Preliminary Design Review

Transcription

1 Preliminary Design Review Rensselaer Rocket Society NASA Student Launch November 3, 2017

2 Contents 1 Summary Team Summary Launch Vehicle Summary Payload Summary Changes Made Since Proposal 6 3 Vehicle Criteria Selection, Design, and Rationale of Launch Vehicle Mission Statement Requirements and Mission Success Criteria Subsystems and Design Alternatives Upper Airframe Lower Airframe Current Vehicle Design Motor Selection Verification Component Analysis Recovery Subsystem Altimeters Parachutes Attachment Hardware and Bulkheads Electronics Tracking Mission Performance Predictions Kinetic Energy Drift Analysis Safety Safety Officer NAR/TRA Mentor Risk Assessment Preliminary Personal Hazard Analysis Material Hazards Facility and Tool Hazards

3 4.5 Preliminary Failure Modes and Effects Analysis (FMEA) Preliminary Environmental Concerns Project Risks Payload Criteria Objective System Overview Software Processing Unit Image Sensor Sensor Position Power Budget Overview Payload Integration Project Plan Requirement Verification Budget Gantt Chart A Requirement Verification 34 B Budgeting 37 C Flysheet 40 D Gantt Chart 44 2

4 List of Figures 1 Team Organization Chart Current Vehicle Design Selected Shape of Fin Thrust Curve for Aerotech K Thrust Curve for Aerotech K Recovery Electronics Schematic Schematic shows the camera field of view, making note of the eventual overlap behind the rocket The electrical schematic shows all connections between components of the payload subsystem Schematic of PiCam v Full schematic of the rocket highlights payload integration and functionality

5 List of Tables 1 Body tube Material Selection Matrix Vehicle Details Statement of Work Items Relevant to Motor Selection AeroTech K828 Specifications AeroTech K480 Specifications Vehicle Mass Requirement and Verification Plan Requirement and Verification Plan Part Component Analysis Altimeter Comparison Kinetic Energy Estimate for Body Sections Calculated Drift Distances for Varying Wind Speeds Processing unit selection study Camera selection study Current best estimate power budget Current best estimate mass budget

6 1 Summary 1.1 Team Summary The Rensselaer Rocket Society (RRS) is a student organization located at Rensselaer Polytechnic Institute (RPI). The RRS operates in the Ricketts Building at RPI. The RRS s faculty advisor is Dr. Jason Hicken, Assistant Professor in the department of Mechanical, Aerospace, and Nuclear Engineering. The Community Mentor for the RRS is John Sicker (NAR Level 2 Certified, NAR Number SR, TRA Level 3 Certified, TRA Number 1017, Faculty Member at SUNY Albany - jsicker@albany.edu). The team organization chart is shown below. Figure 1: Team Organization Chart The RRS mailing address is Rensselaer Rocket Society Department of Mechanical, Aerospace, and Nuclear Engineering Rensselaer Polytechnic Institute 110 8th St Troy, NY Launch Vehicle Summary The launch vehicle will be inches with a body diameter of approximately 4 in. The rocket will have a mass equal to 230 ounces. The launch vehicle will be propelled by an AeroTech K828 motor for a 54mm motor mount. The recovery system will consist of a main and a drogue parachute that shall be deployed via electronic 5

7 deployment. This deployment will be controlled by a set of two redundant altimeters. The primary deployment altimeter is the Perfectflite Stratologger SL100, a barometric pressure based altimeter. The secondary altimeter is the Featherweight Altimeters Raven3, which uses barometric and acceleration readings to measure altitude. Each altimeter has an independent power supply, and is connected only to its set of ejection charges. Additional details of the recovery system can be found in section 3.2. See Appendix A for the Milestone Review Flysheet. 1.3 Payload Summary The proposed payload design will attempt to complete the requirements of the challenges outlined in section 4.4 of the 2018 NASA Student Launch Colleges, Universities, Non-Academic Handbook. The proposed payload design complies with NASA s requirements outlined in the 2018 NSL Handbook section 4.4 in that it will include an on board camera system designed to identify and differentiate between 3 different colored tarps. Two Pi Cam v2 camera modules will be used to capture the images, which will be analyzed in real time by a custom designed software package run on two Raspberry Pi 3 Model B s. A photograph with the colored tarps outlined will qualify as a successful experiment. 2 Changes Made Since Proposal Since the proposal, there have been no significant changes to the design of the lift vehicle, recovery system, payload, or project plan of the RRS for this competition. However, the RRS has gone into greater detail in the planning of each system. 6

8 3 Vehicle Criteria 3.1 Selection, Design, and Rationale of Launch Vehicle Mission Statement The launch vehicle will safely power the payload to an apogee of one mile and then safely return to the ground by a dual-deploy recovery system Requirements and Mission Success Criteria The rocket will reach an altitude of approximately 5,280 ft. This requirement will be successful if the vehicle reaches apogee between 5,000 ft and 5,600 ft. Altitude will be measured by two barometric altimeters, one Perfectflite Stratologger SL100 and one Featherweight Altimeters Raven 3, and will be reported via audible beep post flight. The Stratologger SL100 will be the official altimeter for scoring purposes. The launch vehicle will be reusable with fewer than four independent, separable sections. These requirements will be successful if the rocket is able to be prepared for a re-launch immediately after landing, and if the rocket design utilizes four or less independent, separable 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 not exceed L-class 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 selected is approved by the NAR and TRA, does not exceed L-class, and the igniter is capable of being activated using a standard 12-volt DC firing system. The recovery system of the launch vehicle will be electronic dual deploy with a drogue parachute deployed at apogee and a main parachute deployed at a much lower altitude 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 by the careful selection of parachutes. Additionally, the launch vehicle must be 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 dedicated power supply and arming switch. These requirements will be met by the launch vehicle design Subsystems and Design Alternatives The launch vehicle consists of three subsystems: upper airframe, lower airframe, and recovery. The recovery subsystem is covered in detail in Section 3.2. For these subsystems, there are several components involved. For each of the components, the vehicle design team examined viable design alternatives and their associated merits and drawbacks Upper Airframe The upper airframe refers to vehicle components located in front of the separation point for drogue parachute deployment. In particular, the vehicle design team considered 5 main components in the upper airframe: the nose cone, two upper body tube sections (one to contain the parachute and one to contain the tracker), switch band, and coupler tubes. 7

9 The alternatives for nose cone design largely reduced to two factors: general shape and component material. In terms of nose cone shape, the design team considered conical, ellipsoid, and tangent ogive. Shape consideration was limited to these shapes because most other nose cone shapes are optimized for transonic and supersonic flights. According to research and simulations, ellipsoid and tangent ogive shapes perform comparably (all else held constant), while the conical nose cone shape did not perform as well. The benefit of choosing a conical shape is its relative ease of manufacturing; however, the RRS did not consider manufacturing the nose cone in house, due to project time constraints, effectively nullifying the advantages of a conical nose cone. In terms of component material, polypropylene plastic, G12 (also known as filament-wound) fiberglass, and hand-laid fiberglass were considered. Since the vehicle is not projected to enter transonic speeds, a plastic nose cone would offer sufficient structural strength while contributing far less to the total mass of the rocket; however, reducing the mass of the nose cone moves the vehicle s center of gravity lower, effectively reducing stability. Selecting a G12 or hand-laid fiberglass nose cone would give the component additional strength, with G12 being stronger and more consistent than a hand-laid design, though both options are significantly stronger than polypropylene plastic. The additional mass of either fiberglass option would also give more vehicle stability compared to a plastic nose cone. Another benefit to this material section is that fiberglass nose cones commonly have an open bottom design, allowing for recessed bulkhead placement for parachute packing, though this can be easily solved by removing the integrated bulkhead from a plastic nosecone. However, both G12 and hand-laid fiberglass are significantly more expensive than polypropylene, and given that the flight profile does not require a stronger nose cone, the additional cost of a fiberglass nose cone would not be justified. The choices for alternatives in the two upper body tube sections and switch band components were largely the same, as the parts would have to be compatible with each other. The factors considered included material strength, workability, and cost. The four viable choices in material included cardboard, phenolic resin, reinforced phenolic resin, and fiberglass. The considerations of each material against each factor are shown in Table 1. Table 1: Body tube Material Selection Matrix G12 Fiberglass Reinforced Phenolic Resin Phenolic Resin Kraft-paper Cardboard Failure Mode Compressive Strength Workability Weight Cost Total The above selection matrix is based on a 1-5 scale, where the best material(s) for a particular criterion is graded with a 5, and the worst material(s) for a particular criterion receives a 1. The failure mode is judged based on how a material tends to fail, and how localized damage tends to stay. Workability is judged based on access to necessary tools and techniques, and the likelihood to damage the material during work. Although G12 fiberglass is by far the strongest material considered, it is also the heaviest and by far the most expensive. Additionally, the vast majority of machine shops at RPI do not allow students to work with fiberglass materials. Reinforced phenolic resin offers similar strength to G12 fiberglass, but remains at a high cost, and due to the added thickness of the reinforcement, it would pose additional aerodynamic challenges. Though it would be possible for the RRS to create its own reinforced phenolic resin using either kevlar or fiberglass composites, the extra time and effort involved introduces a large risk to project completion. Standard phenolic resin provides a good balance between sufficient strength, relatively low weight, and low-cost. Kraft-paper cardboard is not expected to be strong enough to withstand the loads placed on the vehicle during flight without reinforcement. Other considerations made for the body tubing included the general dimensions of each section. This largely tied into desired ease of payload integration and overall vehicle length. The vehicle design team decided to create a vehicle with constant diameter along the vehicle length for ease of manufacturing and aerodynamic analysis. As the diameter of the vehicle components increases, more space becomes available for the components of the recovery and payload sections; however, the mass, drag, and cost of the vehicle body tubes also increases. The design team largely considered 3, 4, and 6 airframe diameters. A smaller airframe diameter has the advantage of requiring a significantly smaller (and thus less expensive) 8

10 motor to achieve the same altitude performance, as well as being cheaper in materials costs for the airframe, nose cone, fins (which would not have to be as large for a smaller launch vehicle), and other components. Additionally, a smaller airframe diameter is better suited to the smaller motors the RRS is looking at this year. Alternatives for the coupler sections in the vehicle design are largely constrained by the selection for the vehicle body tubes as well as a desire to procure body tube and couplers from the same vendor. For material selection, the vehicle design team considered the same factors and variable choices as in Table 1. The design team also considered how many couplers would be needed for the vehicle. One coupler is needed at the separation point for the drogue parachute. Adding an additional coupler and splitting the upper airframe into two body tube sections allows for easier transportation, assembly, and payload integration Lower Airframe The lower airframe refers to vehicle components located aft of separation point for drogue parachute deployment. Specifically, the vehicle design team considered 3 main components in the lower airframe: vehicle fins, motor mount assembly, and lower body tube. The alternatives for vehicle fins reduce to several factors including material selection, fin thickness, general shape, and surface area. Fin material options considered include birch plywood and G10 fiberglass. Plywood is an inexpensive and lightweight option, but does not offer the strength necessary to withstand projected flight speeds, without increasing thickness to the point where the weight advantage is nearly negated, and the drag increases well beyond the level of G10. Plywood fins are also much more susceptible to warping and general damage. Fiberglass fins offer the rigidity and strength necessary for this vehicle. Though significantly more expensive, G10 fiberglass fins are still within the RRS budget for this project. As fin thickness increases, the fins become stronger and drag increases. The design team sought to balance these two properties. General shape considerations were limited to trapezoidal fin sets for ease of modeling and analysis. Including a backwards sweep in the shape, as well as finding the optimal length for the tip chord were found to reduce overall drag. Surface area of the fins is closely related to the general shape, and larger surface area fins produce more drag. Design alternatives for the motor mount assembly reduce to a few considerations: the size of the motor mount, motor mount material, centering ring material, and reinforcements. The diameter of the motor mount is entirely dictated by the motor the design team selects. In general, motors do not exist in 38mm diameter that are sufficiently powerful to achieve the required performance. This limited the selection of motor mount diameters to between 54mm and 98mm. However, a 75mm or 98mm motor is excessive for the demands of this competition and the selected launch vehicle, which suggested the use of a 54mm motor and thus 54mm motor mount. The motor mount material was subject to the similar considerations as the vehicle body tube sections, which are summarized in Table 1. The centering ring materials considered included 1 4 birch plywood and G10 fiberglass. Plywood offers sufficient strength, while being easier to work with and much more affordable. Fiberglass offers increased strength, but is significantly more difficult to machine and much more expensive. In addition to these sub-components, the motor mount assembly could include additional reinforcements. Reinforcements considered include fillets along the edge between the motor mount and fins, as well as additional fiberglass weave composite. The lowermost body tube section was subject to the considerations summarized in Table 1. Additionally, the slots for the fins have the option of being extended through the bottom of the airframe. This allows for the RRS to attach the fins to the motor mount assembly before it is inserted into the airframe, which in turn allows for additional reinforcements to be easily made to the motor mount, though these could also be made by not installing the aftmost centering ring until after internal reinforcements have been made Current Vehicle Design The current vehicle design is comprised of the best choices for each alternative subsystem design. Figure 2 and Table 2 show the current vehicle model. The overall vehicle diameter selected was 4 in. Based on availability from the main supplier the RRS has selected, Public Missiles Ltd (PML), the vehicle will have a standard interior diameter of 3.90 in. and a exterior diameter of in. This diameter was selected primarily for ease of integration with the payload system. A 9

11 Figure 2: Current Vehicle Design Table 2: Vehicle Details Component Outer Diameter Inner Diameter Total Length Total Weight Fin Material Nose Cone Material Body Tube Material Motor Mount Booster Length AV Bay Length Nosecone and Tracker Bay Length Specification inches 3.90 inches inches lb G-10 Fiberglass Polypropylene Plastic Phenolic Resin 54mm 30 inches 32 inches (28 inches exposed) inches (16.75 inches exposed nose cone, 12 inches exposed phenolic tubing) larger diameter affords the payload team more space to work with for system design, integration, and assembly, though this was balanced against the need to minimize cost of both airframe and motor. The design team selected a polypropylene nose cone with a diameter matching our vehicle selection, and a in. exposed length. This nose cone was selected primarily due to its ease of availability from the same manufacturer as the airframe. Components from the same manufacturer often have the best fit. The nose cone will be secured to the tracker payload section using small machine screws. All body tubes will be made of phenolic resin material supplied by PML. The benefits of this material selection are affordable price, ease of workability, higher compressive strength, and a better failure mode than traditional cardboard. PML claims that their phenolic resin has up to five times the compressive strength of kraft-paper tubing. Additionally, the brittle nature of the phenolic resin keeps failure points localized, as opposed to the cascading crumpling failures of cardboard. The length of each body tube section can be found in Appendix B. These lengths were selected based on a balance between necessary packing space for parachutes, and necessary lengths for electronic components, and minimizing length to minimize overall weight, drag, and cost. The current vehicle design utilizes two coupler sections. Both couplers will be made of the phenolic resin material supplied by PML. The tracker section will be attached to the main parachute section using one 8 section of coupler. This coupler will be permanently affixed to the tracker section using epoxy, and attached to the main parachute section using shear pins. The aft end of this coupler will have a 1 4 aircraft ply bulkhead, with the anchor point for the main parachute harness. The second coupler will be 12 long and attached centrally, using epoxy, to the switch band (leaving 4 exposed on each end). This coupler section will contain the avionics and primary payload. The ends of this coupler will have two 1 4 stepped bulkheads to allow for easy access and sealing of the electronics bay. The forward end of this coupler will be secured in flight to the aft end of the main parachute section using small machine screws, and the aft end will be secured to the booster section using shear pins. The selected fins are in. thick and made of G10 fiberglass. This combination offers sufficient strength to withstand the forces expected during launch. The fins have a trapezoidal shape with a minimized tip chord and forward sweep on the trailing edge, which is found to minimize drag and prevent damage on landing. A diagram of the fins is shown below in Figure 3. The motor mount will be phenolic resin for reasons discussed above. Due to motor considerations as outlined in Section 3.1.7, the motor mount will be 54mm on the inner diameter. The current selected centering rings are 10

12 Figure 3: Selected Shape of Fin 1 4 thick birch plywood rings that will fit between the OD of the motor mount and the ID of the body tube. Three centering rings will align the motor mount in the airframe, and sandwich the through-the-wall fin tabs Motor Selection The RRS examines several factors when deciding appropriate motors. Motor choices must provide simulated performance that meets all requirements detailed in the statement of work. Particular care was given to follow the requirements in items from the statement of work shown in Table 3 below. The overall vehicle diameter selected was 4 in. Based on availability from the main supplier the RRS has selected, Public Missiles Ltd (PML), the vehicle will have a standard interior diameter of 3.90 in. and a exterior diameter of in. This diameter was selected primarily for ease of integration with the payload system. A larger diameter affords the payload team more space to work with for system design, integration, and assembly In addition to these requirements, consideration was also given to suggested retail price and apparent ease of assembly. From past experience, the RRS has found that motors and kits are often available at discounted price from vendors at launch events. However, we will use MSRP as our consideration point and budget estimate as such discounts cannot be guaranteed until much closer to actual purchase time. With these points in mind, the vehicle design team found two potential motors to use: the AeroTech K828 and the AeroTech K480. The Aerotech K828 is a reloadable 23% K-class motor. It has a thrust curve shown in Figure 4 below. As shown in the thrust curve, the motor burns for 2.50 s and achieves a peak thrust of 1300 N. The complete motor specifications are listed in Table 4 below. The estimated retail value for a complete motor kit (which includes the motor, case, hardware, and nozzle) is $350. The estimated retail value for each additional reload is $140. The Aerotech K480 is a reloadable 77% K-class motor. It has a thrust curve shown in Figure 5 below. As shown in the thrust curve, the motor burns for 4.3 s and achieves a peak thrust of N. The complete motor specifications are listed in Table 5 below. The estimated retail value for a complete motor kit (which includes the motor, case, hardware, and nozzle) is $350. The estimated retail value for each additional reload is $

13 Table 3: Statement of Work Items Relevant to Motor Selection Item Statement Compliance 1.11 The launch vehicle shall use a commercially available solid motor propulsion system using ACPC which is approved and certified by NAR, TRA, or CAR 1.13 The total impulse provided by the launch vehicle shall not exceed 5,120 Newton-seconds (L-class) 1.14 The launch vehicle shall have a minimum static stability margin of 2.0 at the point of rail exit 1.15 The launch vehicle shall accelerate to a minimum velocity of 52 ft/s at rail exit The launch vehicle shall not utilize motors that expel titanium sponges Figure 4: Thrust Curve for Aerotech K828 All motors considered will be commercially available ACPC motors with proper approval and certification All motors considered will have a maximum total impulse of 5,120 Ns High value will be given to lighter motors that offset the CG less, and simulations will be run to verify the static stability margin at least 2.0 cal at rail exit Motors considered will have enough thrust to propel the launch vehicle to at least 52 ft/s at rail exit in simulations Motors with sparking propellant of any kind will not be considered Table 4: AeroTech K828 Specifications Burn Time Total Impulse Average Thrust Peak Thrust Launch Mass Empty Mass Diameter Length 2.50 s 2120 Ns 828 N 1300 N 78.4 oz 30.0 oz 2.13 in 22.8 in The two motors share very similar performance profiles. However the AeroTech K480 overshoots the target apogee by approximately 550 feet. Additionally the K828 has the better thrust to weight ratio. Currently, the RRS has selected the AeroTech K828 as our vehicle motor. The size allows for easy testing, 12

14 Figure 5: Thrust Curve for Aerotech K480 Table 5: AeroTech K480 Specifications Burn Time Total Impulse Average Thrust Peak Thrust Launch Mass Empty Mass Diameter Length 4.3 s Ns N N 73.3 oz oz 2.13 in 22.8 in the thrust to weight ratio is better and the simulations reach the target apogee.table 6 shows the current mass statement of the vehicle design. Table 6: Vehicle Mass Part Mass (oz) Quantity Subtotal (oz) Nose Cone Tracker Tracker Bay Tubing Tracker Bay Coupler and Bulkhead Main Parachute Bay Main Parachute and Hardware Payload and Avionics Bay Booster Airframe Drogue Parachute and Hardware Motor Mount Tube Fins (x3) Centering Rings (x3) Motor Subtotal Verification 13

15 Table 7: Requirement and Verification Plan Read apogee at several launches in different weather conditions Requirement Design Feature Apogee between Rocket 5,280 Mass, ft and 5,600 Rocket ft if left Motor, unaltered Design Reusable Body Strength (fins, airframe, parachutes, etc.) Four or fewer Independent Sections Single Stage Commercially available solid motor propulsion system not exceeding Class L Capable of Launch by 12 V DC firing system Minimum static stability margin of 2 at rail exit General Design Design Phase OpenRocket Simulations, Payload Mass, Rocket material selection, Motor Selection Stress Analysis of Vulnerable Components, Use of Large Factors of Safety in Critical Components Design fewer than four independent Sections (Nosecone and tracker section, Payload/Avionics Section, Lower Body/Booster Section) Design only a single stage rocket General Design, Motor Design Motor Ensure Mass and General Rocket Design allow for a motor not exceeding Class L approved by the TRA and NAR Motor, Motor Retainer Motor, Vehicle Design Select a motor retainer that allows for access to motor, select a motor able to be launched with 12V DC firing system Use simulation data to ensure placement of vehicle CG and CP creates a static stability margin of at least 2 ca Construction Phase Stay as close as possible to original Mass Estimates Use High-Strength Epoxies, Store all components Safely, Design Followed, Components inspected upon order arrival show no signs of damage Design Followed Design Followed Design Followed, stay as close as possible to original mass estimates Design Followed, safely store motors Design Followed, stay as close as possible to original mass estimates, update simulation profiles to reflect any change Testing Phase Launch in appropriate weather multiple times Rocket is launched and recovered in less than four independent sections Rocket is launched with only a single stage motor Motor not exceeding Class L is used at several launches in different weather conditions Launch rocket on standard 12 V DC firing system multiple times Launch rocket several times to ensure stability off the launch rail 14

16 Table 8: Requirement and Verification Plan Part 2 Requirement Design Feature Minimum Motor, Vehicle velocity of Design 52 f/s at rail exit Electronic Dual Deploy Drogue Deploys at Apogee, Main Parachute Deploys at much lower altitude Shear Pins hold rocket sections together until Parachute Deployment Independent Sections have less than 75 ft-lb of KE at Landing Redundant, Safe Altimeters General Recovery Design (parachutes, shock cord, ejection charges, etc.) Parachutes, ejection charges, General Rocket Design, Altimeters Shear Pins, Ejection Charges General Rocket Design (Mass), Parachutes Altimeters, Supporting Recovery Electronics Design Phase Use simulation data to select a motor with sufficient thrust to propel the rocket off the rail with a velocity of at least 52 f/s General Design is Dual-Deploy, Use of a Drogue and Main Parachute, Design for Multiple Separation Points Select at least 2 commercially available altimeters, Design has an independent power supply to each Altimeter Construction Testing Phase Phase Design Followed, Launch rockets stay as close as several times, use possible to original flight data to ensure mas estimates, minimum exit update simulation velocity profiles to reflect any change Design Followed Rocket launched with appropriate recovery system, Drogue and Main parachutes deploy at different points in flight Design Followed, Set Altimeters to required ejection charge deployments (and keep track of which side is deployed when) Design Followed, Shear Pins show no signs of damage before install Design Followed, stay as close as possible to original mass estimates, Parachutes and Recovery system components show no signs of damage upon order arrival Design Followed, Altimeters stored safely, All components of supporting electronics and altimeters tested before installation General Design has Drogue Parachute at First Separation Point and Main Parachute at another, Simulations run with required deployment locations Shear Pin Strength Accurately Calculated and inserted into Recovery Design, Ejection Charge Strength Accurately Calculated Kinetic Energy of Each Independent Section Analyzed Accurately Rocket launched multiple times with parachute deployment at required times Rocket launched multiple times with parachute deployment at required times No damage or hard landing evident in any section after multiple test launches in varied flight conditions Recovery System Operates as Expected, Altimeters report similar apogees in all test launches 15

17 3.1.9 Component Analysis Table 9: Component Analysis Component Description Materials Cost Source Fin Can Fins attached to motor tube G10 Fiberglass + epoxy + Phenolic Motor Tube $36 publicmissiles.com/products/ Lower Airframe Airframe containing Phenolic Airframe Tubing $50 publicmissiles.com/products/ Fin Can and Drogue Parachute Nosecone Nosecone and Main High-Strength Plastic + $10 publicmissiles.com/products/ Parachute Mount forged steel eye bolt + Main Airframe Airframe containing Main Parachute and Payload Section Payload Section Containment system for payload (including bulkheads, couplers, etc.) Recovery Electronics Altimeters and supporting electronics that run recovery system Ejection Charges Charges that eject Parachutes on Altimeter epoxy Phenolic Airframe Tubing $50 mcmastercarr.com Plywood Bulkheads, Phenolic Cardboard Couplers, forged steel eye bolts, steel threaded rod, epoxy, acrylic plastic G10 Fiberglass Sled, Altimeters, Batteries, copper wiring, switches Aluminum Blast Cap holders, Ejection Charges, epoxy, copper wiring command Parachutes Drogue and Main 60 Skyangle main and 24 Rocketman Ballistic Mach II drogue Motor Retainer Tee nuts, screws, and washers $50 publicmissiles.com/products/ $50 sparkfun.com $50 apogeerockets.com $25 b2rocketry.com Steel $42 mcmaster.com 16

18 3.2 Recovery Subsystem Altimeters The three design alternatives for the altimeters are the Raven3, Missleworks RRC2+, and Stratologger SL100. Table 10 outlines the specifications of each altimeter. The Raven3 is the most expensive but is compact and fairly accurate. The Missleworks RRC2+ is cost effective, but it is lacking key specifications such as accuracy and max height. The SL100 is another cheap option as well as the most accurate. Given the club already has all three altimeters purchased, cost is not an issue. All altimeters have a similar power requirement and sample rate. The main difference between the altimeters is the size and accuracy. Given the Missleworks RRC2+ is missing accuracy data, this alternative will not be chosen. Given previous past experience with the Raven3 and Stratologger SL100, these two altimeters will be used. The Stratologger will be chosen as the primary altimeter and the Raven3 will be used as the secondary altimeter. Table 10: Altimeter Comparison Featherweight Raven 3 Missleworks RRC2+ PerfectFlite Stratologger SL100 Price [Dollars] Width [in] Length [in] Thickness [in] 0.55 N/A 0.50 Altitude Accuracy [%] +/- 0.3 N/A +/- 0.1 Operating Voltage [V] Sample Rate [Hz] Maximum Altitude [ft] 100,000 40, , Parachutes The design alternatives for the main parachute are the SkyAngle Classic 60 and Classic II 60 parachutes. Both parachutes have the same drag coefficient and surface area. The Classic II is heavier, 18.2 oz versus 10 oz. However, the Classic II is made from zero-porosity material and has a stronger line attachment - a nickel plated swivel rated for 1500 lbs versus 1000 lbs. Given the club already has Classic II parachutes in stock and has had past success with such chute, the Classic II will be used. The drogue will be a 2 foot Ballistic Mach II, chosen for its strength and positive past experience. OpenRocket simulations have been run with these parachute sizes verifying a safe descent Attachment Hardware and Bulkheads The various materials used for shock cords would be elastic, kevlar, and tubular nylon. Elastic shock cord is too weak for a rocket this size. Kevlar and tubular nylon are both strong materials. Tubular nylon shock cord will be used given positive past experience with this shock cord, in 9 16 width to provide sufficient strength for the mass and projected ejection forces of the launch vehicle. Another key component to the attachment hardware is the eye bolts. Eye bolts come in multiple materials, sizes, and forms. Stainless steel resists corrosion well and therefore, will be used. A forged, closed, shouldered eye bolt can handle the most stress. Given a max acceleration of 317 ft s at main deployment and a mass of lbs, the max force experienced by the eye bolt would be 191 lb f. A eye bolt can withstand vertical loads between lb s. This gives a factor of safety of 2.3 to 2.6. The determining factors for the bulkheads are the material and the dimensions. Bulkheads can be made from fiberglass or plywood. Due to past experience with plywood bulkheads, this material will be used again for the bulkheads. In terms of sizing, RRS uses PML to order body tubing. PML only offers bulkheads for the given diameter of the rocket at 3 16 thick. RRS will purchase a quarter inch thick sheet of aircraft ply plywood 17

19 to manufacture bulkheads from, to provide adequate thickness to ensure a larger factor of safety than the 3 16 bulkheads provided by PML Electronics In terms of electronics, the avionics bay will house two altimeters dedicated solely to parachute deployment: the Raven3 and Stratologger SL100. Each altimeter will have its own independent ejection charges and battery packs. In the case one altimeter loses power or ejection charge fails, the second altimeter will be in place to deploy the parachutes. Figure 6 below shows an electronics schematic for the two altimeters. Figure 6: Recovery Electronics Schematic Blast caps will be mounted on the fore and aft bulkheads of the avionics bay. At the designated altitudes, the altimeter will trigger the charges, separating the body sections, allowing the parachute to unfurl. The drogue will deploy at apogee, the first ejection event, and the main will deploy at about 500 feet, the second ejection event. Assuming a pressure of 12 psi, the main ejection charges will require about 2 g of black powder while the drogue ejection will require 1 g of black powder. Nomex parachute protectors will be used to protect the parachutes from heat during charges firing Tracking Tracking will be performed by a 900 MHz xbee radio, Arduino Uno, GPS reciever, and 9v battery secured in the tracker bay in a similar manner to how the recovery and payload electronics are secured in the avionics bay. This ensures electrical isolation from the recovery electronics. The ground station will consist of a laptop computer, second xbee radio with an external antenna, and a xbee to USB adapter. The ground station will receive location and altitude data from the tracker at a 1 Hz rate, and will use a custom piece of software to plot and display this data. 3.3 Mission Performance Predictions Kinetic Energy Given an estimated mass of lb f and assuming a 60 SkyAngle Classic II main and a 2 foot Ballistic Mach II drogue, the final descent velocity from OpenRocket calculations is 18.5 ft/s. Using kinetic energy equation, the max KE of the entire rocket is ft-lb f. Table 11 shows the estimated KE for each section of the rocket. All body sections meet the 75 ft-lb f requirement Drift Analysis Table 12 lists the calculated drift distances with varying wind speeds given a descent time of seconds found using OpenRocket. The drift values were calculated for a zero degree launch angle, with the assumption that the rocket will drift at the same velocity as the wind while in the air. 18

20 Table 11: Kinetic Energy Estimate for Body Sections Component Mass [oz] Kinetic Energy [ft-lb f ] Nosecone, Tracker Body Tube Payload Bay, Switch Band Booster Table 12: Calculated Drift Distances for Varying Wind Speeds Wind Speed (mph) Drift Distance (ft) Safety 4.1 Safety Officer RRS has identified Sierra Koch as the club Safety Officer. Her responsibilities include ensuring shop safety and the proper handling and storage of hazardous materials. She will oversee the safe construction, launch, and recovery of the vehicle through inspections, supervision, and review of club plans and actions. She will ensure that all participating RRS team members complete the lab safety training administered by the RPI School of Engineering prior to performing any construction or using the lab space. 4.2 NAR/TRA Mentor The RRS s mentor, John Sicker, is a level 2 certified member of the NAR and level 3 certified member of the TRA. He will provide guidance for safe and proper high-powered rocketry practices and will be responsible for all safety inspections. The RRS s Safety Officer, Sierra Koch, will accompany Mr. Sicker on all such inspections. For liability purposes, Mr. Sicker is designated as the owner of the proposed rocket. 4.3 Risk Assessment The following section identifies possible hazards, their causes, and the resulting effects, in addition to providing mitigations and controls for each hazard. Each hazard is assigned a Pre-Risk Assessment Code (Pre-RAC) and a Post-Risk Assessment Code (Post-RAC), based on each hazard s likelihood and severity. Risks with Post-RACs that are categorized as High or Medium risk are unacceptable and undesirable, and every effort will be made to mitigate the risk and reduce the Post-RAC to a Low or Minimum Risk level. 4.4 Preliminary Personal Hazard Analysis The following section outlines likely hazards and their mitigations. This includes material and facility hazards Material Hazards Material hazards include epoxies, resins, solder, and any other chemical that can cause burns, skin irritation, and/or may have risks associated with fumes. In order to reduce and mitigate these risks, the Safety Officer 19

21 will ensure: an orderly workspace, familiarity with MSDS, members wearing appropriate clothing and PPE, and knowledge of first-aid stations and fire extinguishers. Table 4-3 below outlines some of the potential material hazards that the RRS team may be exposed to during the project Facility and Tool Hazards Tools and machines such as lathes, mills, presses, and other equipment found in lab spaces present risks in the form of cuts, bruises, broken bones, and other injuries. Mitigations include having a single, experienced operator on a tool at a time with a spotter; minimization of distraction during tool operation; refraining from wearing loose clothing while operating tools; touching of tools or machines only after they have stopped spinning or have cooled down; proper tool maintenance and handling; and knowledge of locations of first aid stations and fire safety equipment. Independent investigation of lab safety is conducted by the RPI School of Engineering, RPI Facilities Management, and the Troy Fire Marshals on an annual basis. The potential hazards found in the lab and in the field for team members are identified in Table 4-4 below. 4.5 Preliminary Failure Modes and Effects Analysis (FMEA) The following section outlines the launch and flight failure modes that have been identified. These failure modes include motor retainer failure, launch vehicle instability, recovery system failure, and others. Each failure mode has a following mitigation and has been assigned a Pre and Post RAC. 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. 4.6 Preliminary Environmental Concerns Adverse weather conditions such as heavy rains, high winds, or electrical storms may inhibit test launches. This risk is mitigated by scheduling subscale launches before winter weather conditions inhibit test launches and by setting flexible test launch windows for all launches so that launches are easily rescheduled in the event of unforeseen adverse weather conditions. The inability of Mr. Johnson to be present at all test launches and at Huntsville, AL, can be mitigated by ensuring that there is at least one team member present at all launches and at Huntsville who is NAR or TRA certified, as well as with Mr. Sicker s identification of an appropriate certified substitute, if necessary. Damage to the rocket during test launches or the final competition can be mitigated by having spare parts on hand, by constructing the rocket with standard sized, easily-obtainable components, and with thorough engineering analysis to be completed before launch. The concerns that may impact the testing and launch of the rocket have been identified in Table Project Risks Risks to project completion have been identified in Table 4-8. They have been addressed with a mitigation plan and their likelihood and impact is ranked low to high. 20

22 Table 4-1: Color Coded Risk Assessment Codes Severity Likelihood 1 Catastrophic 2 Critical 3 Marginal 4 Negligible A Frequent 1A 2A 3A 4A B Probable 1B 2B 3B 4B C Occasional 1C 2C 3C 4C D Remote 1D 2D 3D 4D E - Improbable 1E 2E 3E 4E Severity- Probability Table 4-2: Different Levels of Risk and Their Acceptance Levels Acceptance Level/Approving Authority High Risk Medium Risk Low Risk Minimal Risk Unacceptable. Documented approval from an independent safety panel, such as the NAR Board, or similar required. Undesirable. Documented approval from the safety officer and/or team mentor required. Acceptable. Documented approval from the system lead or higher-level authority required. Acceptable. Documented approval not required, but an informal review by the system lead is highly recommended. 21

23 Table 4-3: Material Hazards Hazard Cause Effect Pre - RAC Inhalation of Use of certain materials fumes, such as cause the release of paint fumes. dangerous fumes. Skin contact with epoxy, black powder, or other hazardous chemicals. Poor ventilation leads to increased risk. Using or working with a material without the use of PPE such as gloves. Poor knowledge of the chemical or material in use. Inhaled fumes cause damage to lungs and eyes. Presence of fumes can also irritate skin. Skin contact with hazardous chemicals can cause irritation or chemical burns. Improper disposal or cleanup can cause risk in the future. 1B 3B Mitigation Post - RAC Team members will work in wellventilated and spacious areas. Team members will be knowledgeable and wellinformed about the use of MSDS and PPE. Gloves will be worn when working with hazardous chemicals at all times. Team members will be knowledgeable and wellinformed about the use of MSDS and PPE. 3D 4C After use of hazardous materials, the surrounding area as well as any tools used will be well cleaned. Premature propellant or black powder ignition. Launch system failure. Miswired onboard avionics. Avionics failure. Premature ignition can cause severe burns on personnel or damage the vehicle or surrounding area. 1C Motor will be stored in a separate, insulated cabinet away from sources of ignition. Motor will be handled with gloves at all times. 2E Ground tests will be performed to ensure the safety and functionality of the onboard avionics. Vehicle Debris. Catastrophic rocket motor malfunction. Recovery system failure. Unexpected impact during flight. Falling vehicle debris could seriously injure team members or onlookers. Failure of recovery system could lead to dangerous projectile ( lawn dart ). 2D NAR and TRA range regulations will be followed and obeyed at all times. Team members will be trained to watch the skies and avoid falling vehicle debris. 4D The rocket will not be launched in unsafe conditions under any circumstances. 22

24 Ignition Faulty launch switch. system failure; failure to Improper use of launch ignite. system. Catastrophic rocket motor malfunction ( CATO ) Faulty launch system. Motor Failure. Improper igniter positioning or improper ignition. Error in manufacturing. An untrained team member could approach the rocket before it has been deemed safe and experience a launch or failure at a proximity that increases risk of serious injury or death. A CATO could lead to a fire which endangers nearby personal or the environment. A CATO could lead to serious injury or death via shrapnel or burns. 2C 1B NAR regulations require the use of an electrical launch system and electrical motor igniter. Team members will be trained not to approach a rocket on the pad after a failed ignition until sixty seconds have elapsed and it is deemed safe by the RSO. Team members will maintain a safe distance from the launch pad at all times. The RRS will use only commercially available motors to lessen the risk of motor failure. 4C 2E Injury due to unstable rocket flight. Improper motor assembly. Rocket stability is less than 1.0 cal. CP or CG improperly marked. Outdated simulations lead to false stability calculations. Poor static or dynamic rocket stability could lead to undesired or dangerous flight paths or launches. Personnel may be at risk during launches where rockets do not have the necessary stability required for launch. 2C Team members will be briefed on the proper handling and storage procedures of motors. The RRS will properly mark the CP and CG. The RRS will perform simulations regularly with the most current parameters to ensure that a desired stability margin is maintained. 2E 23