1C

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parachutes or not deploy parachutes at all.

malfunction. Code will be reviewed and tested during scale launches. Wiring will be soldered and insulated.

Stepper motor jam

Drag flaps become inoperable, rocket will not compete at peak performance.

3B Stepper motor will be tested against static loads to simulate expected air speeds. Mechanism will be lubricated and inspected for particulates before launch.

3C

Forward flap system causes instability

Unstable rocket could fly towards onlookers.

1B Verification of stability will be performed in OpenRocket and using test launches.

1C

Structural Fins shear off during flight

Rocket would become unstable,

2B Fins will be embedded in body and fixed to the internal motor mount and body with heat­tolerant epoxy resin.

2C

Rocket motor explodes

Debire of rocket would come down with enough speed to cause damage / injure.

1B Test launches, finite elements analysis, and static load tests will be performed to verify integrity of rocket body.

1C

Parachute tears, parachute fails to deploy

Rocket without parachute would come down with enough speed to cause damage / injure.

2B Parachute will be checked for defects before packing. Parachute will be packed properly.

2C

Shock cord breaks

Rocket will split into multiple parts, and may lose a parachute.

2B Recently­manufactured shock cord will be used to ensure integrity. Shock cord will be properly secured and tested under static and jolt loads.Nomex covering will protect from charges.

2C

Drag flaps shear off during flight

Drag flaps could hurt onlooker, may damage rocket.

2B Mechanical components and connections will be shear tested under static loads.

2C

Recovery Parachute ejection failure

Rocket without parachute would come down with enough speed to cause damage / injure.

2B Redundant powder charges will be used for the main parachute. Packing will be inspected before launch.

2C

Parachute ejected prematurely

Parachute deploying too early will rip rocket apart. Additionally peice of

1B Code will be reviewed for correctness. Hard failsafes will be implemented to prevent ejection above a predetermined altitude.

1C

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rocket may fall and hurt onlookers.

Computer system failure

Premature charge detonation could hurt onlooker / damage rocket.

2B Flight computer will be tested using simulated soft landing jolts to ensure component survivability. Wiring will be soldered in place and insulated.

2C

Transmitter failure/interference

Rocket will be unable to be recovered.

2B Transmitter will be tested at range. Interference will be studied.

2C

Launch operations and transportation

Physical damage during handling

Rocket may be dropped / squashed/ crumpled during transit.

1B Each component of the launch vehicle will be transported in a separate, cushioned box. Launch vehicle will be inspected on site for defects.

1C

Live charges on board after launch command

Any premature detonation could hurt nearby people / damage rocket.

1B Avionics will be powered off. Discrete switch will cut off system power entirely.

1C

Premature firing of ejection charges and/or motor

Any premature detonation could hurt nearby people / damage rocket.

1B Wiring will be ensured to proper through tagging. Ejection charges will be disconnected until launch. Igniters will be installed at launch. Computer will have a discrete switch only to be armed during launch to prevent premature signal.

1C

Spectator injury

A rocket that suffers some sort of failure could injure a onlooker.

1B Range will be cleared prior to launch. No large, heavy objects will be installed according to NAR safety code.

1C

The flight procedure will be practiced by the team prior to the final competition launch. Assembly of the rocket will be checked. The team safety officer and team mentor will supervise the final assembly using checklists and will handle the final preparation on the launch pad.

7.4 Hazard Analysis and Environmental Concerns

The team safety has identified a list of common personnel hazards that may be encountered during the manufacture and testing of the vehicle. All participating team members will have completed safety training as administered by the RPI School of Engineering. A comprehensive list of MSDS are freely available to the team. Team members will be briefed on the NAR safety code of conduct by the team safety officer and the team mentor. The following table summarizes the hazards that have been identified with mitigation steps, required safety equipment, and emergency equipment.

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Table 22: Hazard Analysis

Hazard Effect Pre ­ RAC

Mitigation Required Safety Equipment

Emergency Equipment

Post ­ RAC

Drill press rotation

Contact with drill could cause serious injury / death.

1B No loose clothing/hair

Glasses First aid kit 1C

Band saw Contact with saw could cause serious injury / death.

1B No loose clothing/hair

Gloves when working with large pieces, glasses

First aid kit 1C

Epoxy / Paint fumes / irritation

Inhaled fumes cause damage, substance can cause blindness if eye contact.

2B Work in ventilated, spacious area

Face masks, gloves, glasses

First aid kit, eye flushing station

2C

Propellant/powder ignition and skin irritation

Powder can irritate skin. Hot powder can burn skin and flesh.

2B Motor will be stored in separate, insulated cabinet. Motor will be handled with gloves after is has cooled.

Gloves, glasses

First aid kit, eye flushing station, burn kit

2C

Vehicle debris

Falling vehicle debris could seriously injure an onlooker.

2B NAR regulations will be obeyed. Team members will be trained to watch skies and avoid any falling vehicle

Glasses First aid kit 2C

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debire.

General cuts and irritation

Cuts can be caused by a wide range of objects both at launches and in our workshop.

3B Gloves will be worn when working with sharp/irritating implements, and glasses will be worn when operating moving machinery

Glasses, gloves

First aid kit 3C

Several environmental concerns that may impact the testing and launch of the rocket have been identified in the following table.

Table 23: Environmental Concerns

Risk Mitigation Likelihood

Inoperable Wind Conditions / Cloudy

Launches will be planned in advance according to weather forecasts to ensure timely completion.

2

Rainy Conditions All electromechanical and electrical parts will be shielded. Body structure will be able to retain integrity in damp conditions.

2

Operable Wind Conditions Launch rod angle will be adjusted. Parachute deployment altitude will be recalculated and adjusted due to higher­than­expected drift.

3

Vehicle Cannot be Found All charges will be detonated by flight computer before landing to ensure that it is inert. Vehicle will have identifiable, bright paint scheme.

1

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8. Project Plan 8.1 Budget Plan

The following graph shows the total budget plan for the Rensselaer Rocket Society Student Launch project. Note some differences from the projected budget plan in the CDR. Among these include an extra thousand dollars granted us from the MANE Department (Mechanical, Aerospace, and Nuclear Engineering), a significantly lower travel cost (since we decided to drive instead of purchase plane tickets), and several items purchased in the time since the CDR. While this list should encompass everything, there may yet be some unforeseen expenses that come up after the time of this report.

Table 24: Detailed expenses Income

Income in Account (WeR Gold) $1,530.00

Membership Dues $1,400.00

School of Engineering Funding $1,000.00

MANE Department Funding $1,500.00

Fundraising Goal $3,000.00

Total $8,601.67

Vehicle Design Team

Item Price Quantity Total

Forged eye bolts with shoulder (w/ nuts), Stainless steel 316, 1/4"­20 $21.27 2 $42.54

3/16" Birch Plywood (consider getting thicker ones~.25") Centering Rings $3.49 4 $13.96

16.75" Length ­ Nosecone $21.95 1 $21.95

36" Length­cut for motor length from tubing 2.15" diameter $14.99 1 $14.99

54 mm ­ Get from Aeropack, ordinary­not flanged $29.00 1 $29.00

Birch Plywood Bulkheads, 1/4" thick $4.78 6 $28.68

48" Length Section 1, Phenolic $25.99 1 $25.99

36" Length Section 2, Phenolic $20.99 1 $20.99

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Rocketry Warehouse custom G10, .125" thick, 4 fins, See Open Rocket File $45.84 1 $45.84

Lexan Polycarbonate tube, 8" Length 4" OD, 3.75" ID $15.99 1 $15.99

Stainless Steel 316, 6ft, 10­32, McMaster­Carr $12.30 1 $12.30

7" lengths of Phenolic coupler tubing for 3.9" diameter airframe $4.99 3 $14.97

Aerotech K1103X Motor $101.69 2 $203.38

Sanded plywood sheet $10.23 2 $20.46

Wood filler 8oz. $4.99 1 $4.99

Blue Epoxy PC7 2oz. $5.99 1 $5.99

Nuts&bolts $0.27 14 $3.76

Threaded rod $8.58 1 $8.58

Phenolic CT­3.9 $9.97 3 $29.92

HAMR­54 Replacement Sleeve $23.90 1 $23.90

HAMR­54 Replacement Cap $22.90 1 $22.90

Bulkplate, 3.9” $13.31 4 $13.31

Phenolic Airframe Tubing 3.9”x36” $35.94 1 $35.94

Phenolic Airframe Tubing 3.9”x48” $40.94 1 $40.94

Phenolic Airframe Coupler $34.95 1 $34.95

Aluminum Angle Slide 36”x0.5”x0.125” $15.84 1 $15.84

Phenolic Airframe Tubing 2.1” $14.99 1 $14.99

Centering Ring 3.9­2.1 $3.49 4 $13.96

Custom Fins 0.125” $11.46 4 $45.84

54­1706 Motor Casing $200.00 1 $200.00

Makerbot 3D Printer Filament ****** $50.00 1 $50.00

Total $1,076.85

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Recovery Design Team

Item Price Quantity Total

PerfectFlite StratoLogger CF $54.95 1 $54.95

Raven3 Altimeter $115.00 1 $115.00

Skyangle Cert­3 Drogue $27.50 1 $27.50

Skyangle Classic II 52 ­ see FS image and weight of rocket for calculation $82.50 1 $82.50

Ejection Charge $5.00 2 $10.00

Nuts&bolts, galvanized eye­bolt $27.18 1 $27.18

Hex machine screws, SAE flat washers, machine screws $12.82 1 $12.82

2x4 BCX Plywood $9.74 1 $9.74

4"OD x 3 3/4"ID x 12" Polycarbonate Tubes $30.14 1 $30.14

2­foot Diameter Drogue Parachute $75.50 1 $75.50

Skyangle Classic II 52” Main Parachute $82.50 1 $82.50

Raven3 Altimeter $165.00 1 $165.00

Total $692.83

Payload Design Team

Item Price Quantity Total

Nuetrino Microprocessor $20.00 2 $40.00

BMP180 Barometric Pressure/Temperature/Altitude Sensor­ 5V ready $9.95 2 $19.90

Adafruit 10­DOF IMU Breakout ­ L3GD20H + LSM303 + BMP180 $29.95 1 $29.95

Adafruit Ultimate GPS Breakout ­ 66 channel w/10 Hz updates ­ Version 3 $39.95 1 $39.95

CR1220 12mm Diameter ­ 3V Lithium Coin Cell Battery ­ CR1220 $0.95 1 $0.95

TTL Serial JPEG Camera with NTSC Video $39.95 1 $39.95

MicroSD card breakout board+ $14.95 1 $14.95

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Kingston 8 GB microSDHC Class 4 Flash Memory Card SDC4/8GBET $3.92 1 $3.92

2 Samsung INR18650­25R 18650 2500mAh 3.6v Rechargeable Flat Top Batteries (Blue/Green assorted) $10.20 1 $10.20

Nitecore i2 Intellicharge Charger for 18650 AAA AA Li­Ion/NiMH Battery + Case $14.09 1 $14.09

XBee Pro 900 RPSMA $54.95 1 $54.95

Analog UV Light Sensor Breakout ­ GUVA­S12SD $6.50 1 $6.50

SparkFun Humidity Sensor Breakout ­ HIH­4030 $16.95 1 $16.95

Miniature Solar Cell ­ BPW34 $1.50 2 $3.00

Miscellaneous Electrical Components (resistors, capacitors, screw terminals, etc) $36.89 1 $36.89

Cylindrical Battery Contacts, Clips, Holders & Springs 18650 DUAL SOLDER TAIL BATTERY HOLDER $4.44 1 $4.44

Terminal Block, Driver $8.00 1 $8.00

Total $344.59

Travel

Item Price Quantity Total

Hotel Rooms $89.00/night

(5 nights) 3 rooms $1,335.00

Car Rental $206.25/week

(1 week)

1 car $206.25

Gasoline $2/gal (average), 25mpg (average)

48 gallons (x3 cars)

$288.00

Total $1829.25

Educational Engagement

Bubble wrap, duct tape, straws, plates $21.42 1 $21.42

Papercraft $4.31 3 $12.93

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Take out BX bunny/egg 4­pack $2.15 6 $12.90

Total $47.25

Small Scale Launch

Item Price Quantity Total

Estes Pro Series II Leviathan kit $35.00 2 $70.00

Aerotech G80 motor $11.99 1 $11.99

Payload Bay hardware $10.00 1 $10.00

Ejection Charges $5.00 2 $10.00

Total $101.99

Table 25: Summary of income and spending Income $8,601.67

Spending $4,092.76

Difference $4,508.91

*Note that the above numbers for income, spending, and difference are assuming that we reach our $3000 fundraising goal.

8.2 Funding Plan

The original fundraising goal for this project was around $7000, which represents around $1200 for the rocket and around $5000 for travel and lodging, with some surplus set aside for emergencies. At the start of the semester, the RRS had money from membership dues, funding received from the RPI School of Engineering, funding from the RPI Department of Mechanical, Aerospace, and Nuclear Engineering (MANE), and money left over from previous years. This reached a total of $4600. Since the start of the semester the RRS received an additional $1000 in funding from the MANE department; this and other changes are reflected above.

The rest of the fundraising, which was intended to total $3000, was accomplished through a combination of corporate sponsorship and community outreach. The club reached out to over two dozen aerospace companies, either by letter or by email. The

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RRS offered several sponsorship options, each tier unlocked after a certain minimum donation and offering more perks for the company. These perks typically involved some degree of publicity, either on our website, the rocket itself, or during launch footage. The RRS is very disappointed at the lack of positive response to our outreach efforts, but is still hopeful this corporate outreach may yet prove successful. The community outreach strategy proved much more fruitful. The RRS has already raised over $1000 through the WeR Gold program offered here at RPI, a website where alumni and parents are encouraged to donate to particular clubs through a user­friendly process. After supplying WeR Gold with information about the club and current project, pictures, and even a fundraising video posted to Youtube, WeR Gold will feature our club for the rest of the semester in return for a 10% fee, which will be factored into our book­keeping. The RRS has also undertaken a bake sale on campus, and is planning at least one more before the competition. While the funds raised in this manner are small when compared to some of our more ambitious fundraising methods, having a bake sale helps to involve students on campus, presents a fun side­project for club members, and spreads awareness of RRS throughout RPI. Throughout this project RRS has been trying to save money at each turn, while still ensuring solid construction and safety. Frugal decisions such as driving instead of flying to Alabama were all made under the assumption that funds were tighter than they truly were. Fortunately, before the full­scale launch failure, the RRS had a budget surplus on the order of $4000. This will allow the RRS to have ample funding for rebuilding and for any other unforeseen circumstances that may arise.

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8.3 Timeline

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Figure 42: Gantt Chart

8.4 Educational Engagement

8.4.1 National Manufacturing Day Event The RPI School of Engineering sponsored a program for local Troy middle and high school students. The RRS participated in this program by running an event in coordination with the program. For this event, the RRS held a presentation on rocketry and the aerospace industry. Specifically, the presentation covered multiple aerospace organizations, including discussions on NASA, SpaceX, the European Space Agency (ESA), and Orbital Sciences. The RRS talked about manufacturing processes to build rockets with examples of different aerospace materials, such as carbon fiber and fiberglass. Students then saw the results of good and bad manufacturing practices with launch videos. Students were encouraged to ask questions and provide feedback throughout the presentation, as time was limited for a full question and answer or feedback period. This event occurred on November 6.

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8.4.2 STEM Engagement with the Boys and Girls Club The RRS conducted two egg drop events with the Troy Boys and Girls Club in Troy, NY. These events began with a short presentation relating basic Newtonian physics to dropping an egg. Topics covered included forces, gravity, drag, and inertia. Students in grades 2 through 4 answered questions and then each designed their own egg­carrying devices to protect an egg during a fall. After each student completed his or her device, a team member dropped the devices from a height of around 15 feet. The student was then able to evaluate the result of the drop by examining the egg. Most of the students were successful in keeping the egg from cracking, indicating that they had understood the basic concepts the team presented. These events took place on February 29 and March 4.

8.4.3 STEM Engagement with Troy School 2 The team scheduled three dates to travel to Troy School 2 to hold egg drop events similar to those held with the Troy Boys and Girls Club. These were scheduled for March 7, 8, and 9. The team was unable to contact the school administration to confirm these dates as they approached, however, so they were cancelled. Unfortunately, the timing of these events made them impossible to reschedule. Due to the unplanned cancellation of multiple educational events and unexpectedly low turnouts to those events that were held, as of the FRR, the team has not met the requirement of 200 participants in educational engagement programs.

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Appendix

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Appendix A: Scientific Payload Schematics

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Appendix B: Structural Design Assembly Drawings

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Appendix C: Code to Rotate Camera Images

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Appendix D: Motor Thrust Curve

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