2015­2016 NASA USLI Critical Design Review (CDR)

Rensselaer Rocket Society (RRS)

Rensselaer Polytechnic Institute 1999 Burdett Avenue

Troy, NY 12180

Project Name: Red Gemini Task 3.1.1 – Atmospheric Measurements

Task 3.1.6 – Aerodynamic Analysis

Friday, January 15, 2015 1

1. Table of Contents

Table of Contents 2. Executive Summary

2.1 Team Summary 2.2 Launch Vehicle Summary 2.3 Payload Summary

3. Changes Made Since Preliminary Design Review 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 4.2 Recovery Subsystem 4.3 Mission Performance Predictions 4.4 Interfaces and Integration 4.5 Safety

5. Payload Criteria 5.1 Selection, Design, and Verification of Payload 5.2 Payload Concept Features and Definition 5.3 Science Value

6. Project Plan 6.1 Budget Plan 6.2 Funding Plan 6.3 Timeline 6.4 Educational Engagement Plan and Status

7. Conclusion Appendix Appendix A: Scientific Payload Schematics Appendix B: Structural Design Assembly Drawings Appendix C: Image Rotation Code

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2. Executive Summary 2.1 Team Summary This report was produced by the Rensselaer Rocket Society (RRS) of Rensselaer Polytechnic Institute (RPI). The RRS operates out of the Ricketts Building at RPI. The mailing address for the RRS will be the same as that for RPI: 1999 Burdett Ave, Troy, NY 12180. The Community Mentor for the RRS is John Sicker, who is certified to Level 2 with the National Association of Rocketry (NAR) and Level 3 with the Tripoli Rocketry Association (TRA). His NAR member number is 49422, and his TRA member number is 01017. 2.2 Launch Vehicle Summary The launch vehicle design originated with the Liberty 4 from Giant Leap Rocketry. However, the rocket will be approximately 104 inches in height and will have a body 4.02 inches in diameter. The rocket will have a mass of about 13.176 pounds mass. The launch vehicle will be propelled by an Aerotech K1103X motor for a 54mm motor mount. The recovery system will consist of a main and a drogue parachute to be deployed via electronic deployment. This deployment will be controlled by one of two altimeters to be used for redundancy: a StratoLogger CF altimeter and a Raven3 altimeter. The altimeters will only be connected to their independent power supplies and the deployment ejection charges. 2.3 Payload Summary The rocket payload, the Red Hawk, will be composed of two, physically separate systems. The Drag Flap Analysis Module will focus on accomplishing Task 3.1.6, “An Aerodynamic Analysis of Structural Protuberances.” The Environmental Analysis Module will focus on accomplishing Task 3.1.1, “Atmospheric Measurements.” The Drag Flap Analysis Module will chart the movement of the launch vehicle and compare the computer projected path of the rocket to the actual movement of the rocket. After the motor has completed its burn, the on­board computer will determine the angle at which the flaps should be deployed in order to apply an appropriate drag force to hit the target apogee. After deployment of the four drag flaps, barometric pressure sensors below the flaps will take readings of the pressure behind these structural protuberances and compare this data to readings forward of the drag flaps. The Environmental Analysis Module will take readings from several sensors to measure ambient conditions during the rocket’s descent. There will also be a gravity­guided camera system to take the required pictures during descent and after landing through a plexiglass section of body tube. The camera will be oriented correctly through a weight that allows the camera to rotate around the center axis.

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3. Changes Made Since Preliminary Design Review 3.1 Vehicle Changes The vehicle fins have been detailed for their correct design. The recovery system has a more detailed design, including updates to the drift calculations to reflect changes to the mass statement, parachute size, and other factors. Further analysis and calculations have been made to test for vehicle strength, including a buckling analysis. 3.2 Payload Changes Some structural design changes that have been made to the drag flap system include an enlargement of the holes and slots in the entire system, an added ridge on the outside of the drag flap sleeve, and design clarifications with respect to the drive shaft. The added ridge ensures that the drag flaps have adequate room to open without hitting the body tube, and, in installation, the ridge gives the body tube a stopping point when connecting it to the drag flaps sleeve. The slots of the drag flaps sleeve were enlarged in order to assist in the assembly of the printed parts with respect to inserting the struts of the flaps into the inner connections. The electronics for the scientific payload have been separated into two boards that will be connected together. The size of the circuit would outsize the cavity width if put onto a single board. The new general structure of the circuit can be viewed in Figure 7 in Section 5.1.2. The 5V voltage regulator has been replaced with a small circuit that performs the same function with simpler wiring and construction. 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. The funding plan has been updated to include the most recent fundraising successes with the WeR Gold program and the RPI School of Engineering. The educational engagement plan has been updated to include the cancellation of several of the RRS’s events and the inclusion of several new engagement events, including a National Manufacturing Day event and events with local schools.

4. Vehicle Criteria 4.1 Design, and Verification of Launch Vehicle

4.1.1 Mission Statement, Requirements and Mission Success Criteria The launch vehicle will safely power the payload to an apogee of one mile above ground level with a purposeful overshoot of the target by the rocket, allowing for an apogee correction by the Drag Flap Module during the coast phase of the launch. The launch

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vehicle will subsequently be safely and gently lowered to the ground by a dual­deploy recovery system. The rocket will launch with a course that would lead to an apogee between 5,280 ft and 5,600 ft if left unaltered. This requirement will be successful if the vehicle travels between 5,000 ft and 5,600 ft, leaving room for error resulting from error in the payload. To facilitate the measurement and reading of this requirement, a barometric altimeter with the audible beep delivery system will be used for scoring purposes.

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

The launch vehicle will be powered by a single stage. The single stage will use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR) and Tripoli Rocketry Association (TRA). The motor will be less than an L­class motor and will be launched by a standard 12 volt direct current (DC) firing system. This requirement will be met by the design parameters of the rocket if it is powered by one stage, if the motor used is approved by the NAR and TRA, if the motor is less than an L­class motor, and if the motor is capable of being launched on a standard 12­volt DC firing system. The recovery system of the launch vehicle will be electronic dual­deploy with a drogue parachute deployed at apogee and a main parachute deployed at a much lower height afterwards. The stages will be held together by removable shear pins, and, at landing, each independent section of the launch vehicle will have a kinetic energy of less than 75 ft­lbf. The requirements will be met if the launch vehicle successfully deploys its drogue parachute at apogee and the main parachute much later, after breaking the shear pins that hold each parachute in place. The kinetic energy requirement will be met in the design parameters of the parachutes and if the launch vehicle is able to sustain the landing forces associated with its kinetic energy at landing. The recovery system electrical circuits will consist of redundant altimeters that are physically and electronically separate from any payload electronics and power supply. Each altimeter will have a designated power supply and arming switch. These requirements will be met by the launch vehicle design. 4.1.2 Major Milestone Schedule The schedule of major milestones for the launch vehicle’s construction can be found in Section 6.3, which details the entire project plan.

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4.1.3 System Level Design The launch vehicle consists of three subsystems: recovery, aerodynamics and flight stability, and structural. The recovery subsystem encompasses the recovery electronics, ejection charges, shear pins, parachutes, and all features necessary for proper recovery of the launch vehicle. Information about the recovery subsystem will be covered in detail in Section 4.3. The aerodynamics and flight stability subsystem encompasses the selection and preparation of the nosecone and the fins. The main goals of this subsystem are to maintain a low drag coefficient, to monitor the stability of the rocket with respect to the center of pressure and center of gravity, and to integrate the launch vehicle with the recovery and payload sections with respect to the aerodynamics of the launch. The structural subsystem encompasses the analysis of stresses and forces on all of the rocket body components. The structural subsystem will ensure that the launch vehicle is able to withstand the forces associated with the launch and recovery of the rocket and payload. The aerodynamics and flight stability subsystem is a critical portion of the development of the launch vehicle, as well as its integration with the payload. A standard, high­strength plastic nose cone was chosen as the rocket nose cone, as seen in Figure 1. The nose cone will be supplied from Public Missiles Ltd. (PML) with a 3.9 inch diameter shoulder and 16.75 inch exposed length. This design was chosen over high­speed nosecones because of the projected maximum velocity of the rocket. This launch vehicle should not exceed Mach 0.8. Standard shaped nosecones have a proven record of flying faster than Mach 1. The rocket will not near the speeds necessary to cause a change in the aerodynamic profile of the nose cone.

Figure 1: Nosecone Selection

The fin shape has remained much the same since the design in the PDR. Four ⅛ in thick fiberglass were custom ordered to match design specifications, which will be discussed in detail in Section 4.8. They feature a swept back design for greater projected apogee and a relatively small surface area to avoid overstabilization of the vehicle. The final fin design can be seen in Figure 2.

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Figure 2: Selected Fin Design

The structural subsystem of the rocket and its integration with the payload are both critical portions of the development of the launch vehicle. As discussed in the PDR, the primary structural component of the vehicle will be phenolic resin tubing from PML. This material offers ample strength to the airframe, warping resistance, and a surface conducive to adding a layer of fiberglass if additional reinforcement is needed. The structure subsystem of the launch vehicle continues to use the 4.0 inch outer diameter rocket body and 54mm motor tube, as described in the proposal and PDR. In order to accommodate increased size requirements by the payload team, the rocket remains about 104 inches long. A schematic of the rocket structure can be seen below in Figure 3. More detailed drawings can be found in Appendix B.

Figure 3: Launch Vehicle Structure

This model has been analyzed through simulation results in OpenRocket. As detailed in section 4.4, our model is simulated to reach an apogee of 5400 ft above ground level. Additionally, the subscale vehicle modeled after the full scale launch vehicle produced simulation and test results which help to provide a proof of concept and guide the construction of the main project. Specifications and details of the subscale vehicle can be found in section 4.2. The selected motor is the Aerotech K1103X, which has a simulated motor thrust curve shown in Figure 4.

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Figure 4: Thrust Curve of Aerotech K1103X Motor (Thrustcurve.org)

This selection was made with consideration for the drag flap payload system. The K1103X features a high thrust over a short burn time, which will allow for a longer coast time and thus lower deployment angle of the flaps. As a trade­off, this motor will apply much higher forces during launch versus other motors of comparable total impulse. The RRS plans to mitigate this with additional reinforcements to the fin capsule and motor tube components of the structural subsystem.

4.1.4 Design Verification of System Level Requirements

Requirement Design Feature

Verification that Design Meets Requirements

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

Rocket Mass, Rocket Motor, Design

OpenRocket Simulations based on a 5 lb payload mass with the selected motor projects the rocket design to reach an apogee of 5,440 ft.

Reusable

Body Strength (fins,

airframe, parachutes,

etc.)

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

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

damage. The subscale rocket design showed a distinct ability to withstand the

launch forces, even with a similar length and fin structure. Four or fewer Independent Sections

General Design

Three independent sections (nosecone section, payload/avionics section, lower body section) were designed (and mirrored in the subscale launch).

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Single Stage General Design,

Motor Design

Rocket designed with only a single stage motor. Initial construction follows this design, and the subscale launch mirrored this full­scale design.

Commercially available solid

motor propulsion

system under Class L

Motor

OpenRocket Simulations based on a given 5 lb payload mass with the selected rocket design allow for a motor well under Class L motor

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

Capable of Launch by 12 V

DC firing system

Motor, motor retainer

An Aeropack motor retainer that allows for access to motor was selected, and the selected motor is fully capable of being launched with 12 V DC firing system. The subscale launch mirrored this requirement with an Aerotech G80 motor that was launched on a 12 V DC firing system

Electronic Dual Deploy

General Recovery Design

(parachutes, shock cord, ejection charges, etc.)

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

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

Drogue Deploys at

Apogee, Main Parachute Deploys at much lower altitude

Parachutes, ejection charges, General Rocket Design, Altimeters

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

these deployment locations. The subscale launch mirrored these deployments with parachute

deployment at similar locations to the full­scale design.

Shear Pins hold rocket sections together until Parachute Deployment

Shear Pins, Ejection Charges

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

strengths were also accurately calculated for the full scale rocket design. Ejection charge strengths for the subscale rocket were calculated in the same way as the full­scale rocket design, and successfully deployed the

recovery system.

Independent Sections have less than 75 ft­lb of KE at Landing

General Rocket Design (Mass),

Parachutes

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

Parachutes and Recovery system components show no signs of damage upon order arrival. The subscale rocket launch landed safely with no signs of independent sections landing with more than 75 ft­lb of KE at landing in

relatively difficult launch conditions.

Redundant, Safe Altimeters

Altimeters, Supporting Recovery Electronics

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

design has an independent power supply to each altimeter. These altimeters are being stored safely, along with all other supporting electronics components. The subscale recovery system operated as

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expected with installation similar to what will be done on the full­scale launch vehicle.

Table 1: Verification of System Level Requirements 4.1.5 Workmanship The workmanship of any construction project directly correlates to mission success. Workmanship includes the use of safe construction practices in order to produce a quality project using the best standard industry operating procedures With this in mind, the RRS has made it a key goal to maintain the highest levels of workmanship in the construction of its launch vehicle. In particular, it is critical for the RRS to maintain solid workmanship throughout the launch vehicle construction because of the high impulse motor that has been selected for use in this project. Some of the ways that the RRS will strive towards good workmanship include emphasizing the use of personal protective equipment (PPE), using epoxies and tools in the proper manner, and an emphasis on constructing the launch vehicle according to the predetermined design with quality construction methods based on the RRS mentor’s significant experience. The team’s use of workmanship with respect to PPE and use of epoxies and tools in a safe manner is discussed in depth in later sections. Some of the techniques that the RRS will use to facilitate quality construction and workmanship of the launch vehicle include proper epoxy use, a strengthened installation of the fins and motor tube, and ample testing of all critical components of the launch vehicle. Proper epoxy use consists of selecting the correct epoxy to bond two materials for their particular launch environment, sanding each surface to provide ample quality bonding surface, and vigorously mixing both parts of the epoxy in correct ratios for a enough time to ensure that adequate cementing of the epoxy can occur. Testing of each component depends on the use of that individual component. Some examples of this testing includes load testing of samples of body tube not to be used in construction, electronics testing of all circuits and altimeters to be used in the recovery system, and ensuring all components are not significantly warped or damaged based on visual inspection before their installation. The RRS will also rely on a common rocketry construction process for installing the fins and motor tube that has been used by the RRS numerous times and highly recommended by the club mentor. This technique involves epoxying the centering rings on the motor tube, then attaching the fins to this assembly with an accurately constructed fin alignment guide made with high­precision power tools from sheets of plywood. This “fin can” is then secured with additional epoxy or fiberglass sheets before insertion into the lower body tube with epoxy. The techniques used by the RRS and recommended by the club mentor will be followed closely in the construction of the launch vehicle with quality workmanship.

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4.1.6 Planned Additional Testing

The majority of testing for the vehicle can be split into simulation results, component testing, and scale testing. Most of the simulations have been completed, and with detailed results found in section 4.4. Component testing will occur throughout the construction and assembly process. As parts arrive, they will be tested for any defects and basic structural strength. When subsystems such as the fin canister in the lower airframe are assembled, several evaluations will be performed to ensure proper alignment and appropriate structural reinforcement. Details of the subscale launch and results will be found highlighted in section 4.2. Additionally, as outlined in the requirements for the FRR, the completed design will undergo a full test flight. This full scale test in ensure that all systems are functional and secure in preparation for the competition launch. 4.1.7 Design Integrity

The vehicle will have four fins spaced equally apart, with exact measurements specified in Figure 5.

Figure 5: Exact Dimensions of ⅛ in Thick Fiberglass Fin

This shape was selected in order to obtain a higher simulated apogee. Including a sweep to the trailing edge of the fin produced the best results versus other possible shapes. Each fin measures 7.75 inches from the top of the leading edge to the tip of the trailing edge, with a span of 2.75 inches. The fins have this relatively small surface area to avoid overstabilization. The launch vehicle itself is very stable due to its length and does not need additional stability from the fin structure.

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The fins are ⅛ inch thick G­10 fiberglass material, which provide ample strength and mitigate the chance of warping phenomena during flight. These fins are custom ordered from PML, and require little additional work. The main body of the rocket and motor mount use phenolic resin tubing supplied from PML. When specific lengths of tubing need to be cut, care will be taken to ensure that the used section suffers no structural damage. All other structural components, including bulkheads and centering rings will be inspected for any defects, and properly handled to ensure they are not damaged. When constructing all portions of the launch vehicle, following proper procedures and maintaining correct alignments is crucial to design integrity. The construction of the vehicle breaks down into three main sections in this regard: construction of the motor mount/fin can, construction of the payload bay around the Lexan chamber, and remaining assembly. For the motor mount, each of the centering rings must be secured exactly level. To this end, exterior sleeves will be placed on the motor mount to guarantee that each centering ring is attached correctly. Once the centering rings are in place, the fins must be carefully attached. If the fins are not exactly equidistant from each other or perpendicular to the motor tube, there could be complications with inserting the fin can or instability during flight. To assist proper alignment of the fins, a guide has been made out of carefully machined plywood. Two layers of the plywood guides will be used in parallel to achieve fin alignment. Assembly of the portion of the vehicle around the Lexan tube for the camera viewing bay also requires additional care for proper alignment and attachment. Since the Lexan tube is only available with an inner diameter of 3.75 in, a machine lathe will be used to expand this to match the 3.9 in inner diameter of the phenolic resin body tubing. In doing so, the RRS must ensure that the material comes off in an even manner so that couplers provide a strong link between the Lexan and body tube. For the remaining portions of the rocket, following appropriate construction and assembly procedures will ensure that all attachments are strong and aligned. Connections for main chute separation, drogue chute separation, and payload bay loading will undergo heavy scrutiny for strength and reliability during flight with the ability to properly separate at the correct time. Having such a high impact motor as the Aerotech K1103X, sufficient motor mount reinforcement and motor retention becomes an area of concern. The motor mount itself is a 54 mm diameter phenolic resin tube that will contain the motor, along with three centering rings used as structural and alignment support. In order to reinforce this, the fins will be attached between the lower and middle centering rings. Additionally, the fillets along both sides of each fins will have a layer of fiberglass added for more structural support. The motor will have a standard threaded aluminum retainer purchased from PML. Details pertaining to verifications and their corresponding statuses can be found in section 4.1.4

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Detailed dimensional drawings of the launch vehicle, along with its subsystems and major components can be found in Appendix B.

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

Part Mass (oz) Quantity Subtotal (oz)

Centering Ring 2.6 3 7.8

Motor Mount 4.83 1 4.83

Motor Retainer 1.42 1 1.42

Motor 54.42 1 54.42

Body Tubing (44”) 18.7 1 18.7

Drogue Chute 6 1 6

Shock Cord 9.62 1 9.62

Fins (x4) 2.86 1 2.86

Subtotal ­ ­ 6.603 (lbm)

Table 8: Mass of the Lower Airframe

Part Mass (oz) Quantity Subtotal (oz)

Nosecone 10 1 10

Body Tubing (31”) 13.18 1 13.18

Polycarbonate Tubing (8”)

8.48 1 8.48

Bulkheads 4 3 12

Coupler Sections 2.6 2 5.2

Shock Cord 9.62 1 9.62

Main Chute 13.3 1 13.3

Drag Flap System 17.39 1 17.39

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Payload Estimate 16 1 16

Subtotal ­ ­ 6.573 (lbm)

Table 9: Mass of the Main Airframe

Total Mass: 13.176 lbm

As illustrated in tables 8 and 9, this mass estimate of is based on a sum of the masses from the parts that will be used in the rocket. As such, the accuracy of this measurement is limited by data provided from part manufacturers and distributors. Additionally, some parts may be modified during the construction process. This will result in the reduction in mass of some components and the addition of mass on others. Though the design actually saw a reduction of mass since the PDR, this can largely be attributed to a different selection of parachutes which are far lighter than previous parachutes under consideration. Additionally, there still exists a small growth margin in the construction phase for materials not listed in these tables, such as epoxy and fiberglass reinforcements. As such, the actual mass of the rocket is expected to grow around 10­15% from this estimate.

There is a very large margin of mass until the vehicle becomes unable to launch with the selected propulsion system. Besides this hard constraint, the rocket must strictly exceed the target altitude in order for the drag flap system to function. Given this restriction, the RRS launch vehicle has a margin of about five additional pounds before the simulations indicate that the launch vehicle would just reach the target height.

The RRS analyzed the launch vehicle for safety and failure levels with a high level of engineering calculations. The RRS looked at the strength of the rocket with respect to stress­strain, material analysis, and buckling.

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

In a report by Branko Sekulic of the Polytechnic University of Catalonia, the compressive yield strength of cardboard is about 13 MPa, or 1885.5 psi. When compared to the

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expected stress that the launch vehicle will face of 485.6 psi, there is a factor of safety over 3.8.

A complex buckling analysis was performed on the launch vehicle due to the uncommon boundary conditions. The launch vehicle was assumed to be approximately a beam of length 104 inches that was made of purely the Phenolic body tube. The analysis shown in Figure 6 is the derivation of the governing equation for the buckling of the rocket. This equation provides the amount of deformation in the transverse direction due to the buckling of the rocket. Initial solutions of this equation provide approximately zero deformation due to buckling. The RRS worked with an RPI professor to derive this equation and is currently working with this professor to solve the equation with more advanced methods to verify the initial conclusions.

Figure 6: Buckling Analysis of Launch Vehicle

4.2 Subscale Flight Results

The subscale flight was launched using a heavily modified Estes Pro Series II Leviathan. The fin structure of the Leviathan was remarkably similar to the proposed fin structure for the full scale launch vehicle. The kit was modified to include a much longer body tube to simulate the longer body of the full scale rocket. In addition, the modifications allowed for the requisite dual deploy system. The rocket assembly can be seen in Figure 8. The

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dual deploy recovery system in the subscale launch was an electronic dual deploy system. A Stratologger SL100 altimeter was used along with its supporting electronics: the power supply, terminal blocks, screw switches, and wiring. This electronics system powered the initial and main deployments with black powder ejection charges, with masses calculated in the same manner as the full scale recovery system’s ejection charges.

The subscale flight was modeled beforehand on the same OpenRocket software that is being used for the full scale launch. The model is depicted in Figure 6, with the simulated flight graphed in Figure 7.

Figure 6: Sub­Scale Launch Vehicle Model

Figure 7: Simulated Sub­Scale Launch

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The actual subscale launch vehicle flew higher than predicted. The OpenRocket predicted apogee was about 1004 ft. The actual apogee was 1237 ft. The flight data shows that the first recovery event occurred at the apogee, and the second recovery event occurred very close to the predicted altitude of 700 ft. All in all, the actual subscale flight was relatively similar to the predicted flight, and the differences in the data can be attributed to low winds experienced on the day of the actual flight, as well as error in the modeling mass and aerodynamic profile of the rocket.

Figure 8: Sub­Scale Rocket at Launch

The subscale flight has provided significant impact to the work on the full­scale launch vehicle. There are now quantitative results in addition to the engineering analysis to show that the general aerodynamic and structural profile of the full­scale launch vehicle is stable. Additionally,

4.3 Recovery Subsystem 4.3.1 Avionics 4.3.1.1 Flight Computers

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The main electrical components of the recovery system consist of a Perfectflite Stratologger SL100 (Stratologger) barometric altimeter and a Featherweight Raven3 (Raven) barometric and accelerometer altimeter. Both the Raven and the Stratologger are programmed to deploy the drogue parachute at apogee, and to deploy the main parachute at 800 ft AGL during descent. Two altimeters provide system redundancy in the event that one flight computer fails. Table X below details each altimeter’s technical specifications. Table 10: Hardware Specifications

Operating Voltage

Dimensions Sample Rate Altitude Accuracy Maximum Altitude

Stratologger SF100

4­16 V 0.90” W 2.75” L 0.50” T

20 Hz +/­ 0.1% 100,000 ft

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

20 Hz +/­ 0.3% 100,000 ft

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

Figure 9: The Perfectflite Stratologger SL100

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

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Figure 10: The Featherweight Raven3

4.3.1.2 Mounting and Electrical Layout The altimeters and accompanying hardware will be mounted to a sled in the avionics bay, in a way that does not interfere with the altimeters ability to make accurate readings. Specifically, the bay will have static ports so that the pressure in the bay is equivalent to the atmospheric pressure, and the Raven will be mounted so that it’s length is parallel to the body of the rocket, allowing the accelerometer to function properly. Figure X is a block diagram detailing the recovery system circuitry.

Figure 11: Recovery System Electronics Schematic

4.3.1.3 Arming The recovery system electronics are armed through two external, independent switches. Two screw switches externally located on the payload section allow each altimeter to be armed on the pad before launch, the prevent premature firing of ejection charges while the rocket is being handled and set up on the launch pad. 4.3.2 System Level Design

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4.3.2.1 Parachutes The parachutes chosen for the dual­deploy recovery system are the SkyAngle Classic II 52” main parachute from B2 Rocketry and the Ballistic Mach II 2 ft drogue parachute from Rocketman. The main parachute is constructed of zero­porosity 1.9 oz. balloon cloth with ⅜” tubular nylon suspension lines sewn around the canopy, rated for 950 lb. The suspension lines meet at a heavy duty 12/0 nickel­plated swivel joint rated for 1500 lb. Data gathered from OpenRocket simulations indicates that the maximum force on the main parachute and associated hardware will be approximately 90 lb. This equates to the parachute and suspension lines having a factor of safety of 10. The factor of safety is excessive, but to achieve the desired descent rates and kinetic energy, a parachute of this size must be chosen. The drogue parachute is constructed from ballistic­grade rip­stop nylon. Calculations estimate that the maximum force on the parachute and associated hardware will be 13.75 lb. The drogue parachute was chosen for it’s size and manufacturer reputation. The strength of the parachute is not a concern because of RRS's experience flying Rocketman parachutes and their trusted quality. 4.3.2.2 Harnesses and Attachment Hardware The shock cord that tethers the parachutes to the airframe will be ½” tubular nylon cord. The cord for the main parachute will be connected to a 1.5” forged steel eyebolt. The drogue parachute will be tethered to the body in a similar manner, connecting to a 1.5” forged steel eyebolt on the aft bulkhead on the main airframe, and to a 1.5” forged steel eyebolt on the forwardmost centering ring of the motor housing. For each parachute there will be 300 inches of shock cord to allow for adequate separation of vehicle components during descent. The ½” tubular nylon cord is rated to 1000 lb. The forged steel eyebolts are rated for 1200 lb. Using the above estimate of the maximum force experienced on the hardware of 90 lb, there is no concern of attachment hardware failure during parachute deployment. 4.3.2.3 Bulkheads The eyebolts that serve as the parachute’s attachment hardware will be mounted to bulkheads at both ends of the main airframe. The bulkheads will be made of ¼” birch plywood. A schematic of each bulkhead can be seen below in figure X.

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Figure 12: Forward (left) and aft (right) bulkheads (not to scale).

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

Figure 13: Bulkhead hardware arrangement

Using the above estimate of 90 lb in the eyebolt, a strength analysis of the bulkhead can take place. The load will be transferred to the bulkhead from the nuts that secure the eyebolt in place, creating bearing stress on the bulkhead. Using the area of the nuts, the stress is calculated to be 3.80 MPa. Applying a factor of safety of 1.7, the NASA factor of safety for critical components, the stress becomes 6.47 MPa. A simplified finite element

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model was created to analyze the bulkhead, with a uniform pressure of magnitude 6.47 MPa applied to the bulkhead over the area simulating the nut’s bearing load. The boundary conditions in the model reflect how the bulkhead will be supported in the rocket; fully constrained on the outer edge, and with fixed rotation, and fixed X and Z displacement in the central hole. Figure X illustrates the stresses in the bulkhead with the load applied.

Figure 14: Von Mises stresses in forward bulkhead

The finite element analysis (FEA) indicates that the maximum von Mises stress in the bulkhead is 31.9 MPa. The ultimate stress of plywood can be difficult explicitly measure, due to the nature of being a composite, since the manufacturing processes used to create the composite will greatly impact quality. Typical values range between 31.0 MPa and 41.0 MPa however. Assuming an ultimate strength of 31.0 MPa, the design’s factor of safety is slightly less than 1.7, but is still a safe design. The second bulkhead is the one that houses the eyebolt for the drogue parachute. It is located at the lowest point of the main airframe. When the drogue parachute deploys, there will be a load applied to the eyebolt just as there is for the forward bulkhead when the main parachute deploys. However, from above, the maximum load that the aft bulkhead will experience from the drogue parachute is 13.75 lb. Due to the fact that the load is considerably less than the load on the forward bulkhead, a strength analysis of the aft bulkhead is not necessary. Being made from the same material, in a nearly identical arrangement with a considerably smaller load indicates that the design is as safe or safer than the forward bulkhead. 4.3.2.5 Ejection

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Aluminum blast cups will be placed on both the forward and the aft bulkheads of the main airframe. These cups will be filled with black powder and ignited by E­matches connected to the altimeters. The ejection charges for the parachute deployment require 1 g of black powder charge per 11” of body length containing the parachute. Approximately 1.7 g of black powder will be required for the main parachute deployment, and 2.1 g of black powder for the drogue parachute deployment. At each ejection event, the altimeters will go send a signal to the E­matches to ignite the black powder charges. The charges will separate the airframe and allow the unfurling of the parachutes. Both parachutes will be protected from the heat effects of the charge firings with kevlar parachute protectors. Drogue parachute deployment will occur at apogee and main parachute deployment will occur at 800 ft AGL. 4.3.3 Discussion of Kinetic Energy and Drift 4.3.3.1 Kinetic Energy Analysis At the maximum predicted velocity of the rocket, the lower airframe will have a kinetic energy of approximately 35484.2 ft • lb, and the main airframe will have a kinetic energy of 62978.5 ft • lb. At the deployment of the main parachute, the main airframe will have a kinetic energy of 367.9 ft • lb. The lower airframe will have a kinetic energy of 207.0 ft • lb. When the rocket touches down, the main airframe will have a kinetic energy of 52.2 ft • lb, and the lower airframe will have a kinetic energy of 29.4 ft • lb. Both components of the rocket will land safely with kinetic energy below 75 ft • lb. 4.3.3.2 Drift Analysis For the Upper and Lower Body Table 10 lists the drift distances with respect to the wind speeds for launch. The calculations were made assuming a one degree launch angle, with the wind blowing in the same direction as the launch angle.

Table 11: Drift Analysis at Various Wind Speeds and Vertical Launch

Wind Speed Drift Value

0 mph 211 ft

5 mph 782 ft

10 mph 1209 ft

15 mph 1912 ft

20 mph 2150 ft

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The calculations show that for any launch below 20 mph, the rocket will land safely in the one square mile field to be easily recovered. 4.4 Mission Performance Predictions 4.4.1 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 the drag flap system to decrease the coasting velocity. Additionally, the vehicle must land within a reasonable distance of the launch site. To this end, the recovery system will be tailored so that the main parachute is not deployed too early, mitigating drift.

4.4.2 Vehicle Simulations Using OpenRocket with a motor configuration for the Aerotech K1103X, a simulated model of the rocket was produced as shown in Figure 8.

Figure 15: Simulated Model for Flight Simulation

On this model, there is a stability margin of 6.12 Calibers with the motor included. The simulated center of mass is approximately 48 inches down from the top of the nosecone, and the simulated center of pressure is approximately 72 inches from the top. Using this model and the selected motor, a flight simulation was created which is shown in Figure 9.

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Figure 16: Flight Simulation from OpenRocket

Note that this simulation does not take into account the effects of the drag flap system. Given that, the apogee reached is 5400 feet, with a maximum acceleration of 677 feet/second2, and flight time of 132 seconds.

4.5 Payload Integration

The payload will be integrated into the overall vehicle using a payload bay in the center of the rocket. This will consist of an electronics bay and two parachute bays, separated by bulkheads. The drag flap module will also be located within his section. A ridge on the sleeve of the drag flaps module will ensure that the printed sleeve will fit smoothly into the body tube of the launch vehicle without slipping through the tube too far. The first bulkhead at the rear of the rocket will have the drogue attachment point and deployment system. The drag flap module is forward of this bulkhead. The electronics bay will be forward of the drag flap module so that the processing segment of the electronics is adjacent to both the motor and the camera system, minimizing wire length needed. The camera assembly section will be located forward of the electronics, and will consist of the swing arm and related connection structure. So that the rods forming the main supports of the aft segment of the payload bay do not interfere with the camera’s view, an additional bulkhead will be located immediately aft of the plexiglass window. This will also serve as the attachment point of the swing­arm assembly to the rocket

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superstructure. Another bulkhead will be located forward of the camera structure, to which the forward parachute may be attached. The bulkheads will have shock cord and parachute attachment points in the form of forged eyebolts and quick links to facilitate easy access to the payload and electronics, as well as aluminum charge cups to hold the black powder of the ejection charge. There will also be connection points for the electric matches that will trigger these ejection charges. Electronics arming and disarming will take place using screw switches accessible from outside the rocket. Shear pins will connect the aft segment of the overall payload bay to the base of the rocket. These will sever at the first charge event to deploy the drogue, and a similar mechanism will deploy the main parachute at 800 feet above ground. Data from the electronics section will be transmitted back to the ground station via an antenna. Actual physical modifications to the payload will require removal of the electronics sled as only the arming switches allow an external action to influence the payload. 4.6 Safety and Environment

4.6.1 Safety Officer The RRS has identified Chris Andre as the safety officer. However, Chris Andre will relinquish his duties in January 2016 and take leave from RPI for an internship. Elections for a new safety officer will take place as soon as classes for next semester begin. The safety officer’s responsibilities include ensuring shop safety and hazardous material procedures, which are partly accomplished through safety quizzes administered by the RPI School of Engineering and MANE Department. The safety officer will oversee the safe construction and launch of the pertinent rocket vehicles through supervision and inspections. 4.6.2 Preliminary Checklists The safety officer and team mentor will oversee the final assembly of the rocket and its subsystems prior to flight, as well as the launch pad preparation operations for all RRS launches. Two preliminary checklists have been developed for each phase that cover the subscale, full­scale, and final launches. The pre­launch operations checklist broadly covers the assembly of the major structural components prior to the rocket arriving at the launchpad. The launch pad operations checklist covers the procedures to be completed prior to launch. The checklists are detailed in tables 11 & 12.

Category Description Required Actions

Parachute Deployment

Protective insulation will be inserted between the ejection charges and the parachute to prevent burning, and the parachute will be properly folded and

protective insulation insertion, parachute folded, parachute inserted.

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then inserted.

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

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

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

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

Drag Flaps Drag flaps will be checked for unobstructed movement.

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

Table 12: Pre­Launch Operations Checklist

Category Description Required Actions

Rocket Assembly Rocket assembly is complete.

RSO Safety Check RSO will check final rocket before launch.

RSO Approval. RRS SO Approval. RRS Mentor Approval.

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

Data link is clear.

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

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

Ejection Charges Ejection charges will be installed.

Ejection charges installed

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

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

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

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

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

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

Ejection Charges Ejection charges will be armed.

Ejection charges are connected and will not short.

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

RRS SO Approval. RRS Mentor Approval.

Data Retrieval Final altitude will be retrieved.

Altitude has been retrieved by SL official.

Post­Flight Inspection Inspection will be performed on rocket following flight.

RRS SO Approval. RRS Mentor Approval.

Table 13: Launchpad Operations Checklist 4.6.3 Failure Modes The following table summarizes the launch and flight failure modes that have been identified. They are addressed with a mitigation plan and a risk likelihood/impact (range of 0­3 and 0­5, respectively). The status of each mitigation plan is noted.

System Failure R/I Mitigation Status

Motor Dislodgement from housing

1/5 Employ retainer structure with three bulkhead centering rings.

Factored into design

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.

In progress

Stepper motor malfunction

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.

In progress

Forward flap system causes instability

1/5 Verification of stability will be performed in OpenRocket and using test launches.

In progress

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Structural Fins shear off during flight

2/4 Fins will be embedded in body and fixed to the internal motor mount and body with heat­tolerant epoxy resin.

Factored into design

Rocket motor explodes

1/5 Test launches, finite elements analysis, and static load tests will be performed to verify integrity of rocket body.

In progress

Parachute tears, parachute fails to deploy

2/5 Parachute will be checked for defects before packing. Parachute will be packed properly.

Checked at launch

Shock cord breaks

2/5 Recently­manufactured shock cord will be used to ensure integrity. Shock cord will be properly secured and tested under static and jolt loads.

Shock cord acquired. Testing in progress

Drag flaps shear off during flight

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

Testing of drag flap components has begun

Recovery Parachute ejection failure

2/5 Redundant powder charges will be used for the main parachute. Packing will be inspected before launch.

Checked at launch

Parachute ejected prematurely

1/3 Code will be reviewed for correctness. Hard failsafes will be implemented to prevent ejection above a predetermined altitude.

In progress

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.

In progress

Transmitter failure/interference

2/3 Transmitter will be tested at range. Interference will be studied.

In progress

Launch operations and transportation

Physical damage during handling

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.

At launch

Live charges on board after

1/3 Avionics will be powered off. Discrete switch will cut off system power

Factored into design

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launch command

entirely. and launch procedure

Premature firing of ejection charges and/or motor

1/5 Wiring will be ensured to be 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.

Factored into design, will be checked at launch

Spectator injury

1/5 Range will be cleared prior to launch. No large, heavy objects will be installed according to NAR safety code.

Factored into design, checked at launch

Table 14: Failure Modes

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

Hazard Mitigation Required Safety Equipment

Emergency Equipment

Drill press rotation No loose clothing/hair

Glasses First aid kit

Band saw No loose clothing/hair

Gloves when working with large pieces, glasses

First aid kit

Epoxy/Paint fumes/irritation

Work in ventilated, spacious area

Face masks, gloves, glasses

First aid kit, eye flushing station

Propellant/powder Motor will be stored Gloves, glasses First aid kit, eye

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ignition and skin irritation

in separate, insulated cabinet. Motor will be handled with gloves after it has cooled.

flushing station, burn kit

Vehicle debris NAR regulations will be obeyed

Glasses First aid kit

General cuts and irritation

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

Glasses, gloves First aid kit

Table 15: Hazard Analysis

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

Risk Mitigation Likelihood

Inoperable Wind Conditions / Cloudy

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

2

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

2

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

3

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

1

Table 16: Environmental Concerns

4.7 Launch Operation Procedures 4.7.1 Recovery Preparation Insulation will be placed within the rocket body so that the parachute is not damaged by the ejection charge. The parachute will be folded so that it may unfurl without tangling. The ejection charges are to be properly filled and sealed. The charge trigger altimeter is

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to be tested on the ground before a launch to verify that it is programmed and wired correctly. 4.7.2 Motor Preparation The appropriate motor delay will be selected based on previous simulations. The motor will be mounted within the casing and the integrity of the interface with the rocket body will be checked. The retention cap is screwed on. The igniter charge is installed securely before launch. 4.7.3 Launcher Setup The rocket will be slid onto the launch rod horizontally and then raised to a vertical position. 4.7.4 Igniter Installation The igniters will be installed shortly before launch, after the rocket is situated on the launch rod. The igniters will be carefully handled and shorted such that static discharges do not trigger the igniters. The igniter will be fully inserted and stopped off so that it is firmly held in place. The igniter will then be connected to the launch console. 4.7.5 Troubleshooting A static test of the flight computer telemetry will be conducted on the pad to ensure that the radio link between the RRS team and the rocket is operational. The flight computer will be checked for full functionality and sensor readings will be checked to ensure that they are connected/functioning properly. 4.7.6 Post­Flight Inspection The rocket body will be checked for visible structural damage. The flight computer will be tested for functionality and the wiring checked. The interfaces between separating fuselage components will be checked for damage, and the parachutes and shock cords will be checked for tears and burns. The drag flap system will be checked to ensure that the actuation system has not become damaged under the force of landing. The competition altimeter will be read by a competition official.

5. Payload Criteria 5.1 Testing and Design of Payload Equipment 5.1.1 Drag Flap Subsystem The drag flap subsystem of the rocket will be used to accomplish multiple requirements simultaneously. These are Vehicle Requirement 1.1, and Payload Requirement 3.1.6. Vehicle Requirement 1.1 states that the rocket must deliver the payload to an altitude of 5,280 feet. The rocket motor has been selected such that if the rocket were to fly without the drag flaps, the final altitude would be slightly above this height. The drag flaps will be dynamically controlled in­flight to decrease the vertical velocity of the rocket and adjust the maximum height of the rocket to exactly 5,280 feet at apogee.

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

Figure 1: Diagram of Drag Flap Subsystem Assembly

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

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shaft is connected to the motor shaft with a small cylindrical coupler made of aluminum. The coupler will grip the drive shaft with epoxy and a press fit with the threads of the drive shaft, and it will connect to the motor with epoxy and a small set screw that grips the flat of the motor shaft. As this slider moves up and down, it causes the flaps to rotate in and out. The flaps are designed to extend to a maximum angle of 37.89 degrees. This maximum deflection results in an increase of 27.37 square inches of additional vertical projected area. Initial work has begun on performing a detailed computational fluid dynamics analysis on this assembly. The some of the goals of this analysis are to develop a function which relates the angle of the flaps to the increase in drag force, to model the pressure distributions around the drag flaps, to simulate the suction effect that is created by the pressure distributions around the drag flaps, and to model the flow velocity distributions around the drag flaps. The RRS has obtained a license for use of a student version of the ANSYS Fluent CFD program. The launch vehicle and drag flap system were modeled in this program all together and in sections in order to maximize the effectiveness of the meshes used in the model. These models were then simulated in high­speed flow, up to the maximum speed that the launch vehicle is projected to reach, approximately 0.8 mach. Since this flow is mainly turbulent, viscous flow, a particular model for the flow must be chosen. The RRS is working with RPI professors in order to further this analysis for accurate flow modeling. For initial simulations, the k­epsilon model was used. An image of the simulation of the flow velocity can be seen in Figure 2, and an image of the simulation of the contours of the static pressure on the drag flaps can be seen in Figure 3.

Figure 2: Contours of Velocity Magnitude

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Figure 3: Contours of Static Pressure on Drag Flaps The motor which actuates the drag flaps will be controlled using an Arduino Zero clone (Neutrino). The Neutrino will be connected to a motor driver, and to one of the accelerometers. The design of the control system is in progress. A simulation software has been

developed to characterize controller performance under an array of varying flight conditions, including body mass, ground air density, launch rod angle, and coefficient of drag. The software is sensor­aware and simulates measurement noise to facilitate design for robustness. Using simulation results from OpenRocket, the simulator's accuracy has been verified. For a simulated flight without controls applied and perfect standard conditions, the observed apogee is 1658 meters (~5440 feet). The state history of the simulated flight is seen below in Figure 4.

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Figure 4: State History of First Simulated Flight A demonstration of the sensor fuzzing capabilities is seen below:

Figure 5: Demonstration of Added Sensor Noise

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

Figure 6: Altitude and Velocity Graph of Possible Control Programs The state domain (altitude and velocity) is plotted in Figure 6. The greyed out regions represent states for which there is no possible control program that can be employed to guide the rocket to exactly 5280 feet. These regions are found by propagating rockets backwards through time from the target height with control values of 0.0 and 1.0. These values represent the range that the flaps are extended, with 0.0 and 1.0 (fully closed, fully open) being the extreme values. The lower region represents states for which there must be negative control (impossible: u < 0.0); these states do not carry enough energy to reach 5280 feet even with no control for the flight. The upper region represents states

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for which there must be control higher than possible (u > 1.0); these states carry more energy than can be bled off in time even with maximal control throughout the flight. The white region therefore represents the states for which there is a valid control program that can be employed to guide the rocket to exactly 5280 feet in ideal conditions. It is termed the region of feasibility, and is unique to the rocket's geometry and weather conditions. The region of feasibility is parameterized in the drag flap drag coefficient, which is assumed to be 1.28 in this case. An example launch simulation is plotted on the graph. The boost phase expectedly propels the rocket past the minimal control region into the region of feasibility. The region of feasibility is seen to narrow significantly; this happens because the control effectivity decreases with dynamic pressure/velocity, which can be seen in the trend along the x­axis. An important point on the graph is the intersection of the flight data with the x­axis (off chart), which represents the final altitude. Both regions of infeasibility converge on the desired altitude of 5280 feet. The example launch overshoots the target, which the competition rocket is designed to do. The reason that the velocity signal derived from the barometric altimeter reading cannot be used effectively without further processing is visible in the previous graph. Note that there is a slim margin between the simulated flight path of the rocket and the lower bound that averages 20 m/s (~60 ft/s). A control system based on the velocity signal may easily shed too much energy and fall into the infeasibility region due to the high randomness of the signal. This is observed in simulations with an active controller as seen in the following figures:

Figure 7: Simulation with Control System Based on Velocity Signal

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Figure 8: Graph of Control System Based on Velocity Signal with Added Noise

Figure 9: Graph of Possible Control Programs Based on Velocity Signal Figure 7 depicts the control program that the rocket creates on­the­fly. The rocket attempts to find an unchanging control value that will guide the rocket to its final apogee

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based on the measured state. In this case, the drag flap motor is not modelled, and so there is extreme chatter in the program that is caused by wildly varying velocity measurements. In Figure 8, it can be seen in the last graph that the expected altitude (blue line) based on measured state varies considerably. The mostly constant segment after 4 seconds represents the time after which the control system is activated. It is able to achieve the target altitude most of the time, however, as time progresses, the control authority is seen to drop. This is expected because the controller is overcompensating due to overestimated velocity points and bringing the energy down, but the undercompensation cannot bring it up at a similar rate. This also suggests that unbounded controllers like PID controllers may cause issues if they are not properly tuned. The rocket undershoots the target altitude. In Figure 9, this failing is visualized on the feasibility graph. The flight path of the rocket crosses the lower bound prematurely, signifying that there is no longer a possible control to reach the target altitude. The intersection point of the flight path on the x­axis coincides with the final apogee of 1588 meters. In the following figures, an example simulation is carried out with perfect sensor readings that targets the ideal 1609 meter apogee perfectly:

Figure 10: State History of Third Simulated Flight

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Figure 11: Graph of Third Simulated Flight with No Bleed

Figure 12: Graph of Possible Control Programs Based on Third Control System Type In Figure 10, it is seen that the controller is applying the expected control. It does not considerably change during the flight. There is no chatter due to perfect sensor readings.

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In Figure 11, there is no bleed in the control authority. In Figure 12, the flight path correctly straddles the region of feasibility and converges along with the upper and lower bounds on the target apogee. These results and tools lay the groundwork for the evaluation of potential control systems. The next step is designing a filter for the measurements so that the state can be estimated accurately. The team has identified PID controllers as a candidate for control, due to their ubiquity, which will soon be designed and evaluated. The motor which will be used to actuate the drag flaps is a bipolar stepper motor. The motor specifications are shown below in Table 16.

Table 16: drag flap motor specifications

Steps per Revolution 200

Supply Voltage 4.5 V

Current per Phase 0.67 A

Holding Torque 950 g­cm

The drag flap subsystem is made up of ten separate plastic pieces, all of which will be 3D­printed using a fifth generation Makerbot. All prints are done using Makerbot’s PLA (Polylactic Acid) plastic. The part’s layer thickness and infill percentage vary based on where the part is located and what types of stresses the part will encounter. The drag flap doors, for instance, are printed at 100 micron layer thickness to allow for a smoother exterior surface and at a 100% infill, essentially making the part a solid piece of plastic. All parts were printed at temperatures between 215­225 degrees centigrade to ensure proper layer adhesion. A few manufacturing issues were encountered while printing the parts. One of these issues was a total failure of a print. Print failure was often caused by filament jams or feed failures. Normally, the print operator can notice an impending failure and correct it; however, due to the long print times of some of the parts (upwards of 15 hours), the print could not always be monitored by the print operator. Another manufacturing and assembly issue was considered is the thermal effects of the printing process. Due to thermal enlargement that is a natural byproduct of the 3D printing process, all holes in the system needed to be enlarged after initial prints. The team performed preliminary finite element analysis (FEA) to verify the strength and robustness of the design, but more detailed analysis will come once the team has CFD data in order to determine the total drag force on the flaps. One challenge in the analysis comes from the anisotropic nature of 3D­printed parts. This challenge has been analyzed using mechanical properties testing.

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Once it came to the attention of the project team that one or more load­bearing components of the final rocket design were to be 3D­printed, it was decided that it was necessary to conduct tests on a sample of 3D­printed material to determine its mechanical properties, particularly its Young’s modulus and yield strength. First, a dogbone sample was 3D­printed such that the grain of the sample was perpendicular to the direction in which it would be loaded. This was predicted to be the weakest possible orientation. The dimensions of this sample are tabulated below:

Gauge Length: Width: Thickness:

20 mm 6.40 mm 3.37 mm

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

Figure 5: Adjusted Load vs Elongation of 3D­Printed Material

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Figure 6: Engineering Stress vs Strain of 3D Printed Material

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

Young’s Modulus:

Yield Strength: Crosshead Displacement Speed:

Extension at failure:

Maximum Load:

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

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5.1.2 Atmospheric Measurement Subsystem

Figure 7: Electrical Schematic of Scientific Payload

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

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

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

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The XBee will be wired to the hardware serial on the second Neutrino. GPS data will be sent in flight. After landing, the primary Neutrino will perform some brief processing before sending out the data logged on the MicroSD card. The GPS module will be wired to the second set of hardware serial ports on the second Neutrino. Data will be saved and transmitted live across the XBee in order to locate the rocket during flight and descent. The camera module will be mounted on a collar that will allow it to rotate freely in a horizontal plane. A diagram of the camera system can be seen in Figure 8 below. A weight equal to or slightly greater than the mass of the camera will also be attached to the collar that will allow the camera to orient itself into the correct position upon landing to capture both the ground and sky. There will be a threaded rod attached radially to the rotating collar. A weight disc will be able to slide along this rod, and will be secured by two nuts. By moving the weight along the threaded rod, the moment arm of the weight will change, allowing for the weight to be finely tuned so that the camera is horizontal when the system is at rest. A diagram of this concept can be seen below in Figures 9 and 10. The camera will be oriented such that the photos taken during descent will be oriented correctly in order to meet the requirement of having the sky toward the top of the photo and the ground toward the bottom. Images taken once the vehicle is landed will need to be rotated digitally by 90 degrees in order to orient the sky toward the top of the photo and the ground toward the bottom. The initial code for rotating the images can be found in Appendix D. This design assumes the rocket body will land on its side. A polycarbonate ring integrated into the body will allow light to pass through to the camera.

Figure 8 ­ Camera subsystem

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Figure 9 ­ Weight concept

Figure 10 ­ Weight concept

The electrical schematic can be found in Appendix A. One aspect not displayed in Appendix A is the presence of an electrical slip ring in the camera system. This slip ring, which can be seen in Figure 8, shall be used to transmit power and signal to and from the camera while the system is rotating. 5.2 Payload Concept Features and Definition The RRS first entered USLI in 2011 with 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. As explained in the proposal, the RRS independently developed a new active drag control system and a new atmospheric measurement system. The active drag control system is a set of four drag flaps that will slow the rocket down according to a predetermined position, acceleration, and velocity profile. This drag flap system is creative in that it originated from a rocket fairing concept. This concept centered on body flaps ejecting partially from the side of the rocket. This concept was adapted to ensure that the flaps would stay connected to the rocket body. This would allow the flaps to manipulate the drag coefficient of the rocket in a more

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stable configuration. The main science objective of the drag flaps is to measure the suction effect caused by the flow moving around the flaps during their opening and closing. This effect has been documented in multiple scholarly publications on the computational fluid dynamics (CFD) analysis of aircraft air brakes. The idea for measuring the suction effect of the drag flaps at its most basic level came from the suction effect created when opening doors in climate controlled rooms. The challenge in this portion of the payload design lies in: the stress analysis of the drag flap system that ensures it is a safe and durable mechanism, the use of CFD analysis to measure fluid dynamics and the suction effects around these structural protuberances, and the use of CFD analysis to determine a drag coefficient curve to correctly manipulate the rocket’s movement towards the target apogee. The atmospheric measurement system idea came from the idea of digitally manipulating images of the horizon during the rocket descent and landing using code. This idea originated in learning to manipulate images using student­developed programming. This idea was then altered to only check if the image matches the target image orientation. The camera system provides mechanical challenges due to the number of degrees of motion needed for camera translation and rotation around the central axis of the rocket. The camera system provides large challenges for computer science students and electrical engineers in the RRS because of difficulties arising from data transmission. The atmospheric measurement system also includes several sensor arrays that provided challenges to the electrical engineers of the RRS, including access to UV light and barometric pressure at the outside of the rocket body for measurement inside the payload section. All in all, the payload section provides numerous mechanical, electrical, aeronautical, and computer science challenges that must be surmounted by the RRS. The previous RRS club and member experience in these arenas, as well as support from the Rensselaer community will provide a strong guide for success in Project Red Gemini. 5.3 Science Value

There are a number of objectives that were established for this project. Many of these objectives are related to data collection. Data will be collected during this launch from a camera, mounted within the rocket, aerodynamic sensors, located beneath the drag flaps, sensors for measuring humidity, temperature, barometric pressure, solar irradiance, UV light levels, a GPS module, a wireless transmitter capable of sending data to the ground team, and an accelerometer to determine payload orientation during flight and recovery. The camera is going to be mounted and oriented in such a fashion that it is able to include the horizon and the ground within the context of the same photo. This camera will be able to store its data within an SD card that is also mounted on the rocket for extraction after recovery. The aerodynamic sensors will be able to analyze the turbulent flow behind the drag flaps to determine the feasibility of the design for accurate adjustment of the apogee of a rocket mid­flight. It will also be used to collect data so that an on­board microprocessor can determine the drag coefficient of the protuberances.

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The other sensors will be set up to simultaneously record data on the SD card while transmitting the data to the ground team. The GPS module and the altimeter will be used to determine the height of the rocket. This, in turn, will be used to determine the control of the drag flaps to adjust the apogee. It will also send data to the ground team via the transmitter. The payload success criteria for the payload is that the payload can successfully transmit at least two images during the descent of the rocket, and at least three images after landing (where the images have the horizon in frame with the sky at the top of the frame and the ground in the bottom of the frame), that the payload can successfully gather the required sensor data (pressure, temperature, relative humidity, solar irradiance, and UV radiation) every 60 seconds after landing and transmit/store the data for up to 10 minutes after landing, and finally, that the effects of the structural protuberances (the drag flap system) can be accurately measured and transmitted back to the ground station. For most of the scientific payload, sensors will be passively collecting and transmitting data, so the logic, approach, and method of investigation are fairly straight­forward. The drag flap system however, involves actively measuring pressure in front of and underneath the flaps in order to analyze the suction effect. According to several early CFD models that have been tested, there should be an increase in pressure underneath the flaps that increased as the flaps are deployed. In particular, the Fluent software that was run since the Preliminary Design Review has shown that the greatest pressure differential can be found at the outermost radial edge of the flaps. In order to investigate this, barometers can be placed beneath the flaps to measure the increase in pressure and barometers placed in front of the drag flaps. A comparison can be made between the readings from both data sets in order to see if there is truly a suction effect being created by the drag flaps. In this test, the angle in which the drag flaps are deployed at is one of the largest variables, since according to the CFD tests, the larger the angle they are deployed at, the greater the suction force. For a control, the barometer in front of the drag flaps will be used, since it should not be affected by the structural protuberances. Since a suction effect will be accompanied by a change in pressure, the collecting the pressure data from beneath the drag flaps and then comparing it to pressure data taken in front of the drag flaps should show a difference in pressures, mainly that the pressure beneath the drag flaps is higher than the ambient pressure taken by the sensors in front of the drag flaps. The barometric sensors chosen have a resolution of 0.03hPa / 0.24m. This resolution allows for a fairly accurate reading of pressure in the different areas of the rocket. The temperature gathering sensor on the sensor also has a +/­ 2C accuracy, which might affect the resolution of data gathered. Sources of error include the accuracy of the sensors which could lead to inaccurate readings. The preliminary process procedure is as follows. Before launch, pressure reading will be taken from both sets of barometers and logged. After motor burn out when the drag flaps begin to deploy, measurements from both sets of barometers can be collected again and a difference between the two can then be

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calculated and then logged with a timestamp in order to compare pressure the effects of the angle of the drag flaps and the pressure differences.

6. Project Plan 6.1 Budget Plan

Income

Income in Account $1,530.00

Membership Dues $1,600.00

School of Engineering Funding $1,000.00

MANE Department Funding $500.00

Fundraising Goal $3,000.00

Total $7,630.00

Vehicle Design Team

Item Price Quantity Total

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

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

16.75" Length ­ Nosecone $21.95 1 $21.95

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

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

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

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

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

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

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

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

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

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Aerotech K1103X Motor $101.69 2 $203.38

Total $490.58

Recovery Design Team

Item Price Quantity Total

PerfectFlite StratoLogger CF $54.95 1 $54.95

Raven3 Altimeter $115.00 1 $115.00

Skyangle Cert­3 Drogue $27.50 1 $27.50

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

Ejection Charge $5.00 2 $10.00

Total $289.95

Payload Design Team

Item Price Quantity Total

Nuetrino Microprocessor $20.00 2 $40.00

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

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

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

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

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

MicroSD card breakout board+ $14.95 1 $14.95

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

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

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

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XBee Pro 900 RPSMA $54.95 1 $54.95

Analog UV Light Sensor Breakout ­ GUVA­S12SD $6.50 1 $6.50

SparkFun Humidity Sensor Breakout ­ HIH­4030 $16.95 1 $16.95

Miniature Solar Cell ­ BPW34 $1.50 2 $3.00

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

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

Total $336.59

Travel

Item Price Quantity Total

Hotel Rooms $122.00 15 $1,830.00

Airfare $405.00 8 $3,240.00

Total $5,070.00

Small Scale Launch

Item Price Quantity Total

Estes Pro Series II Leviathan kit $35.00 2 $70.00

Aerotech G80 motor $11.99 1 $11.99

Payload Bay hardware $10.00 1 $10.00

Ejection Charges $5.00 2 $10.00

Total $101.99

Income $7,630.00

Spending $6,289.11

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Difference $1,340.89

6.2 Funding Plan

For this project, the RRS fundraising goal is $3,000. At this point, the RRS has sufficient funding for the construction and testing of the rocket. This funding came from club membership dues, funding received from the RPI School of Engineering, and funding received from the RPI Mechanical, Aerospace, and Nuclear Engineering (MANE) Department. The remaining funding will go towards travel and lodging at the competition. The rest of the fundraising will be through corporate sponsorship and community outreach. The club has already reached out to corporate contacts and to various other aerospace companies, either by letter or email. The RRS sponsorship offer includes a request for funding and, in return, a certain amount of publicity either on the RRS website, the RRS rocket, or even during launch footage. The rewards increase in tiers according to the amount donated. It is the hope of the RRS that this method proves mutually beneficial, in helping to reach the RRS fundraising goal and providing advertisement for the companies that choose to get involved. The community outreach will be accomplished through a program that exists at RPI called WeR Gold. This program sponsors student­run projects by reaching out to alumni and friends of Rensselaer. The RRS was successful in applying for funding. The information the RRS has submitted to WeR Gold will now be used by the program organizers to create a donation site to encourage donations to this project. This information includes several pictures, a video, quotations by club members, and a project mission statement.

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6.3 Timeline

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6.4 Educational Engagement Plan and Status

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

6.4.2 Test Launch Rensselaer students and local high school students will be invited to see the test launch of the RRS rocket. They will be able to observe launch procedures and will learn about rocketry safety and its importance. After the launch RRS will hold a discussion on the science of rocketry. We will schedule this event around the test launch of the full­scale rocket.

6.4.3 STEM Engagement with Local Schools RRS has contacted several elementary schools in Troy, NY and spoken to administrators there about coordinating STEM events organized and lead by the RRS. The team has opportunities to hold programs with the Troy Boys and Girls Club and the local Civil Air Patrol. The RRS plans to hold events early in the spring semester in order to complete the educational engagement requirements by FRR. These STEM events may include an egg drop, model rocket build and launch, and other hands­on learning opportunities. The RRS is in the process of planning the events and scheduling dates.

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Appendix

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Appendix A: Scientific Payload Schematics

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Appendix B: Structural Design Assembly Drawings

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Appendix C: Code to Rotate Camera Images

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