NASA USLI Preliminary Design Review (PDR) Rensselaer Rocket Society (RRS)

Transcription

1 NASA USLI Preliminary Design Review (PDR) Rensselaer Rocket Society (RRS) Rensselaer Polytechnic Institute 110 8th St Troy, NY Project Name: Andromeda Task 3.3: Roll Induction and Counter Roll Friday, November 4 th,

2 1. Table of Contents Table of Contents List of Figures List of Tables List of Acronyms 2. Executive Summary 2.1 Team Summary 2.2 Launch Vehicle Summary 2.3 Payload Summary 3. Changes Made Since Proposal 3.1 Vehicle Changes 3.2 Payload Changes 3.3 Project Plan Changes 4. Vehicle Criteria 4.1 Selection, Design, and Verification of Launch Vehicle Mission Statement Requirements and Mission Success Criteria Subsystems and Design Alternatives Upper Airframe Lower Airframe Current Vehicle Design Motor Selection Verification Plan Risk Analysis Component Analysis 4.2 Recovery Subsystem System Level Design System Components 4.3 Mission Performance Predictions Mission Performance Criteria Vehicle Simulations Kinetic Energy Analysis Drift Analysis 4.4 Safety Safety Officer Preliminary Checklists of final assembly and launch procedures 2

3 Preliminary Personal Hazard Analysis Material Hazards Facility and Tool Hazards Preliminary Failure Modes and Effects Analysis Preliminary Environmental Concerns Project Risks 5. Payload Criteria 5.1 Selection and Design of Payload Roll-Blade Subsystem Microprocessor System Integration with Vehicle Subsystem 5.2 Payload Concept Features and Definition 5.3 Science Value 5.4 Verification of Payload System Level Verification Defense of Current Leading Design 6. Project Plan 6.1 Budget Plan 6.2 Funding Plan 6.3 Timeline 6.4 Educational Engagement Status Soda Bottle Rockets and Open House National Manufacturing Day Appendix Appendix A: Example Milestone Review Flysheet Appendix B: Structural Design Assembly Drawings Appendix C: Detailed Budget 3

4 List of Figures Figure 4-1: Current Vehicle Design...12 Figure 4-2: Selected Fin Shape in OpenRocket...13 Figure 4-3: Thrust Curve for Animal Motor Works L Figure 4-4: Thrust Curve for Cesaroni L Figure 4-5: Recovery System Electronics Schematics...31 Figure 4-6: Model for Flight Simulation (Launch-Ready Motor)...32 Figure 4-7: Model for Flight Simulation (After Motor Burnout)...33 Figure 4-8: Flight Simulation from OpenRocket...33 Figure 5-1: "Roll-Blade Subsystem"...44 Figure 5-2: Complete Payload System Render...46 Figure 5-3: Payload Integrated with Vehicle...47 Figure 5-4: NASA NLF15 Airfoil Profile...49 Figure 5-5: Reynold's Number vs Altitude During Cruise...49 Figure 5-6: Shape of Cam Track...50 List of Tables Table 2-1: Payload Timeline...7 Table 4-1: Body Tube Material Selection Matrix...10 Table 4-2: Statement of Work Items Relevant to Motor Selection.14 Table 4-3: Specifications of AMW L Table 4-4: Specifications of Cesaroni L Table 4-5: Mass Statement of Current Vehicle Design...18 Table 4-6: Requirement and Verification Plan...19 Table 4-7: Risk Assessment Code Matrix...21 Table 4-8: Definition of Risk Levels...21 Table 4-9: Definition of Risk Impact Values...22 Table 4-10: Definitions of Risk Likelihoods...22 Table 4-11: Possible Risks and Mitigation Tactics...23 Table 4-12: Component Analysis...26 Table 4-13: Altimeter Comparison Part Table 4-14: Altimeter Comparison Part Table 4-15: Parachute Select...30 Table 4-16: Variables in Calculation of Final Descent Velocity..34 Table 4-17: Landing Kinetic Energies of Independent Sections...34 Table 4-18: Calculated Drift Distances for Varying Wind Speeds.35 Table 4-19: Pre-Launch Operations Checklist...36 Table 4-20: Launchpad Operations Checklist...37 Table 4-21: Material Hazards...38 Table 4-22: Facility Hazards...40 Table 4-23: Failure Modes...41 Table 4-24: Environmental Hazards to Project Completion

5 Table 4-25: Risks to Project Completion...43 Table 5-1: System Level Verification...48 Table 5-2: Current Leading Payload Designs...50 Table 6-1: RRS Budget Summary

6 2. Executive Summary 2.1 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 Jody Johnson (NAR Level 3 Certified, NAR #85182 SR, TRA Level 3 Certified, TRA #10973). The mailing address for the RRS 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 98 inches with a body diameter of approximately 6 in. The rocket will have a mass equal to slugs. The launch vehicle will be propelled by an Cesaroni L910 motor for a 75mm motor mount. The recovery system will consist of a main and a drogue parachute that shall be deployed via electronic 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 a Featherweight 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 4.2. See Appendix A for the Milestone Review Flysheet. 2.3 Payload Summary The payload design attempts to complete the requirements of the challenges outlined in section 3.3 of the 2017 NASA Student Launch Colleges, Universities, Non-Academic Handbook. To achieve this, the payload includes two cam systems that deploy two sets of opposing blades to induce roll and counter-roll. The Payload Module utilizes redundant gyroscopes to monitor roll and dynamically deploys blades to match an idealized rotational model. The two sets of three blades are asymmetric airfoils (NASA NLF15) with fixed angles of attack of 5 and -5, respectively. Two independent motors drive the cam, and are controlled by a central microprocessor, which deploys each set as they are needed. Following the Roll-Counter-roll phase, the microprocessor will command the motors to equally extend both sets of blades, to act as active drag control, slowing the rocket to reach the target altitude. 6

7 The Payload Timeline will proceed in the following manner: Table 2-1: Payload Timeline Phase Zeroing Phase Roll Phase Counter-Roll / Rezeroing Phase Active Drag Phase Action Post motor-burnout. Gyroscopes will detect residual roll and deploy blades to eliminate remaining angular velocity. First set of blades deployed. Angular position and velocity tracked by gyroscopes and magnetometer while deployment length is adjusted accordingly. First set of blades is retracted while second set is deployed in order to counter the induced roll. Angular position and velocity are again tracked and the deployment length is adjusted accordingly. Following Rezeroing. Both sets of blades are deployed at equal lengths in order to actively decrease vertical velocity of the rocket. The microprocessor actively monitors altitude and velocity in order to match an idealized altitude model. 3. Changes Made Since Proposal 3.1 Vehicle Changes A ballast mass of one pound was added to the nosecone to shift the center of gravity forward. Fin shape was slightly changed to further increase vehicle stability. Choice of motor was then changed from a Cesaroni L545 to a Cesaroni L910 to compensate for the increase in mass and greater drag from an increased fin surface area. 3.2 Payload Changes Updates to the payload include the addition of detail related to the motors driving the cam system and integration into the vehicle. Additional discussion of the payload system and integration can be found in sections and 5.1.3, respectively. 3.3 Project Plan Changes The project plan has been updated to more closely represent the remaining design, construction, and testing that will occur during the project. Dates of some aspects of the plan have been updated to reflect more accurate start and projected end dates. The project plan also includes much more detail in the project development of the payload sections, including stress testing of payload materials. 7

8 4. Vehicle Criteria 4.1 Selection, Design, and Verification 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 a barometric altimeter be reported via audible beep post flight. 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 separable. 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 is capable of being launched on 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 8

9 4.2. 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 4 main components in the upper airframe: the nose cone, upper body tube, middle body tube, and coupler tubes. 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 want to consider making our own nose cone. In terms of component material, polypropylene plastic and G10 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 G10 fiberglass nose cone would give the component additional strength, and the additional mass would give more vehicle stability. 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. An obvious drawback to fiberglass is the increased component cost. The choices for alternatives in the upper and middle body tube components were largely the same. 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

10 Table 4-1: Body Tube Material Selection Matrix G10 Fiberglas s 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 G10 fiberglass is by far the strongest material considered, it is also by far the heaviest and most expensive. Additionally, the vast majority of machine shops at RPI do not allow students to work with fiberglass materials. Reinforced phenolic resin offers a good compromise between G10 fiberglass and standard phenolic resin, but remains at a high cost. 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 affordability. Kraft-paper cardboard is not expected to be strong enough to withstand the loads placed on the vehicle during flight with 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 and cost of the vehicle body tubes also increases. The design team largely considered 4 and 6 airframe diameters. 10

11 Alternatives for the coupler sections in the vehicle design are largely constrained by the selection for the vehicle body tubes. For material selection, the vehicle design team considered the same factors and variable choices as in Table 4-1. 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. 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 our budget. As fin thickness increases, the fins become stronger and drag increases. The design team sought a balance these two properties. Since general shape and surface area are not discrete factors, there is not a discrete set of alternatives for fin design. As such, we will include how the design team considered these factors and their effect on vehicle design, then present our final selection in section 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 with enough power in the 54mm diameter range are extremely long (usually over 24 in length). By considering motors in a 75mm diameter range, the length needed for the motor mount is drastically reduced. The motor mount material was subject to the similar considerations as the vehicle body tube sections, which xare summarized in Table 4-1. The centering ring materials considered included ½ birch plywood and G10 fiberglass. Plywood offers decent 11

12 strength, while being easier to work with and much more affordable. Fiberglass offers increased strength, but is a bit more difficult to work with 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 4-1. 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 Current Vehicle Design The current vehicle design is comprised of the best choices for each alternative subsystem design. Figure 4-1 shows the current vehicle model. Figure 4-1 Current Vehicle Design The overall vehicle diameter selected was 6 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 in. and a standard 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. The design team selected a G10 fiberglass nose cone with diameters matching our vehicle selection, and a 24 in. exposed length. This nose cone was selected primarily because of manufacturer availability. Components from the same manufacturer often have the best fit, and PML does not sell a plastic nose cone for their 6 in. air frames. As an additional benefit, this nose cone comes with an open bottom design, which allows the RRS to easily place ballast in the nose cone and recess the bulkhead for recovery hardware. The nose cone will be secured to the upper airframe by shear pins during flight, and will serve as the main parachute separation point. 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 12

13 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 restriction in the handbook, which prohibits forward canards. The current vehicle design utilizes two coupler sections. Both couplers will be made of the phenolic resin material supplied by PML. Each will be 12 in long, to provide a secure connection at each separation point. One will be located at the drogue separation point. This coupler will be involved in payload integration. It will be permanently fixed to the upper airframe by structural screws, and attached to the lower airframe by shear pins for drogue parachute deployment. The second coupler will be located between the middle and upper body tube sections. It will contain the vehicle's avionics bay and be permanently fixed to both the middle and upper body tube sections by way of structural screws. 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, which is found to minimize drag. The shape of the fins is shown below in Figure 4-2, and the exact dimensions can be found in Appendix B. Figure 4-2: Selected Fin Shape in OpenRocket The motor mount will be phenolic resin for reasons discussed above. Due to motor considerations as outlined in Section 4.1.5, the motor mount will be 75mm on the inner diameter. The current selected centering rings are 0.5 in thick birch plywood rings that will fit between the OD of the motor mount and the ID of the body tube. Two centering rings will align the motor mount in the airframe. The fins will be assembled onto the motor mount before insertion in the lower airframe to allow for reinforcement. These reinforcements will include both epoxy fillets along the fins, as well as fiberglass strip reinforcement. The lowermost body tube section will be slotted through the bottom of the airframe to allow for the insertion of the preassembled fin can. 13

14 4.1.5 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 4-2 below. Table 4-2: 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 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 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 Animal Motor Works L777 and the Cesaroni L

15 The Animal Motor Works (AMW) L777 is a reloadable 23% L-class motor. It has a thrust curve shown in figure 4-2 below. Figure 4-3: Thrust Curve for Animal Motor Works L777 As shown in the thrust curve, the motor burns for 3.96 s and achieves a peak thrust of 1000 N. The complete motor specifications are listed in Table 4-3 below. The estimated retail value for a complete motor kit (which includes the motor, case, hardware, and nozzle) is $325. The estimated retail value for each additional reload is $

16 Table 4-3: Specifications of AMW L777 Burn TIme Total Impulse Average Thrust Peak Thrust Launch Mass Empty Mass Diameter Length 3.96 s 3137 Ns 793 N 1000 N 130 oz 68.3 oz 2.95 in 19.6 in The Cesaroni L910 is a reloadable 12% L-class motor. It has a thrust curve shown in figure 4-3 below: Figure 4-4: Thrust Curve for Cesaroni L910 16

17 As shown in the thrust curve, the motor burns for 3.16 seconds and achieves a peak thrust of 1048 N. The complete motor specifications are shown in Table 4-4 below. The estimated retail value of the casing and all affiliated hardware is $300. The estimated retail value of each reload kit is $140. Table 4-4: Specifications of Cesaroni L910 Burn TIme Total Impulse Average Thrust Peak Thrust Launch Mass Empty Mass Diameter Length 3.16 s 2869 Ns 906 N 1048 N 92.3 oz 44.1 oz 2.95 in 13.8 in The two motors share very similar performance profiles. The differences in average thrust and total impulses can be attributed to the 0.8s difference in burn time. Unsurprisingly, the simulation profiles for both motors yielded acceptable projected apogees. A major difference between these two motors is their assembled length. The Cesaroni L910 is nearly 6 in shorter than the AMW L777. In addition, the AMW L777 is a much heavier motor both on the launch pad and post burnout. Initial cost for the Cesaroni L910 is much higher, since the casing kit does not come with a motor reload. In speaking with Jody Johnson, the design team learned that Cesaroni motors have a good reputation for being simple and easy to assemble. Based on pictures of AMW motor kits, the AMW L777 would be more complicated to assemble. Currently, the RRS has selected the Cesaroni L910 as our vehicle motor. The shorter length allows for a shorter motor mount assembly and thus shorter airframe. Its lighter mass, especially after motor burnout creates a more stable vehicle by raising the CG. Simulations also confirm that this motor meets both the requirements for static stability margin and minimum velocity at rail exit. The design team also valued Cesaroni s reputation for easy motor assembly, as well as greater product availability. Table 4-5 shows the current mass statement of the vehicle design. Payload mass is estimated at 80 oz, which is a heavy-handed estimate used until the payload design team creates a more accurate mass statement for their system. Mass growth is also expected due to general assembly, especially in the use of epoxy which is difficult to estimate during the design phase. 17

18 Table 4-5: Mass Statement of Current Vehicle Design Part Mass (oz) Quantity Subtotal (oz) Nose Cone Airframe Tubing Motor Mount Motor Retainer Centering Rings Fins (x3) Couplers Main Chute Drogue Chute Estimated Payload Ballast Motor Subtotal

19 4.1.6 Verification Plan Requirement Apogee between 5,280 ft and 5,600 ft if left unaltered Reusable 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 Design Feature Rocket Mass, Rocket Motor, Design Body Strength (fins, airframe, parachutes, etc.) General Design General Design, Motor Design Motor Motor, Motor Retainer Motor, Vehicle Design Table 4-6: Requirement and Verification Plan Verification Design Phase Construction Phase Testing 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 section, Payload/Avionics Section, Lower Body Section) Design only a single stage rocket Ensure Mass and General Rocket Design allow for a motor not exceeding Class L approved by the TRA and NAR 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 l 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 Read apogee at several launches in different weather conditions 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 19

20 Requirement Minimum velocity of 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 Design Feature Motor, Vehicle Design 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 Verification Design Phase Construction Phase Testing 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 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 Select at least 2 commercially available altimeters, Design has an independent power supply to each Altimeter Design Followed, stay as close as possible to original mas estimates, update simulation profiles to reflect any change Design Followed 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 Launch rockets several times, use flight data to ensure minimum exit velocity Rocket launched with appropriate recovery system, Drogue and Main parachutes deploy at different points in flight 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 20

21 4.1.5 Risk Analysis The following tables explain the Risk Assessment Codes (RACs) used to evaluate the potential hazards in the NASA USLI launch. RACs are presented for the initial hazard, as well as for the hazard remaining after controls and mitigations have been applied. Table 4-7 identifies the color-coded RACs, which will be referred to later when assessing individual risks. Table 4-8 outlines the different levels of risk and their acceptance levels. Table 4-9 explores definitions of severity as the risks apply to personnel, equipment, and the environment. Finally, Table 4-10 offers both qualitative and quantitative definitions of probability. In all cases, individuals involved in each task will be advised of the risks involved and proper safety precautions to be taken. Table 4-7: Risk Assessment Code Matrix Risk Assessment Codes (RACs) Likelihood Impact Value 1 - High 2 - Medium 3 - Low A - High 1A 2A 3A B - Medium 1B 2B 3B C - Low 1C 2C 3C Table 4-8: Definitions of Risk Levels Level of Risk and Level of Management Approval High Risk Moderate Risk Low Risk Undesirable. Documented approval from the NASA USLI Safety Officer. Acceptable. Documented approval from the supervisor or supervisors directly responsible for overseeing the launch. Acceptable. Documented approval not required, but an informal review by the supervisor or supervisors directly responsible for overseeing the launch. 21

22 Table 4-9: Definitions of Risk Impact Values Impact Value Definitions Description Personnel Safety and Health Facility/Equipment Environmental 1 - High Loss of life or a permanent, disabling injury. 2 - Medium Moderate injury or occupational-rela ted illness. 3 - Low First aid injury occupational-rela ted illness. Loss of facility, systems or associated hardware. Moderate damage to facilities, systems, or equipment. Minimal damage to facility, systems, or equipment. Irreversible severe environmental damage that violates law and regulation. Mitigable environmental damage without violation of law or regulation where restoration activities can be accomplished. Minimal environmental damage not violating law or regulation. Table 4-10: Definitions of Risk Likelihoods Likelihood Definitions Description Qualitative Definition Quantitative Definition A - High B - Medium C - Low High likelihood to occur immediately or expected to be continuously experienced. Expected to occur several times or occasionally within time. Very unlikely to occur and an occurrence is not expected to be experienced within time. Probability >0.1 (>10% chance) 0.1 > Probability > 0.01 (1-10% chance) 0.01 > Probability > 0 (<1% chance) 22

23 Table 4-11: Possible Risks and Mitigation Tactics Hazard Cause Effect Pre-RAC Mitigation Post-RAC Personnel exposure to harmful chemical substances (resin, epoxies, etc.) Mishandling of chemical materials Personnel injury requiring medical treatment 3A Designated personnel will wear PPE (Personal Protective Equipment) such as gloves and splash-proof eyewear during handling of chemicals. Personnel will be familiarized with the MSDS of used chemicals and their station. 3C Personnel unexpected contact with heavy machinery - Misuse of machinery or equipment -Malfunction of machinery Personnel injury requiring medical treatment Possible permanent injury or death 1B Only a single, experienced operator on one tool at a time. Minimize distractions Refrain from wearing loose clothing Handling of tools only after they are finished cooling or spinning. Proper tool maintenance and knowledge of first aid stations. 1C Personnel exposure to harmful voltage Contact with energized electrical systems Personnel injury requiring medical treatment Possible permanent injury or death 2B Systems will not be handled by personnel while energized during testing. Gloves may be worn to protect against accidental contact with electrical components. 2C Personnel exposure to toxic fumes -Chemical spillage resulting in airborne particles -Improper mixing of chemicals Personnel injury requiring medical treatment Possible permanent lung damage or death 2B All personnel will be familiar with all chemicals involved 2C Inhibited Test Launches - Adverse weather conditions - Lack of personnel with necessary NAR or TRA certification. Loss of vital test data Inability to launch rocket 2B Test launches can be scheduled before harsh winter conditions. Flexible launch windows can be scheduled. Presence of an additional team member with necessary NAR or TRA certification. 2C 23

24 Hazard Cause Effect Pre-RAC Mitigation Post-RAC Late arrival of parts -Issues with shipping company/posta l service -Error on the part of the shipper Delay of assembly, possibly pushing assembly past deadline 2B All parts will be ordered with enough time before they will be needed. Expedited shipping will be selected as appropriate. 2C Malfunction of mechanical parts or systems -Fracture, yielding, or some other form of mechanical failure in an essential mechanical component Loss of vital test data Inability to launch rocket Possible in-flight failure of critical systems 1B Redundancy of mechanical systems Thorough materials testing Generous factors of safety Multiple small-scale tests 1C Malfunction of electrical parts or systems -Short circuit -Failure of analog electrical component -Failure of digital electrical component -Failure of microprocessor Loss of vital test data Inability to launch rocket Possible in-flight failure of critical systems 1B Redundancy of electrical systems Thorough inspection Multiple small-scale tests 1C Malfunction of software or code -Exposure to signals, data, or numbers not programmed to process - Bugs - Infinite loop or other programming error Loss of vital test data Inability to launch rocket Possible in-flight failure of critical systems 1B Thorough checking by multiple programmers Multiple small-scale tests 1C Failure of sensor -Faulty wiring -Exposure to elements -Exposure to excessive acceleration -Internal issues Loss of vital test data Inability to launch rocket In-flight failure of critical systems 2B Redundancy of sensors vital to rocket operation Thorough inspection Multiple small-scale tests 2C 24

25 Hazard Cause Effect Pre-RAC Mitigation Post-RAC Failure of drag flap system -Failure of any associated mechanical or electrical system or code Inability to control rocket s vertical displacement or speed during ascent 2B Redundancy of subsystems Thorough inspection Multiple small-scale tests 2C Failure of either parachute -Mechanical failure of cord, fabric, etc. -Failure to deploy Loss of vital test data Possible destruction of subsystems or entire rocket 1B Assuring only high-quality materials are used Thorough inspection Multiple small-scale tests 1C 25

26 4.1.6 Component Analysis Table 4-12: 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 Fin Can and Drogue Parachute Phenolic Airframe Tubing $50 publicmissiles.com/products/ Nosecone Nosecone and Main Parachute Mount High-Strength Plastic+forged steel eye bolt+epoxy $10 publicmissiles.com/products/ Main Airframe Airframe containing Main Parachute and Payload Section Phenolic Airframe Tubing $50 mcmastercarr.com Payload Section Containment system for payload (including bulkheads, couplers, etc.) Plywood Bulkheads, Phenolic Cardboard Couplers, forged steel eye bolts, steel threaded rod, epoxy, acrylic plastic $50 publicmissiles.com/products/ Recovery Electronics Altimeters and supporting electronics that run recovery system G10 Fiberglass Sled, Altimeters, Batteries, copper wiring, switches $50 sparkfun.com Ejection Charges Charges that eject Parachutes on Altimeter command Aluminum Blast Cap holders, Ejection Charges, epoxy, copper wiring $50 apogeerockets.com Parachutes Drogue and Main Two Skyangle Parachutes $25 b2rocketry.com Motor Retainer Aeropack Motor Retention System Aluminum $42 macperformancerocketry.com 26

27 4.2 Recovery Subsystem System Level Design The main parachute selection is based on positive past club experience with SkyAngle Classic series parachutes. The team compared the SkyAngle Classic with the SkyAngle Classic II and chose the SkyAngle Classic II. The zero-porosity silicone-coated balloon cloth of the Classic II will have a lower rate of descent and and increased stability, as compared to the SkyAngle Classic. The allowable opening shock on the zero-porosity material will be lower than that of a higher porosity fabric, but will not have a significant negative effect on the suspension line, line attachments, or eyebolts. The lowest allowable strength between these components is in the swivel connections, which can support a 1500 lbf load. This allowable strength is considerably stronger than needed for anticipated loads. SkyAngle parachutes use a non-circular surface area, which does not directly translate to a drag coefficient typically used in parachute calculations. The effective surface area is the surface area of the parachute if the non-traditional shape is translated to a traditional surface area calculation. The main parachute has an effective surface area of 15.8 square feet.the manufacturer also lists loads to which the parachutes have been subjected without failure. Based on the manufacturer's recommendations, the selected parachute size is acceptable for rockets in the weight range of our rocket. All eye bolt connections need to be strong enough to withstand the forces associated with parachute deployment. Eye bolts are sold as either open or closed shapes. The closed shape is able to withstand higher loads than open shapes due to basic principles of shear flow. Closed eye bolts are manufactured with either welded or forged steel. Forged eye bolts are built as load-bearing eye bolts, with much higher allowable stresses. This manufacturing method constructs the eye bolt as a solid piece of steel, rather than having welds within the system that are susceptible to stress concentrations. Eye bolts have the option of including a shoulder at the base of the ring above the shank. Since the eye bolts will be holding a moving shock cord, the load may not be completely in the axial direction of the eye bolt. The shoulder allows the eye bolt to withstand shear loads, as will be seen in this system, more effectively. Stainless steel is preferable to steel or iron due to the high ultimate tensile strength (UTS) of stainless steel. In addition, the stainless steel has high corrosion resistance compared to iron and steel, even when coated. It is most convenient to use standardized sizes to minimize cost and system complexity. The eyebolts we selected are forged stainless steel 316, ¼ -20 and have a shank length of two inches. The manufacturer has a recommended work load limit of 500 lbs. This is within expected loads. The two most common materials for shock cord in high-powered rocketry are kevlar and tubular nylon. Kevlar has more resistance to the heat and corrosive environment of ejection gasses, but has a much lower UTS compared to tubular nylon. Tubular nylon also has more elasticity when initially loaded, therefore 27

28 making it a more effective material for use in large rockets. Based on these factors, tubular nylon will be used for the shock cord System Components Table 4-13:Altimeter Comparison Part 1 Featherweight Raven 3 PerfectFlite Pnut PerfectFlite Stratologger SL100 Price $155 $64.95 $54.95 Dimensions 0.80 W 1.80 L 0.55 T 0.63 W 2.36 L 0.48 T 0.90 W 2.75 L 0.50 T Altitude Accuracy Operating Voltage +/- 0.3% +/- 0.1% +/- 0.1% V V 4-16 V Sample Rate 20 Hz 20 Hz 20 Hz Maximum Altitude 100,000 ft 100,000 ft 100,000 ft 28

29 Table 4-14: Altimeter Comparison Part 2 Pros Cons Featherweight Raven 3 Includes an accelerometer Quality software for post flight analysis 8 minutes of high rate data and 45 more minutes recorded per flight Software for flight simulations included Deploys parachutes with audible ematch continuity Accelerometer-base apogee deployment Includes backup deployment 1.5 seconds after the main deployment Very expensive Needs to be mounted vertically PerfectFlite Pnut Lithium ion battery is rechargeable Post flight locator siren battery lasts for two weeks Fairly inexpensive Deploys parachutes with audible ematch continuity Limited to an input of 4.2 volts No post flight software Charging temperature must be above 32 degrees Fahrenheit PerfectFlite Stratologger SL100 Post flight locator siren Deploys parachutes with audible ematch continuity High resistance to false triggering due wind and sunlight Provides real-time data serial output Main chute deployment altitude is adjustable Fairly inexpensive Selectable apogee delay Post flight software not included with purchase The Featherweight Raven 3 is the selected altimeter because of the accompanying post flight analysis software. The benefit of the included software outweighs the cost of the altimeter and the vehicle design can easily accommodate for the vertical mount. The PerfectFlite Stratologger was chosen as the redundant backup altimeter because of the adjustable deployment altitude. Unpredictable atmospheric conditions eliminates the Pnut with its low tolerance for cold temperatures. The post flight software available for each altimeter was the most significant deciding factor. 29

30 Table 4-15: Parachute Selection SkyAngle Classic 60 SkyAngle Classic II 60 Total Surface area 39.3 square feet 39.3 square feet Tested Drag Coefficient Cd Suspension Line Length 60in 60in Weight 10oz 18.2 oz Price $90.00 $99.00 Material Low-porosity 1.3oz silicone-coated ripstop nylon Zero-porosity 1.9oz silicone-coated balloon cloth Line Attachment 1000lb size 9/0 nickel-plated swivel 1500lb size 12/0 nickel-plated swivel The mass statement from Section was used to determine loads on the recovery system components. To Keep drift within the acceptable limits, the RRS has chosen a 3 foot Ballistic Mach II drogue chute. This parachute was chosen for its strength and flight characteristics, based on available data from the manufacturer. Data collected from OpenRocket simulations indicates descent velocity after main chute deployment to be ft/s, consistent with manual calculations conducted by the RRS. The ejection charges for the parachute deployment require 1 gram of black powder charge per 6 inches of interior body length containing the parachute, according to estimates made by Vern Knowles of Vern s Rocketry. This leads to an approximate calculated ejection charge size of 1.7 g of black powder for main parachute deployment and 1.7g of black powder for drogue parachute deployment. The attachment scheme between the rocket body and the main parachute consists of shock cord connecting the main parachute to a 1.5 inch eyebolt mounted to the main airframe forward ½ thick birch bulkhead. The drogue parachute and the body will be attached in a similar manner; there will be eyebolt attachments on the forwardmost motor centering ring and on the rear of the payload section. Both the fore and aft shock cords will be 250 long, or about 3 body lengths, to ensure adequate separation of vehicle components during descent. Simulation data from OpenRocket indicates that the maximum 30

31 acceleration the rocket will experience during descent is 300 ft/s 2 and occurs when the main parachute deploys, at an altitude of 700 feet. The force exerted on the rocket during deployment can be calculated using the mass of the rocket. This force is used to ensure that the loads on the attachment hardware do not exceed the ultimate strength of the components. The electronic components of the recovery system consists of a StratoLogger CF altimeter and a Raven 3 altimeter, and can be seen in detail in Figure 4-4 below. The design of the electronics allows for externally accessible switches on the exterior of the payload section to arm and disarm the charges and altimeters. The two altimeters provide redundancy, as the flight system is capable of functioning on either should one fail mid-flight. At the appropriate altitudes, each altimeter will send signals to trigger the its set of charges for each ejection event. Thus, each altimeter can trigger charges for each separation event, and the charges themselves are redundant. If one fails, parachute deployment will not be adversely affected. These will be black powder charges, mounted in blast cups of the fore and aft bulkheads of the main airframe. These charges will separate the rocket sections and allow unfurling of the parachutes. The drogue will be deployed at apogee when the first pair of charges are detonated, and the main chute will be deployed at 700 feet when the second charge pair triggers. Kevlar parachute protectors will insulate the parachutes from the heat effects of the charge firings. Figure 4-5: Recovery System Electronics Schematic Testing of the recovery system will be done via simulation software for initial validation, and final testing and modification will occur on scale model flights. Current simulation data has been obtained from OpenRocket.. Ground tests of the ejection charges and electric matches may also occur. 4.3 Mission Performance Predictions Mission Performance Criteria The rocket has several criteria that must be met on launch day. The primary goal is to obtain an apogee as close to 5280 feet as possible. This will be achieved by the selection of a motor that will overshoot the target, then usage of 31

32 the drag flap system to slow the coasting velocity down. Additionally, the vehicle must land within a half mile of the launch site. To this end, the recovery system will be tailored so that the main parachute is deployed in a manner thus drift is mitigated Vehicle Simulations Figure 4-4 in Section shows the simulated motor thrust curve for a Cesaroni L910 Motor. Utilizing OpenRocket and this motor configuration, a simulated model of the rocket was produced as shown in Figure 4-6. Figure 4-6: Model for Flight Simulation (Launch-Ready Motor) On this model, there is a stability margin of 1.82 Cal with the motor included. The simulated center of gravity is approximately 63 in down from the top of the nose cone, and the simulated center of pressure is approximately 74 in from the top. To verify that the simulation model adheres to requirements 1.14 and 1.15 in the statement of work, simulations were analyzed at the point of rail exit. The simulated velocity at rail exit is 78.2 ft/s, which complies with the minimum exit velocity of 52 ft/s. A custom MatLab tool was used to determine the mass of the motor at the top of the launch rail. Using this data, we determined that the static stability margin at the top of the rail is 2.08 Cal, which complies with the minimum stability margin of 2 Cal. Additionally, the vehicle design team analyzed static stability after motor burnout. As shown in Figure 4-7, this analysis was performed by removing the motor from an OpenRocket profile, then adding a mass component of equal size and weight to a burnout L910 motor. 32

33 Figure 4-7: Model for Flight Simulation (After Motor Burnout) After motor burnout, the center of gravity in this configuration is approximately 59 in down from the top of the nose cone. This change in center of gravity creates a static stability margin of 2.46 during the coast phase. The vehicle design team will be coordinating with the payload design team to ensure that inducing a roll rotation does not compromise overall vehicle stability. This will be primarily achieved through CFD analysis. Using this model and the selected motor, a flight simulation was created which is shown in Figure 4-8. Figure 4-8: Flight Simulation from OpenRocket Note that this simulation does not take into account the effect that the drag blade payload system will have on flight performance. Given that, the apogee reached is 5344 ft, with a maximum acceleration of 279 ft/s 2, and flight time of 106 seconds. 33

34 4.3.3 Kinetic Energy Analysis Table 4-16: Variables in Calculation of Final Descent Velocity Variable Value mg lb ρ slug / ft 3 C D 1.89 A 15.8 ft 2 Final Velocity ft / s The final descent velocity was calculated using the following equation V = 2mg ρac D Where mg is the weight of the rocket, A is the effective area of the main parachute, ρ is the air density, and C D is the drag coefficient. The values are given in Table The final velocity of ft / s was used to calculate the kinetic energy of the vehicle sections shown in Table Table 4-17: Landing Kinetic Energies of Independent Sections Vehicle Section Nose Cone Upper Airframe Lower Airframe Kinetic Energy (lb ft) Drift Analysis Table 4-18 lists the calculated drift distances with varying wind speeds. 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. 34

35 Table 4-18: Calculated Drift Distances for Varying Wind Speeds Wind Speed (mph) Drift Distance (ft) Safety Safety Officer 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. He will monitor or designate a monitor for all RRS lab meetings Preliminary Assembly and Launch Procedures 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 pre-launch 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 checklists are detailed in Table 4-19 and Table

36 Table 4-19: Pre-Launch Operations Checklist Category Description Required Actions Parachute Deployment Avionics Payload Roll Blades Protective insulation will be inserted between the ejection charges and the parachute to prevent burning, and the parachute will be properly folded and then inserted. 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. Payload is inserted into payload tube and fixed. Shear pins are inserted. Roll Blades will be checked for unobstructed movement. protective insulation insertion, parachute folded, parachute inserted. 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 inserted in right orientation, payload bay secured, shear pins inserted. Verify that Roll Blades can extend and retract fully by running control system test. 36

37 Table 4-20: Launchpad Operations Checklist Category Description Required Actions Rocket Assembly Rocket assembly is complete. Ensure all components and independent sections are properly connected and secured RSO Safety Check Avionics Avionics Ejection Charges Motor Ignition Launch Configuration Avionics Ejection Charges Final Safety Check RSO will check final rocket before launch. Avionics will be checked for proper communication and data relay. Avionics charge and ignition switch will be turned off to prevent premature ignition. Ejection charges will be installed. Motor igniter will be installed. Delay charge will be installed. Leads will be connected during handling to prevent static electricity discharges. Rocket will be placed on the launch rod and given the desired angle. Avionics charge and ignition switch is turned on to arm the system. Ejection charges will be armed. Final safety check will be conducted by RRS SO, RRS Mentor. RSO Approval. RRS SO Approval. RRS Mentor Approval. Data link is clear. External switch is OFF. Confirm status light is also OFF. Ejection charges installed 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. Rocket rail buttons are on launch rod. Selected angle is confirmed by RSO. External switch is ON. Confirm status light is also ON. Ejection charges are connected and will not short. RRS SO Approval. RRS Mentor Approval. Data Retrieval Final altitude will be retrieved. Altitude has been retrieved by SL official. Post-Flight Inspection Inspection will be performed on rocket following flight. RRS SO Approval. RRS Mentor Approval. 37

38 Preliminary Personal Hazard Analysis The following section outlines safety risks and their mitigations that are likely to be encountered in the construction and testing of the rocket. All participating team members have completed safety training as administered by the RPI School of Engineering Material Hazards Solid rocket boosters, epoxies, resins, solder, and other chemicals and materials needed for the construction of the rocket can cause burns, skin and/or eye irritation, and/or may have risks associated with fumes. Mitigations include an orderly workspace; familiarity with relevant MSDS; wearing appropriate clothing and PPE; and knowledge of locations of first aid stations. Table 4-21 below outlines some of the potential material hazards to which team members may be exposed during the project. Table 4-21: Material Hazards Hazard Effect Pre - RAC Mitigation Required Safety Equipment Emergenc y Equipment Post - RAC Epoxy / Paint fumes / irritation Inhaled fumes cause damage, substance can cause blindness if eye contact. 2B Work in ventilated and spacious area, knowledge of MSDS Face masks, gloves, glasses First aid kit, eye flushing station 2C Propellant/powde r ignition and skin irritation Powder can irritate skin. Hot powder can cause burns. 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, fire extinguishers 2C Vehicle debris Falling vehicle debris could seriously injure an onlooker. 2B NAR and TRA Glasses regulations will be obeyed. Team members will be trained to watch skies and avoid any falling vehicle debris. First aid kit, eye flushing equipment 2C Facility and Tool Hazards Tools such as lathes, mills, presses, and other equipment found in lab spaces present risks in the forms of cuts, burns, bruises, broken bones, and damage to eyes. 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. 38

39 The RRS safety standards include important stipulations regarding fire safety. Emphasis is placed on maintaining three properly cleared exit paths, all marked with emergency lighting and exit signs. No smoking or any other activity that uses a lighter or open flame is allowed in RRS lab building. The RRS also has access to two fire extinguishers, one inside the lab and the other just outside the doorway. Each of these is checked monthly by the safety officer. The RRS lab does not contain any equipment that uses electrical power exceeding 12 volts. However, all RRS members are required to take safety briefings that include information regarding devices that exceed 12 volts. All devices and equipment that use electricity are checked regularly for signs of damage or wear, and particular attention is payed to frayed or improperly insulated wires. 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 4-22 below. 39

40 Table 4-22: Facility Hazards Hazard Effect Pre - RAC Mitigation Required Emergenc Safety y Equipment Equipment Post - RAC Injury related to drill press and mill rotation Contact with drill could cause serious injury / death. 1B No loose clothing/hair Glasses First aid kit 1C Injury related to band saw, belt sander, and other similar equipment Contact with saw could cause serious injury / death. 1B No loose clothing/hair Gloves when working with large pieces, glasses First aid kit 1C 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 debris. Glasses First aid kit 2C 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 Fire Fire can be caused by a wide range of accidents, including electrical and chemical sources 1C Maintain properly cleared exits. No smoking or any other activity that uses a lighter or open flame. Proper usage of tools and chemicals. N/A First Aid kit, Fire Extinguishers, personal cell phones, burn kit 1D Safety Officer Unavailable No identifiable person is in charge of ensuring adequate safety operations followed 2B Numerous students involved in the club are Level 1 certified by TRA and NAR. All participants briefed on safety operations. N/A Officer chain of command structure is identified in RRS constitution. 2D Tripping hazard in lab Unclean lab space could injure persons attempting to leave lab. 3B Maintain properly cleared exits. Keep Lab space trash free. N/A N/A 3D 40

41 4.5.4 Preliminary Failure Modes and Effects Analysis 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 out of three and impact out five. Table 4-23: Failure Modes System Failure Risk Impact Mitigation Motor Dislodgement from housing 1 5 Employ retainer structure with three bulkhead centering rings. Computer / Control Microcontroller failure 1 3 All wiring will be finalized and tested before launch to prevent electrical malfunction. Code will be reviewed and tested during scale launches. Wiring will be soldered and insulated. Stepper motor jam 1 2 Stepper motor will be tested against static loads to simulate expected air speeds. Mechanism will be lubricated and inspected for particulates before launch. Forward flap system causes instability 1 5 Verification of stability will be performed in OpenRocket and using test launches. Structural Fins shear off during flight 2 4 Fins will be embedded in body and fixed to the internal motor mount and body with heat-tolerant epoxy resin. Rocket motor explodes 1 5 Test launches, finite elements analysis, and static load tests will be performed to verify integrity of rocket body. Parachute tears, parachute fails to deploy 2 5 Parachute will be checked for defects before packing. Parachute will be packed properly. Shock cord breaks 2 5 Recently-manufactured shock cord will be used to ensure integrity. Shock cord will be properly secured and tested under static and jolt loads. Drag flaps shear off during flight 2 2 Mechanical components and connections will be shear tested under static loads. Recovery Parachute ejection failure 2 5 Redundant powder charges will be used for the main parachute. Packing will be inspected before launch. Parachute ejected prematurely 1 3 Code will be reviewed for correctness. Hard failsafes will be implemented to prevent ejection above a predetermined altitude. Computer system failure 2 4 Flight computer will be tested using simulated soft landing jolts to ensure component survivability. Wiring will be soldered in place and insulated. Transmitter failure/interference 2 3 Transmitter will be tested at range. Interference will be studied. 41

42 System Failure Risk Impact Mitigation Launch operations and transportation Physical damage during handling Live charges on board after launch command Premature firing of ejection charges and/or motor 1 5 Each component of the launch vehicle will be transported in a separate, cushioned box. Launch vehicle will be inspected on site for defects. 1 3 Avionics will be powered off. Discrete switch will cut off system power entirely. 1 5 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. Spectator injury 1 5 Range will be cleared prior to launch. No large, heavy objects will be installed according to NAR safety code. 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 Preliminary Environmental Concerns Several environmental concerns that may impact the testing and launch of the rocket have been identified in the following table. 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. Johnson 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

43 Table 4-24: Environmental Hazards to Project Completion Risk Mitigation Likelihood Inoperable Wind Conditions / Cloudy Rainy Conditions Operable Wind Conditions Vehicle Cannot be Found Launches will be planned in advance according to weather forecasts to ensure timely completion. All electromechanical and electrical parts will be shielded. Body structure will be able to retain integrity in damp conditions. Launch rod angle will be adjusted. Parachute deployment altitude will be recalculated and adjusted for higher-than-expected drift. All charges will be detonated by flight computer before landing to ensure that it is inert. Vehicle will have identifiable, bright paint scheme Project Risks Risks to Project Completion have been identified in Table They have been addressed with a mitigation plan and their likelihood and impact is ranked low to high. Table 4-25: Risks to Project Completion Risk Definition Mitigation Likelihood Impact Time Team may have difficulties completing project in competition timeline. The team has instituted weekly and monthly progress points to ensure that work is being done in a timely manner. Medium High Resources The team must procure all necessary resources for the competition. The team has prepared detailed bills of materials. Components will be ordered early in advance. Medium High Budget There is a large budget associated with participating in the NASA USLI competition. The team has a semesterly fee to join. Additionally the treasurer has set up a donation site that helps fund the team. Low Medium Scope The competition covers a vast scope of challenges. The team has gained a large number of members with a variety of engineering skills. Low Low 43

44 5. Payload Criteria 5.1 Selection and Design of Payload This subsection fulfills Vehicle Payload Experiment Report Criteria 5.1 (objective, experiment and successful results), 5.2 (system level review of design), and 5.4 (current leading choices) Roll-Blade 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 3.3 (Option 2). 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 active drag control, the final altitude would be slightly above this height. The drag control will be dynamically controlled in flight to decrease the vertical velocity of the rocket and adjust the maximum height of the rocket to exactly 5,280 feet at apogee. The roll-blade subsystem will also fulfill Option 2 of Payload Requirement 3.3, Roll Induction and Counter Roll. Two sets of blades will induce both the roll and counter roll. Additionally, the performance of the blades will be simulated and studied using ANSYS Fluent CFD software. This analysis will help to verify the robustness and performance of the subsystem. Figure 5-1: Roll-Blade Subsystem The blade system is, in reality, two sets of an identical system. Each system consists of the 3 blades, a straight plate and a cam plate. One of the two systems is flipped and placed on top of the other, creating the complete 44

45 Roll-Blade Subsystem pictured above. Having the two sets on the same plane diminishes turbulent effects that would propagate if the sets were separated vertically along the rocket. This will allow for easier CFD analysis. The blades are have the airfoil shape of a NASA NLF15 airfoil and will serve to impart a roll moment about the rocket in order to accomplish Option 2 of Payload Requirement 3.3. The airfoils will be attached to threaded rods which will be passed through the tracks of two plates - the cam plate and the straight plate either above or below them. The threaded rods will be fitted with rollers and nuts. The rollers will be in place to reduce resistance in the tracks as the actuation occurs; and the nuts will be in place at the end of the rods in order to constrain them. A third set of plates will be utilized in order to mount the motors that will drive the actuation. This set will be on the top and bottom of the pictured system. A picture of the complete mounting and integration of the system can be found in section Microprocessor System The scientific subsystem will contain a central motherboard containing the processing components, batteries, memory, and most of the scientific instruments on the rocket. The remaining instruments will be connected to the central board via screw terminals. An Arduino or Arduino clone will serve as the processor of the board. This was chosen because of simple implementation and former experience. A magnetometer will be used to to determine the orientation of the rocket. The Arduino will read a serial stream from an additional payload stratologger altimeter to ensure that the rotation of the rocket begins above a specified altitude. The magnetometer readings will be used to ensure that the rocket rotates the correct amount. To turn the rocket, the Arduino board will be connected to two different H-Bridge chips that will allow the stepper motors to turn with more power than the Arduino can natively supply. The motion of the motors will be determined by a either a PID or PI controller that will be fed the output of the magnetometer. The motor s motion will directly translate to the extension of the blades. Additionally, the board will contain a ublox GPS receiver, an XBee wireless transmitter, and an SD card module. The components will be individually tested on a breadboard. The wiring will then be recorded and adapted to a PCB that will be ordered from either OSH park or Seeed. This provides a balance between the versatility of a breadboard design without the hassle of wiring many connections together. 45

46 5.1.3 Integration with Vehicle Subsystem This subsection fulfills Vehicle Payload Experiment Report Criteria 5.6 (preliminary interfaces between payload and vehicle). Payload integration into the vehicle is made simple by having the integration in mind throughout the design process. The payload exists as a black-box with respect to the vehicle, being secured by twelve screws through the holes in the top and bottom of the struts as seen in figure 5-2. Fig 5-2: Complete Payload System Render The payload assembly slides into the the bottom of the upper airframe and is constrained in the axial and rotational directions by twelve screws from the exterior of the rocket body. A visualization of the payload assembly housed inside the rocket body can be seen below in figure

47 Fig 5-3: Payload Integrated with Vehicle 5.2 Payload Concept Features and Definition This subsection fulfills Vehicle Payload Experiment Report Criteria 5.4 (advantages of current leading choice). The RRS s 2011 USLI team had similar payload objectives. In 2011, the RRS used an active drag control system to manipulate the target apogee and a camera system to capture the horizon in the correct orientation. The RRS also developed a different active drag control system for USLI last year (2015) which consisted of a drag-flap system. An active drag control system is being developed for 2016 which aims to improve on the system used in It consists of two sets of three blades (6 blades total), capable of changing the rocket s drag coefficient during flight. Each set of blades is independently actuated by its own cam mechanism to allow deployment and retraction as necessary to accomplish mission objectives, namely altitude requirement, and roll induction and counter roll (req. 3.3). Two sets of fixed, oppositely-pitched blades are necessary for inducing both positive and negative roll torque for req Rotation of each cam mechanism is accomplished by a motor, controlled by a central microprocessor. Rotation of the cam geometry induces radial movement of the associated blade set. The cam mechanism allows a single motor input to identically affect each blade of the set, thus greatly minimizing the potential of uneven deployment (an unstable-flight scenario). The number of critical moving parts is also smaller for the cam mechanism compared to alternatives (rack and pinion, or crank mechanisms would require many moving parts to redirect inputted motion for each individual blade). Additionally, the natural cylindrical envelope of the cam mechanism makes efficient use of the natural cylindrical interior of the rocket body. 5.3 Science Value Data will be collected for the blade payload during launch from redundant gyroscopes and two dedicated accelerometers onto an onboard SD card. Blade control will be reliant on real-time information from the gyroscopes and magnetometers and will attempt to match an idealized rotation profile. Initial blade control will be based on a theoretical model of the mechanism and forces during flight. Assessment of these theoretical models will be conducted post-launch through analysis of gyroscope, 47

48 accelerometer, motor outputs and measured deviation from the idealized rotation profile. Theoretical models will be iteratively improved and optimized using the collected data and the effectiveness of the blade payload system as a joint active drag system and roll-control system will be assessed. Supplementary position data will be outputted by the onboard GPS. 5.4 Verification of Payload System Level Verification This subsection fulfills Vehicle Payload Experiment Report Criteria 5.2 (system level review of design) and 5.3 (cons of alternatives). Table 5-1: System Level Verification Blade Design Geared Blades Reaction Wheel Pros - Unique airfoil design - Wind turbine-based design - Heritage cam design from Motors not bearing direct load - Can alter angle of attack - Fewer blades necessary - Mechanically simple to design and implement - Flight proven - Minimal aerodynamic interactions Cons - Difficult aerodynamic analysis - Friction - More difficult to manufacture (waterjet and CNC) - Uses aerodynamic forces to induce roll - Complex (more parts to fail) - Motors bear direct load - Uses aerodynamic forces to induce roll - Adds a lot of weight - Not a new design Cold Gas Thrusters - Flight proven - Solid State - Not a mechanical system (Requirement 3.3.3) Defense of Current Leading Design This subsection fulfills Vehicle Payload Experiment Report Criteria 5.2 (system level design review), and 5.4 (current leading choices). As discussed in Table 5-1, the current leading payload design is the blade design. One of the biggest strengths this design has is that the motors do not bear a direct load, as they do in the geared blade design. While the blades are more difficult to analyze, this can be combatted by using CFD. The friction between the track plates and mating rods will be reduced by adding miniature rollers. The following table was designed in order to satisfy PDR Payload requirement #4 - a defense of the current leading payload design. 48

49 Table 5-2: Current Leading Payload Designs Airfoil Fig 5-4: NASA NLF15 Airfoil Profile The airfoil chosen for the blades is NASA NLF15. Chosen for high L/D at our Reynold s Numbers - (~300, ,000) and for structural stability a 16% thickness to chord ratio was selected (at 40% chord). The Reynold s numbers were found using variable atmospheric conditions varying with elevation as well as variable velocity from the OpenRocket simulations. Fig 5-5: Reynold s Number vs Altitude During Cruise Rollers To assist with blade deployment attempt utilized a similar cam system and experienced a large amount of friction due to tolerancing and manufacturing issues. 49

50 3 blades The number of blades is determined by the maximum possible efficiency of the blades, relative to the amount of spin they can cause and the drag produced. More blades increases the solidity of the total disc area, which further decreases the aerodynamic efficiency of the blades thereby decreasing tip speed. Tip speed is directly related to the rotating rate of the blade system. Three blades has been historically determined to be the optimum solution for solidity and turbine blade count in similar applications. Material Choices The blades will be constructed of Delrin - a low friction polymer that mates well with metal. It has a fairly high tensile yield strength at 73 MPa, it is around 10-50% the strength of aluminum and approximately twice that of PTFE. Another major benefit to delrin is its ease of machining using conventional methods available at RPI s manufacturing centers. The plates will be constructed of aluminum (possibly 6061-T6). While relatively dense, aluminum has a very high tensile yield strength at 276 MPa. It has the capability to be thinner than the Delrin and the sheet can be water jetted. Shape of track The shape of the cam track is derived from a polynomial, with a rise of 1.75 over an angular displacement of 150 degrees. The polynomial shape is chosen to have continuous displacement, velocity, acceleration, and jerk. The slot is to be waterjet, and includes tolerances for the rollers of the cam assembly. Fig 5-6: Shape of Cam Track 50

51 6. Project Plan 6.1 Budget Plan Table 6-1 provides a summary of important sections, as well as our projected bottom lines. Appendix C contains the fully detailed budget. Table 6-1: RRS Budget Summary Budget Summary Vehicle Design Team Expenses $1, Recovery Team Expenses $ Payload Team Expenses $ Travel Costs $3, Income $10, Total Expenditures $5, Total Budget $4, Funding Plan The RRS is very well funded for this year s competition. In addition to funds remaining from last year, we have received $1,000 from the RPI School of Engineering and about $800 in membership dues. We also expect to raise money through a program at RPI that connects donors with student groups. This program, called wer Gold, sponsors student-run projects by reaching out to alumni and friends of Rensselaer. We were successful in applying for funding last year, and we are applying for funding again this year with a goal of $3,000. At this time, we have sufficient funds for the construction and testing of our rocket. The remaining funds will be used for travel and lodging at the competition and saved to enable the club to compete in future years. 6.3 Timeline The RRS continues to follow the Gantt chart schedule that was created during the proposal stage of the competition. Thus far, the schedule has been followed as close as possible. Both the initial ordering process and construction of the subscale vehicle began a few days late. Due to the detail and depth of the Gantt chart, there is no easily readable way to include it in this report. To this end, the RRS has updated the Gantt chart hosted here [1] [1] rrs.union.rpi.edu/nasa2017.html 51

52 6.4 Educational Engagement Status Soda Bottle Rockets and Open House The RRS taught sixty-five incoming Rensselaer freshmen students about rocketry. We used water rockets, as explained at this link [2]. The launch pad was built beforehand. First, we held a preliminary discussion to educate the students about basics of rocketry and gave instruction for building a soda bottle rocket. Soda bottles, cardboard, and other materials for constructing the rockets were provided to the students. Students were divided into teams to design and constructed their own rockets. After the construction period we launched the rocket. We concluded the event by bringing the students to the RRS meeting and design room to discuss and reflect on what the students observed and learned. This also served as an open house for the Rensselaer community. Following the soda bottle rockets discussion, attendees were presented a short info session about general rocketry, the Rensselaer Rocket Society and the RRS s rocket. General rocketry included a history of rocketry, types of rockets, and modern and future uses of rockets. The presentation of the RRS will introduced the members of the team and briefly discuss some individuals majors and interests. A presentation was given introducing the RRS s rocket including specifications, an overview of the objectives for the rocket, and an analysis of construction of the rocket National Manufacturing Day As a part of National Manufacturing Day at RPI, the RRS gave a brief presentation on manufacturing practices in the aerospace industry and how they correlate to practices that the RRS use as well. Afterwards we led the students through a hands on exploration of our drag flap system from last year. We met a total of 32 students, in two waves, and we had them break down into five groups each group was allowed to work with a miniature version of the drag flap system we had 3D printed out prior. [2] spaceflightsystems.grc.nasa.gov/education/rocket/bottlerocket/about.htm 52

53 Appendix 53

54 Appendix A: Milestone Review Flysheet 54

55 55