2016-2017 NASA USLI Critical Design Review (CDR)

Rensselaer Rocket Society (RRS)

Rensselaer Polytechnic Institute

110 8th St

Troy, NY 12180

Project Name: Andromeda

Task 3.3: Roll Induction and Counter Roll

Friday, January 13th, 2017

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1. Table of Contents 1. Table of Contents

1.1 List of Figures 1.2 List of Tables

2. Executive Summary 2.1 Team Summary 2.2 Launch Vehicle Summary 2.3 Payload Summary

3. Changes Made Since PDR 3.1 Vehicle Changes 3.2 Payload Changes 3.3 Project Plan Changes

4. Vehicle Criteria 4.1 Design and Verification of Launch Vehicle

4.1.1 Mission Statement 4.1.2 Requirements and Mission Success Criteria 4.1.3 Selected Design Components 4.1.4 Vehicle Design 4.1.5 Risk 4.1.6 Design Integrity

4.1.6.1 Fin Design Integrity 4.1.6.2 Proper Material Selection 4.1.6.3 Motor Mounting and Retention 4.1.6.4 Mass Statment

4.2 Subscale Flight Results 4.2.1 Subscale Vehicle and Scaling Factors 4.2.2 Launch Conditions and Simulation 4.2.3 Subscale Flight Analysis 4.2.4 Full-scale Design impact

4.3 Recovery Subsystem 4.4 Mission Performance Predictions

4.4.1 Mission Performance Criteria 4.4.2 Vehicle Simulations 4.4.3 Kinetic Energy Analysis 4.4.4 Drift Calculations

4.5 Safety

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4.5.1 Launch Operations Procedures 4.5.2 Launch Vehicle Safety Analysis 4.5.3 Payload Safety Analysis 4.5.4 Environment Safety Analysis

5. Payload Criteria 5.1. Design of Payload Equipment

5.1.1 Mechanical Design 5.1.2 Electrical Design

5.2 Risk Assessment

6. Launch Operations Procedures

7. Project Plan 7.1 Testing 7.2 Requirement Verification 7.3 Team Derived Requirement Verification 7.4 Budget and Funding Plan 7.5 Timeline

Appendix

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1.1 List of Figures Figure 3-1: Increased Stability at Point of Rail Exit as Shown in OpenRocket Figure 3-2: Model of Assembled Payload Subsystem Figure 4-1: Final Vehicle Fin Dimensions Figure 4-2: Simulated Thrust Curve of Cesaroni L910 from Thrustcurve.org Figure 4-3: Model of Rocket Assembly Shown as Independent Sections Figure 4-4: Subscale Vehicle Model in OpenRocket Figure 4-5: Simulated Launch of Subscale Vehicle Figure 4-6: OpenRocket Simulation given drogue does not deploy Figure 4-7: Electrical schematics for altimeters Figure 4-8: Model for Flight Simulation (Launch-Ready Motor) Figure 4-9: Model for Flight Simulation (Top-of-Rail Motor) Figure 4-10: Model for Flight Simulation (After Motor Burnout) Figure 4-11: Flight Simulation from OpenRocket Figure 5-1: CAD Model of Assembled Payload Subsystem Figure 5-2: Linear Plate (left) and Cam Plate (right) from Top View Figure 5-3: Exploded View of Payload Subsystem Components Figure 5-4: Target Location of Payload Subsystem

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12 13 14 17 18 22 23 24 25 25 26 39 40 40 41

1.2 List of Tables Table 4-1: Vehicle Risk Table 4-2: Mass Statement for Launch Vehicle on the Pad Table 4-3: Landing Kinetic Energies of Independent Sections Table 4-4: Calculated Draft Distances for varied Wind Speeds Table 4-5: Material Risks to health and Risk mitigation Table 4-6: Color Coded Risk Assessment Codes Table 4-7: Different Levels of Risk and Their Acceptance Levels Table 4-8: Personnel Hazard Risk Assessment Table 4-9: Launch Vehicle Safety Risk Analysis Table 4-10: Payload Safety Risk Analysis Table 4-11: Environmental Safety Risk Analysis Table 4-12: Environmental Impact on Rocket Risk Analysis Table 5-1: Payload Phase Timeline Table 5-2: Electrical Components Table 5-3: Payload Risk Assessment Table 6-1: Launch Operations Procedures Table 7-1: Testing Plan Table 7-2: Requirement Verification Table 7-3: Team Derived Requirements Table 7-4: Vehicle Verification Table 7-5: Payload Verification Table 7-6: RRS Budget Summary

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2. Executive Summary 2.1 Team Summary The Rensselaer Rocket Society (RRS) is a student organization located at Rensselaer Polytechnic Institute (RPI). The RRS operates in the Ricketts Building at RPI. The RRS’s faculty advisor is Dr. Jason Hicken, Assistant Professor in the department of Mechanical, Aerospace, and Nuclear Engineering. The Community Mentor for the RRS is Jody Johnson (NAR Level 3 Certified, NAR #85182 SR, TRA Level 3 Certified, TRA #10973). The mailing address for the RRS is:

Rensselaer Rocket Society Department of Mechanical, Aerospace, and Nuclear Engineering

Rensselaer Polytechnic Institute 110 8th St

Troy, NY 12180 During this phase of the project, the RRS had some organizational changes. Victoria Castillo and Cassandra Castillo stepped down as the vehicle and recovery team leads, respectively. The new vehicle subsystem team lead will be Sean Beacham, a Junior Aeronautical Engineering student. The new recovery subsystem team will be Rebecca Caswell, also a Junior Aeronautical Engineering student.

2.2 Launch Vehicle Summary The launch vehicle will be approximately 102 in long, with a body diameter of approximately 6 in, and have three vehicle fins of custom dimensions. The launch vehicle will be propelled by a Cesaroni L910 75mm motor. The recovery system will consist of an electronic dual-deployment system that will deploy a drogue parachute at apogee and a main parachute at a much lower altitude during descent. The recovery system will be controlled by a set of two completely independent, redundant altimeters. The primary altimeter will be a PerfectFlite Stratologger SL100. The secondary altimeter will be a Featherweight Raven3. Each altimeter is powered independently, and is connected only to ejection charges for parachute deployment. 2.3 Payload Summary The payload design attempts to complete the requirements of the challenges outlined in section 3.3 of the 2017 NASA Student Launch Colleges, Universities, Non-Academic Handbook . To achieve this, the payload includes two cam systems that deploy two sets of opposing blades to induce roll and counter-roll.

The Payload Module utilizes redundant gyroscopes to monitor roll and dynamically deploy blades to match an idealized rotational model. The two sets of three blades are asymmetric airfoils (NASA NLF15) with fixed angles of attack of 5° and -5°, respectively. Two independent motors drive the cam, and are controlled by a central microprocessor, which deploys each set as they are needed. Following the roll-counter-roll phase, the

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microprocessor will command the motors to equally extend both sets of blades, to act as active drag control, slowing the rocket to reach the target altitude.

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3. Changes Made Since PDR 3.1 Vehicle Changes The changes made to the vehicle subsystems reflect feedback received from the PDR. Chiefly, the design team sought to increase the static stability margin of the launch vehicle at the point of rail exit. To achieve this, the upper airframe section was extended by 4 in and the ballast mass in the nose cone was increased from 16 oz to 24 oz. Additionally, the mass estimate in the model for the fin can reinforcement was reduced from 20 oz to 12 oz, which is a more realistic estimation. As a result of these changes, the static stability margin at point of rail exit has increased from 2.07 Cal to 2.24 Cal in OpenRocket simulations.

Figure 3-1: Increased Stability at Point of Rail Exit as Shown in OpenRocket

3.2 Payload Changes The majority of the changes to the payload subsystem are to the system integration. The payload can is the skeletal structure that supports the mechanical payload while also easing the integration to the vehicle subsystem. Attached to the skeletal structure will be 6 guides that will constrain the blades radially as they are deployed, seen below. Each guide will be made up of two aluminium strips that will be fastened to the straight plates.

Figure 3-2: Model of Assembled Payload Subsystem Screws will be able to fasten the can to the structure by utilizing the L-brackets that flare up and down from the skeletal structure shown above. In total, 12 screws will secure the payload system into the body tube. Further detail is given in 5.2.1.

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3.3 Project Plan Changes The project plan was again updated to more closely represent the remaining work for the project. At this phase, most design work has been completed. Acquisition of the majority of the vehicle subsystem components was slightly delayed during the ordering process. As such, some items on from the Gantt chart have been pushed back. This should not cause a notable delay in the construction process. Additionally, the project plan now includes all verification and testing plans, as specified in the NSL Handbook for the CDR. These sections were distributed throughout the Vehicle and Payload sections of the PDR.

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4. Vehicle Criteria 4.1 Design and Verification of Launch Vehicle

4.1.1 Mission Statement The launch vehicle will safely propel the payload to an apogee of at least one mile AGL, then safely return to ground via a dual-deploy recovery system such that the vehicle may be readily reused. 4.1.2 Requirements and Mission Success Criteria The rocket will reach an altitude of approximately 5,280 ft. This requirement will be successful if the vehicle reaches apogee between 5,000 ft and 5,600 ft. Altitude will be measured primarily by a barometric altimeter and reported via audible beeps post-flight. A secondary barometric and accelerometer based altimeter electronically separate from the primary altimeter will provide redundancy for the recovery subsystem. The launch vehicle will be reusable with fewer than four independent, separable sections. These requirements will be successful if the vehicle is able to be prepared and re-launched immediately after landing, and if the vehicle design utilizes less than four independent, separable sections. The launch vehicle will be powered by a single stage. This single stage will use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR) and Tripoli Rocketry Association (TRA). The motor will not exceed L-class and will be launched by a standard 12-volt direct current (DC) firing system. This requirement will be met by the design parameters of the rocket if it is powered by a single stage using a selected motor that is approved by the NAR and TRA, does not exceed L-class, and can be fired by a standard 12-volt DC ignition system. The recovery system of the launch vehicle will be electronic dual deploy with a drogue parachute deployed at apogee and a main parachute deployed at a much lower altitude during descent. The independent sections will be held together by removable shear pins, and each independent will have a kinetic energy of less than 75 ft-lbf at landing. The requirements will be met if the launch vehicle successfully deploys the drogue parachute at apogee and the main parachute much later. Shear pins holding the independent sections together should cleanly break during recovery deployment. The kinetic energy requirement will be met by selection of appropriate parachutes and verified by the calculation of kinetic energy at landing using each section’s estimated mass. The recovery system electrical circuits will consist of redundant altimeters that are physically and electronically separate from any payload electronics and power supply. Each altimeter will have a dedicated power supply and arming switch. These requirements will be met if there are redundant altimeters for the recovery system that are physically and electronically separate from all other

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electronics, and are powered independently of all other systems and armed by dedicated arming switches. 4.1.3 Selected Design Components The PDR outlined several design alternatives for the various components of the vehicle subsystem. This section presents the selected alternatives and the justifications for each selection.

4.1.3.1 Upper Airframe: The nose cone will be made of G10 fiberglass and will have a tangent ogive shape. There are a variety of benefits to selecting fiberglass over the alternative polypropylene plastic. Although the plastic should have been sufficiently strong, using the more rigid material will increase the factor of safety. A ballast will be required in the nose cone to increase stability. By using the heavier material for the nose cone, the weight of the ballast, which does not contribute to the structural strength of the rocket, is reduced. With respect to the ballast, plastic nosecones are generally manufactured with a closed base, restricting design choices for methods of ballasting. The most integral reason for selecting G10 fiberglass was due to the manufacturer of the body tubes for the RRS rocket not offering a 6-inch diameter plastic nose cone. Changing manufactures between connecting parts can cause flaws in the fit of the components due to the difference in tolerances and manufacturing processes between them. Since PML is the selected manufacturer, the RRS was limited to the shapes offered by PML. Of the shapes considered, the tangent ogive was the available design option that performed best in OpenRocket simulations.

The material for the body tubes was selected before writing the preliminary design review, but the selection process is briefly reiterated below. The RRS put forth four material options and five criteria to rate each material on in the preliminary design review. The phenolic resin outscored the alternatives by a significant margin. Thus, the RRS has chosen phenolic resin as the material for the body tubes. The most beneficial properties of the phenolic resin include its strength, cost, and weight. In the categories of weight and cost, the resin scored comparably with kraft cardboard, receiving nearly perfect marks. The main benefit of using resin was its compressive strength. The RRS has performed limited experimentation comparing the strength of the kraft board and the resin. The tests provided enough data to deem the resin as significantly stronger. The greatest drawback of this material is its workability. As the resin is a brittle material, the RRS will have to be cautious when completing the manufacturing of the rocket body. Another design choice

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regarding the body tubes was whether to use a constant or a changing diameter along the length of the rocket. A changing diameter would add multiple layers of complexity to the calculations and manufacturing processes required to finish the rocket. For this reason, the RRS decided to use a constant diameter. The chosen diameter was approximately 6-inches over the 4-inch alternative. The body will be comprised of three body tube sections to ease transportation. The only change in the body tube structure since the preliminary design review was an extension to the uppermost body tube in an effort to increase stability. Concerns were brought forth during the PDR presentation about the static stability margin. The rocket only exceeded the minimum stability metric by 3.5% with the previous design. With the change in length, in conjunction with an increase in nose cone ballast, the RRS’s rocket exceeds by 12%. This change can be clearly seen and will be noted in section 4.1.4. The slots required for the payload fins to be deployed create a large change in the maximum stress experienced by the body tubes in this section. The slots are rectangular in shape with filleted corners of 0.125” radius. After preliminary analysis, the estimated increase in stress was deemed significant and therefore the system will require further reinforcement. The RRS has decided to support the perimeter of the openings with fiberglass and epoxy. In addition, increasing the radius of the fillet at the corners decreases the maximum stress. A design change of this metric could further support the system. Before beginning work on the section housing the payload, the RRS intends to perform finite element analysis on this particular section of the vehicle to better understand the stress concentration in the region.

4.1.3.2 Lower Airframe The fins of the rocket will be made of G10 fiberglass. The other material mentioned in the preliminary design review was plywood. Plywood would have allowed for ease of customization as the RRS has access to a laser cutter, but due to the lack of material strength offered by plywood and the simplicity of the fin shape, fiberglass was deemed superior. The price of the fiberglass fins is also well within the budget of the RRS. Although the team decided early on in the process that the fins were to be trapezoidal, the final profile and thickness were chosen after analyzing the relationships between surface area, thickness, tip length, and drag. Specifically, the thickness of the fins was selected to reduce the risk of fluttering due to aerodynamic forces. A diagram of the profile is shown below in Figure 4-1

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Figure 4-1: Final Vehicle Fin Dimensions Figure 4-1 shows the final design of the vehicle fins for the launch vehicle. Note that the 1.515” tab length is interior to the vehicle body, and attaches the fins to the motor mount.

The design choices made for the motor mount assembly included whether to use reinforcing fillets connecting the mount to the fins and the material selection for both the mount itself and the centering rings. Reinforcing the fins to the motor mount assembly will greatly increase the integrity of the lower airframe with a minimal amount of weight increase. The RRS has decided to use reinforcements made of epoxy fillets and fiberglass strips set in West Systems epoxy. The material selected for the centering rings was plywood. Plywood is much less expensive than the alternative G10 fiberglass and allows for easy modification. In addition, the RRS does not have access to the necessary resources for machining fiberglass. The criteria for the material of motor mount are nearly identical to those of the body tubes. For this reason and the simplicity of ordering these parts from one manufacturer, the RRS decided to use the phenolic resin for the motor mount. The size of the motor mount will allow for the use of 75mm diameter motors.

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4.1.3.3 Motor The RRS has selected the Cesaroni 2856-L910-CS-0 Reloadable 75mm APCP motor (also simply referred to as the Cesaroni L910 motor) to power the launch vehicle. The simulated thrust curve is shown below in Figure 4-2.

Figure 4-2: Simulated Thrust Curve of Cesaroni L910 from

Thrustcurve.org As shown in the figure, the motor burns for 3.16 seconds and achieves a peak thrust of 1048 N. The average thrust is 906 N, with a total impulse of 2869 Ns (12% L class). The motor weighs 92.3 oz on the launch pad, and has an empty mass of 44.1 oz. Its total length is 13.8 in. These specifications have several desirable characteristics. The motor burnout time is relatively tame compared to short burn motors, while still allowing sufficient coast time for the payload subsystem to operate. The thrust profile is also fairly evenly distributed across the motor burn time, which allows for a more manageable maximum thrust value. One of the leading advantages of the Cesaroni L910 is its relatively low weight and short assembled length; both attributed to its 2-grain structure. The lighter motor increases stability, and allows for a shorter motor mount and lower airframe section. In addition to the specifications and characteristic above, the motor also propelled the launch vehicle to an acceptable simulated apogee before correction from the payload subsystem, as discussed in Section 4.4.

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4.1.4 Vehicle Design The final overall dimensions of the vehicle are a length of 102 inches, an inner body tube diameter of 6.007 inches, and an outer body tube diameter of 6.155 inches. The design choice involving the change in length from the PDR is outlined in section 4.1.3. An image of the vehicle separated into the lower airframe, upper airframe, and nose cone sections is shown below in Figure 4-3.

Figure 4-3: Model of Rocket Assembly Shown as Independent Sections

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4.1.5 Risk Table 4-1: Vehicle Risk

Risk Likelihood Impact Risk Reduction Plan

Failure of body tube sections

Medium Change of trajectory, potential mission failure, payload damage

Complete structural analysis, and ensure that weak points are reinforced

Failure of coupler Low Change of trajectory, potential mission failure, payload damage

Inspect components upon arrival and perform careful workmanship

Fin misalignment Medium Unexpected trajectory and apogee, potential mission failure

Use a fin alignment guide that is precision manufactured

Motor mount failure during burn

Low Potential damage to all components, unexpected trajectory, potential mission failure

Inspect mount and centering rings upon arrival and make strong bonds

4.1.6 Design Integrity

4.1.6.1 Fin Design Integrity The fin design selected was influenced by two primary goals: reduce drag to increase projected apogee, and withstand the forces expected during motor burnout. To satisfy the first goal, several fin shapes were iteratively tested to find what produced the highest projected apogee, all else held constant. From this process, the design team identified that a combination of a slightly swept back shape, as well as a relatively small tip chord produced optimal results. It was also discovered that while decreasing the tip chord generally improved results, bringing it to a sharp point negatively affected simulation results compared to a small, non-zero length tip chord. As discussed above, G10 fiberglass was selected for its rigidity and strength. From club experience, fins of 0.125” thickness provide ample structural integrity without increasing drag substantially. This thickness should also ensure that no fluttering phenomenon occur during flight. 4.1.6.2 Proper Material Selection Proper material selection is discussed throughout sections 4.1.3 and 4.3.2. 4.1.6.3 Motor Mounting and Retention The Cesaroni L910 motor will be mounted in the launch vehicle using a standard configuration. A 3 in diameter, 15 in long piece of phenolic resin tubing from PML will serve as the motor mount, mounted in the launch vehicle by two ½ in thick birch plywood centering rings of appropriate

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inner and outer diameters. The centering rings will be bound by high strength epoxy to the motor mount, as level as possible. Using a fin alignment guide, the fins will be attached to form the fin canister, which will then be reinforced along the edges between the fins and the motor mount. The fin canister will then be bonded at the centering rings to the vehicle airframe using high strength epoxy. The motor will be retained using a Aeropack non-flanged aluminum 75mm motor retainer. The retainer will be secured to the motor mount by J.B. Weld high temperature epoxy, taking care to do so in a level fashion, with no excess epoxy on any part of the motor retainer. 4.1.6.4 Mass Statment Table 4-2 below shows an update mass statement for Project Andromeda. There were several changes compared to the mass statement included in the PDR. The ballast mass was increased for reasons discussed above and in section 4.4.2. The mass estimate for the payload and fin canister reinforcements has decreased, while unaccounted for parts such as bulk plates and shock cord are now included. Over all, the mass of the launch vehicle on the pad increased from 386 oz to 408 oz.

Table 4-2: Mass Statement for Launch Vehicle on the Pad Part Mass (oz) Quantity Subtotal (oz)

Nose Cone 28.2 1 28.2 Ballast 24 1 24

Airframe Tubing (All sections) 60 1 60 Coupler 5.4 2 10.8

Bulkheads 3.5 3 10.5 Motor Mount 11.5 1 11.5

Motor Retainer 4 1 4 Centering Rings 2.06 2 4.12

Fins (x3) 22.7 1 22.7 Fin Can Reinforcements 12 1 12

Estimated Payload 72 1 72 Main Parachute 34 1 34

Drogue Parachute 6 1 6 Shock Cord (250”) 8.02 2 16.04

Motor 92.3 1 92.3 Total - - 408.16

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4.2 Subscale Flight Results 4.2.1 Subscale Vehicle and Scaling Factors In conforming with task 1.16 of the competition requirements, the RRS has modeled, created, launched, and recovered a subscale version of the launch vehicle to be used in for the Andromeda project. To the extent possible, the design team sought to create a 1:2 scale vehicle of the full-scale rocket. This scaling factor was especially enforced with respect to the subscale vehicle’s overall length, body diameter, and fin shape. The RRS designed and created the subscale rocket with the intention of modifying an Estes Pro Series II mid-power rocket kit. In addition to their reasonable price point, these particular kits are used by new members at the beginning of the school year as an introduction activity. As such, the club has a fair inventory of compatible parts that could be used for modification. The Estes Pro Series II Scion kit was used as a starting point for the subscale rocket. Figure 4-4 below illustrates the final model of the subscale vehicle. The vehicle body for this kit has a 3 in diameter, which fits the 1:2 scaling. Normally, these kits rely on motor ejection charges to operate single deployment recovery systems. As such, the Scion kit only has one separation point at the nose cone for parachute deployment. In order to modify the kit to use a dual deploy recovery system, an additional airframe section was added and conjoined by a 3D printed coupler section. The coupler section was enclosed by bulkheads and switch band, and also acted as the avionics bay for the vehicle. It was permanently fixed to the upper body section via epoxy, and friction fitted to the bottom airframe to serve as an additional separation point. With the additional body section and original nose cone, the overall length of the subscale vehicle was about 49 in. This is just 2 in less than half of the full-scale’s length of 102 in. Another advantage of using the Scion kit was that the fins required only a small modification to match a scale profile of the design for the full-scale vehicle. All dimensions of the subscale fins matched the 1:2 scaling ratio with the exception of the tip chord, which were kept at 1.5 in long due to manufacturing constraints.

Figure 4-4: Subscale Vehicle Model in OpenRocket

The subscale flight utilized a single Featherweight Raven 3 altimeter, secured on a payload sled within the avionics bay. This altimeter was programed to ignite the initial ejection charges for the lower separation point at apogee. The second

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ejection charges were programed to eject the nose cone and deploy the parachute at 550 ft AGL during descent. The RRS used an Aerotech G80-10 motor to power the subscale test launch. The motor had its fused ejection charge completely trimmed off to facilitate the use of the dual deploy recovery system. This motor was not chosen according to any scaling factor relative to the selected Cesaroni L910 motor selected for the full-scale launches. Rather the Aerotech G80-10 was chosen for being the largest motor available without needing certification, thus making it far easier to obtain. The power of the motor still allowed for sufficient altitude to test a dual deploy recovery system.

4.2.2 Launch Conditions and Simulation Ground wind speeds at launch were around 5 mph. The launch rail was set without angle into the wind as the low wind speed did not cause drift to be a concern. The skies were very clear providing great visibility throughout the entire flight of the rocket. The temperature at the launch site was 16 degrees Fahrenheit. Figure 4-5 below shows a simulation in OpenRocket using the model and launch day conditions from above.

Figure 4-5: Simulated Launch of Subscale Vehicle

In this simulation, the projected apogee for the subscale test launch is 1220 ft AGL with a maximum acceleration of 292 ft/s2 and a total flight time of 36.8 seconds.

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4.2.3 Subscale Flight Analysis All aspects of the subscale flight went as expected. The rocket was propelled straight off the pad and cocked slightly into the wind. After reaching a max altitude of 1290 ft AGL as reported by the on board Featherweight Raven 3 altimeter, the dual deploy recovery system separated the rocket at the lower separation point, falling separated without drogue until the dual deploy system successfully separated the rocket at the nose to deploy the main at 550 ft AGL. The rest of the landing was unremarkable.

4.2.4 Full-scale Design Impact The nominal flight of the subscale indicates a flight worthy design for the full-scale rocket, as the rocket was a direct scale model. As such, no major design changes will occur to the full-scale rocket as a result of the subscale flight. One notable difference between the subscale and the full-scale rockets is the lack of ballast in the nose cone in the subscale. Adding a scaled amount of ballast to the nose cone of the subscale would overstablize the rocket and require a larger motor to match the thrust profile. As such, the ballast was left out of the subscale.

4.3 Recovery Subsystem

4.3.1 System Level Design The final components for the recovery subsystem were selected from the alternates - these include parachutes, altimeters, and attachment hardware. The SkyAngle Classic II 60 was chosen for the main parachute. The SkyAngle Classic II was chosen over the SkyAngle Classic mainly due to the difference in material. The Classic II is made from zero-porosity silicone-coated balloon cloth which means a lower descent rate, better stability, increased toughness, and a lower opening shock compared to a higher porosity material. For the drogue chute, the Ballistic Mach II drogue was chosen for its strength and flight characteristics. The two foot diameter chute was chosen to keep drift within acceptable limits. The altimeters chosen for the final product were the Featherweight Raven 3 and the PerfectFlite Stratologger SL100. The deciding factor for the altimeters was the post flight software. The Raven 3 comes with post flight software which makes up for the high cost. The backup altimeter, the Stratologger, has post flight software that must be purchased separately but the main benefit is an adjustable deployment altitude. The Raven 3 is the best option due to its post flight software, however, the Stratologger is a suitable back up. A 1.5’’ forged stainless steel 316, ¼”-20 eye bolt with a shank length of two inches along with ½ ‘’ tubular nylon shock cords were chosen for the attachment hardware. A closed eye bolt can withstand higher loads than an open eye bolt and this is due to shear flow. A forged eye bolt was chosen again for its load-bearing abilities - the eye bolt is constructed as a single piece rather than

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being constructed of welded parts which could lead to points of failure. The eyebolt will have a shoulder to counter the effect of forces in non-axial directions. Stainless steel was chosen for its high ultimate strength and resistance to corrosion. According to the manufacturer, the eye bolt has a workload limit of 500 lbs. The nylon shock cord was chosen over kevlar for a few reasons. Though Kevlar has more resistance to the heat and corrosion, it has a much lower ultimate strength compared to tubular nylon.The tubular nylon is also more elastic making it more effective for a large rocket.

4.3.2 System Components

4.3.2.1 Parachutes For the dual deployment system, the two parachutes are the SkyAngle Classic II 60 and the Ballistic Mach II. The main parachute is a SkyAngle Classic II 60 with a surface area of 39.3 square feet and a weight of 18.2 oz. It has zero-porosity silicone-coated balloon cloth allowing for a lower descent rate, more stability, and increased toughness. The line attachment consists of a 12/0 nickel-plated swivel joint that can withstand 1500 lb. From OpenRocket simulations, it is found the maximum acceleration on the rocket to be 428 ft/s2 at an altitude of 600 ft just after deployment resulting in a max force of 325 lbf. Thus, there is a factor of safety of 4.6 for the line attachments. SkyAngle parachutes use a non-circular surface area, which does not directly translate to a drag coefficient typically used in parachute calculations. The effective surface area is the surface area of the parachute if the non-traditional shape is translated to a traditional surface area calculation. The main parachute has an effective surface area of 15.8 square feet.The manufacturer also lists loads to which the parachutes have been subjected without failure. Based on the manufacturer's recommendations, the selected parachute size is acceptable for rockets in the weight range of our rocket. A 2 foot diameter Ballistic Mach II will be used for the drogue. It is made from ballistic-grade rip-stop nylon. OpenRocket simulations calculate a max force of 24 lbf on the drogue and associated hardware at apogee. Given positive past experience with this type of parachute as well as given the small force experienced at apogee relative to the force experienced by the main chute at its deployment, it is reasonable to assume the components can withstand these forces.

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4.3.2.2 Harnesses, Attachment Hardware, Bulk heads The shock cords that connect the airframe with the parachutes will be ½” tubular nylon. A ½” thick birch plywood bulkhead with the 1.5’’ eyebolt will be used for the main chute. For the drogue, there will be eyebolt attachments on the forwardmost motor centering ring and on the rear of the payload section. Both the fore and aft shock cords will be 250’’ long or about 3 body lengths to ensure adequate separation of vehicle components during descent. The eye bolts are rated for 500 lbs and the tubular nylon is rated for 1200 lbs, given the max force estimate of 325 lbf, the factor of safety for each component is within a safe region. A stress analysis was done on the eyebolt and bulkhead attachment for two conditions: 1) both parachutes deploy 2) the drogue chute does not deploy. Calculations for the first condition are previously given in section 4.3.2.1 with a max force on the main of 325 lbf and a max force on the drogue of 24 lbf. For the second condition, the max force on the main parachute will be 171 lbf at about 700 ft when it is deployed. The OpenRocket simulation for the second condition can be found in Figure 4-7. Focusing on the forward bulkhead, the vertical force of 325 lbf will be transferred through the eyebolt. Thus, the eyebolt will experience a tensile stress. Given the cross sectional area of the eyebolt shank being 0.05 square inches, the stress is calculated to be 6500 psi. Stainless steel has an ultimate strength of 84 ksi resulting in a factor of safety of 12.9. For the second condition, with a force of 171 lbf, the factor of safety would be even greater. There will also be a bearing stress on the bulkhead as the force on the bolt is transferred to the bulkhead. The bearing stress is calculated from the projected area of contact (the diameter and thickness of the area of contact between the bulkhead and the bolt) calculated as 0.125 square inches, giving a stress of 2,600 psi. Given the difficulty of finding an exact value for the ultimate strength of plywood, a typical range of strengths is between 4,500 to 6,000 psi. This returns a minimum factor of safety of 1.7 which is the recommended NASA value for parachutes. Again, the second condition would result in an increased factor of safety than the first condition. The second eyebolt bulkhead attachment is on the forwardmost centering ring of the motor mount. When the drogue deploys at apogee, the max

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force experienced will be 24 lbf. The load is within the load bearing capabilities of the eyebolt and considering the force is much less than the previous bulkhead, it is assumed the stresses are within the factor of safety.

Figure 4-6: OpenRocket Simulation given drogue does not deploy

4.3.2.2 Ejection The ejection charges for the parachute deployment require one gram of black powder charge per six inches of interior body length containing the parachute, according to estimates made by Vern Knowles of Vern’s Rocketry. This leads to an approximate calculated ejection charge size of 1.7 g of black powder for main parachute deployment and 1.7 g of black powder for drogue parachute deployment. The black powder charges will be mounted in blast cups of the fore and aft bulkheads of the main airframe. Upon detonation, these charges will increase the pressure in their respective body sections by approximately 12 psi. Assuming the black powder gasses only do work on the bulkheads, which have an area of 28.5 square inches, this will create an ejection force of 340 lbf. The ejection force will be supported by six 2-56 nylon screws, each having a minimum shear force of 31 lbf. This creates a factor of safety of 1.8 for each ejection event. The separation of the rocket sections allows for the unfurling of the parachutes. The drogue will be deployed at apogee when the first pair of charges are detonated, and the main chute will be deployed at 700 feet when the second charge pair triggers. Kevlar parachute protectors will insulate the parachutes from the heat effects of the charge firings.

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4.3.2.4 Electronics The electronic components of the recovery system consists of a PerfectFlite Stratologger SL100 altimeter and a Raven 3 altimeter. The schematics for each altimeter can be seen in Figure 4-7 below. The design of the electronics allows for externally accessible switches on the exterior of the payload section to arm and disarm the charges and altimeters. GPS will be used for the tracking system with an Xbee wireless transmitter with a frequency of 2.4 GHz.

Figure 4-7: Electrical schematics for altimeters

The Raven 3 will be mounted vertically to ensure proper functionality. It uses pressure readings as well as an on board accelerometer to determine the altitude. Flight data is recorded with a sample rate of 20 Hz and can be downloaded post flight using a USB connection. The Raven 3 has 4 output channels, two of which will be used for the main and drogue. It has a max altitude of 100,000 feet and dimensions of 0.80” wide, 1.80’’ long, and 0.55’’ thick. The altitude accuracy is within +/- 0.3%. The Stratologger uses pressure sensors to determine the altitude. Flight data is recorded at a frequency of 20 Hz and is stored in the altimeter’s onboard internal memory. It has two output channels for the main and the drogue. The max altitude is 100,000 feet and has dimensions of 0.90” wide, 2.75’’ long, and 0.50’’ thick. The altitude accuracy is within +/- 0.1% and has an adjustable main deployment altitude within a minimum of one foot increments. The two altimeters provide redundancy, as the flight system is capable of functioning on either should one fail mid-flight. At the appropriate altitudes, each altimeter will send signals to trigger its set of charges for each ejection event. Thus, each altimeter can trigger charges for each separation event - the charges themselves are redundant. If one fails, parachute deployment will not be adversely affected.

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4.4 Mission Performance Predictions 4.4.1 Mission Performance Criteria The vehicle subsystem has several performance criteria that must be met on launch day. Primarily, the launch vehicle must obtain an apogee as close to 5280 ft AGL as possible. OpenRocket, the launch simulation software that the RRS utilizes, does not take into account the effects the payload system will have on launch apogee. These effects include increased drag during the roll induction and counter-roll induction phases of the payload, as well as the intentional coast control during the last phase of the payload deployment. Therefore, the RRS has created simulations in OpenRocket that slightly overshoot the target apogee. Additionally, the launch vehicle must land within a half mile of the launch site, and each independent section must land with no more than 75 ft-lbs of kinetic energy. To meet these requirements, the recovery system is tailored to minimize lateral drift in various wind conditions, while slowing the descent of the rocket such that no independent section lands with excessive kinetic energy.

4.4.2 Vehicle Simulations Figure 4-2 in Section 4.1.3.3 shows the simulated motor thrust curve for the selected Cesaroni L910 Motor. Using this motor configuration in OpenRocket, a simulated model of the rocket was produced as shown in Figure 4-8.

Figure 4-8: Model for Flight Simulation (Launch-Ready Motor)

On this model, the launch vehicle with a loaded motor has a stability margin of 2.04 Cal The simulated center of gravity is approximately 64.9 in down from the top of the nose cone, and the simulated center of pressure is approximately 77.4 in from the top of the nose cone. The design team once again used this simulation model to verify requirements 1.14 and 1.15 in the statement of work by analysing simulation conditions at the point of rail exit. The simulated velocity at rail exit is 79.2 ft/s, which complies with the minimum exit velocity of 52 ft/s. The same MatLab tool that was employed in the PDR was again used to determine the mass of the motor at the top of the launch rail. A mass component of equal size and mass to the motor at the top of the launch rail was inserted into the simulation model, as shown in Figure 4-9. The static stability margin at the top of the rail is 2.24 Cal, which complies with the minimum stability margin of 2 Cal. This stability margin was increased the 2.07 Cal reported in the PDR, in response to feedback the RRS received during the PDR presentation. The design team also updated the simulation model in a similar manner to include a dummy mass equivalent to a burnt-out L910 motor.

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Figure 4-9: Model for Flight Simulation (Top-of-Rail Motor)

As shown in Figure 4-10, the post-burnout center of gravity is approximately 60.6 in down from the top of the nose cone. With this center of gravity, the static stability margin is 2.74 Cal during the coast phase.

Figure 4-10: Model for Flight Simulation (After Motor Burnout)

The RRS recognizes that the process of inducing a roll-axis rotation may affect the the center of pressure of the launch vehicle, and thus alter the stability of the rocket. CFD analysis is currently on going to explore the effects of roll induction, but unfortunately are not completed for this report.

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Using this simulated model with the selected motor, a flight simulation shown in Figure 4-11 was created.

Figure 4-11: Flight Simulation from OpenRocket

In this simulation, the uncorrected apogee reached is 5363 ft, with a maximum acceleration of 284 ft/s2 and a total flight time of 95.5 seconds.

4.4.3 Kinetic Energy Analysis The final velocity was obtained from OpenRocket Simulations and was approximately:

v final = 23.9 ft/s This value was then used in conjunction with the mass values from Table 4-3 for all subsequent kinetic energy calculations.

Table 4-3: Landing Kinetic Energies of Independent Sections

Independent Section Kinetic Energy

Nose Cone - Nose Cone, Ballast = 52.2 oz

28.96 ft⋅lbf

Upper Airframe - 48’’ Phenolic Tubing, 2 Couplers, 2 Bulkheads, Payload = 126.72 oz

70.31 ft⋅lbf

Lower Airframe - 30’’ Phenolic Tubing, Motor Retainer, Motor Mount, Centering Rings, Fins, Fins Reinforcement, Motor Casing = 121.5 oz

67.41 ft⋅lbf

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The kinetic energy of each independent section does not exceed the maximum allowed kinetic energy of 75 ft-lbf.

4.4.4 Drift Calculations The calculated drift distances based on varying wind speeds can be found in Table 4-4 below. These values were calculated for a zero degree launch angle and assume that the vehicle drifts at the same velocity as the wind during a descent time of:

t descent = 78.9 s This descent value was obtained from OpenRocket simulations.

Table 4-4: Calculated Draft Distances for varied Wind Speeds

Wind Speed (mph) Drift Distance (feet)

0 0

5 579

10 1157

15 1736

20 2314

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4.5 Safety 4.5.1 Safety Officer The RRS has identified Philip Hoddinott as the acting safety officer. His responsibilities include ensuring shop safety and hazardous material procedures, which is partly accomplished through safety quizzes administered by the RPI School of Engineering. He will oversee the safe construction and launch of the pertinent rocket vehicles through supervision and inspections. He will monitor or designate a monitor for all RRS lab meetings.

4.5.2 Launch Operations Procedures The launch operations procedures are detailed in section 6. 4.5.3 Personal Hazard Analysis The RPI Rocket Team has conducted a thorough evaluation of all possible hazards that may affect any rocket personnel. The safety officer Philip Hoddinott is in charge of ensuring the safety and security of all rocket team activities. He has the right to allow or prohibit team members from working on projects. He is responsible for making sure all team members are properly trained on safety when it comes to handling chemicals, using machine tools, and fabricating the rocket.

All team members are required to take RPI’s lab safety course. This enssuers they know how to safely handle machine tools, materials, and chemicals. The following table summarizes hazardous materials, chemicals, and machine tools team members may encounter.

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Table 4-5: Material Risks to health and Risk mitigation

Material Risk to personnel health

Mitigation Required Safety Equipment

Emergency Equipment

Black Powder Skin, eye, and respiratory irritation. Fire and explosive risk.

Keep away from ignition source. Have fire extinguisher on hand. Work in ventilated, spacious area. Wear eye protection and protective gloves.

Face masks, gloves, glasses Fire extinguisher and fire blanket.

First aid kit, eye flushing station Fire extinguisher

Hydrogen Peroxide Fire risk Skin, eye, and respiratory irritation

Work in ventilated, spacious area. Wear eye protection and protective gloves.

Face masks, gloves, glasses

First aid kit, eye flushing station

Acetone Skin, eye, and respiratory irritation

Work in ventilated, spacious area. Wear eye protection and protective gloves.

Face masks, gloves, glasses

First aid kit, eye flushing station

West System 105 Epoxy

Skin, eye, and respiratory irritation

Work in ventilated, spacious area. Wear eye protection and protective gloves.

Face masks, gloves, glasses

First aid kit, eye flushing station

JB Weld Skin, eye, and respiratory irritation

Work in ventilated, spacious area. Wear eye protection and protective gloves.

Face masks, gloves, glasses

First aid kit, eye flushing station

Scotch weld Skin, eye, and respiratory irritation

Work in ventilated, spacious area. Wear eye protection and protective gloves.

Face masks, gloves, glasses

First aid kit, eye flushing station

Loctite Super Glue Skin, eye, and respiratory irritation

Work in ventilated, spacious area. Wear eye protection and protective gloves.

Face masks, gloves, glasses

First aid kit, eye flushing station

Cesaroni Pro75 Profire Igniter

Fire / burn risk Will be handled on site, outside with operating wearing face and body protection. Fire extinguisher will be on hand.

Gloves and Glasses.

Fire extinguisher on hand. First aid kit.

Cesaroni Pro75 Profire Motor

Fire / burn risk Will be handled on site, outside with operating wearing face and body protection. Fire extinguisher will be on hand.

Gloves and Glasses.

Fire extinguisher on hand. First aid kit.

Solder Skin burns and respiratory irritation

Work in ventilated, spacious area. Wear eye protection and protective gloves.

Face masks, gloves, glasses

First aid kit

Sanding Skin abrasions and respiratory irritation

Work in ventilated, spacious area. Wear eye protection and protective gloves.

Face masks, gloves, glasses

First aid kit

G10 Fiberglass board Skin, eye, and respiratory irritation

Work in ventilated, spacious area. Wear eye protection and protective gloves.

Face masks, gloves, glasses

First aid kit, eye flushing station

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The following tables explain the Risk Assessment Codes (RACs) used to evaluate the potential hazards in the NASA USLI launch. RACs are presented for the initial hazard, as well as for the hazard remaining after controls and mitigations have been applied. Table 4-6 identifies the color-coded RACs, which will be referred to later when assessing individual risks. Table 4-7 outlines the different levels of risk and their acceptance levels. In all cases, individuals involved in each task will be advised of the risks involved and proper safety precautions to be taken

Table 4-6: Color Coded Risk Assessment Codes

Probability

Severity

1 Catastrophic 2 Critical 3 Marginal 4 Negligible

A – Frequent 1A 2A 3A 4A

B – Probable 1B 2B 3B 4B

C – Occasional 1C 2C 3C 4C

D – Remote 1D 2D 3D 4D

E - Improbable E 2E 3E 4E

Table 4-7: Different Levels of Risk and Their Acceptance Levels

Severity-Probability Acceptance Level/Approving Authority

High Risk Unacceptable. Documented approval from the MSFC EMC or an equivalent level independent management committee.

Medium Risk Undesirable. Documented approval from the facility/operation owner’s Department/Laboratory/Office Manager or designee(s) or an equivalent level management committee.

Low Risk Acceptable. Documented approval from the supervisor directly responsible for operating the facility or performing the operation.

Minimal Risk Acceptable. Documented approval not required, but an informal review by the supervisor directly responsible for operation the facility or performing the operation is highly recommended. Use of a generic JHA posted on the SHE Web page is recommended, if a generic JHA has been developed.

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The following table shows the personal hazard risks associated with the project, their risk ratings, and their mitigation strategies.

Table 4-8: Personnel Hazard Risk Assessment Risk Cause Overall Effect Risk

Rating Mitigation Strategy Post -

Mitigation Risk Rating

Irritation of skin. Exposure to chemicals (epoxy, black powder, super glue).

Team member may suffer, delayed timeline.

3C Protective gloves will be worn while handling chemicals. Team members to be instructed on how to remove gloves without contact to skin.

3E

Irritation of eyes.

Chemicals / particulate / fast moving debris come in contact with eyes.

Serious risk to vision and health.

1C Safety glasses will be worn during all activities involving chemicals and machinery.

1E

Irritation of lungs.

Chemicals / particulate inhaled.

Serious risk to lungs and health.

1B Masks will be worn during all activities involving chemicals. Chemicals will only be handled in well ventilated areas. Soldering will be done in well ventilated areas.

1E

Chemical / Heat Burn.

Skin contact with soldering iron. Chemical contact with skin.

Serious risk of burns and/or scarring.

2C All personnel soldering must receive training. Gloves are to be worn when using chemicals. All personnel will be made aware of first-aid kits and proper burn treatment.

2E

Black powder ignition near team member.

Static discharge. Premature voltage discharge.

Skin burns. Hearing damage.

1C All personnel working with black powder will be grounded. Thick gloves will be worn.

2E

Severe static shock.

Static buildup on equipment or team member.

Damage to rocket electrical components. Premature black powder ignition.

2C Team members handling electronic equipment will be grounded.

2E

Limb caught in machine.

Untied long or loose hair. Loose fitting clothing

Serious risk of permanent injury or death.

1B Team members working with machines will tie long hair back and wear tight fitting clothing.

1E

Fire in Lab. Fire can be caused by a wide range of accidents, including electrical and chemical sources.

Serious risk of permanent injury or death.

1C Maintain properly cleared exits. No smoking or any other activity that uses a lighter or open flame. Proper usage of tools and chemicals.

1D

Tripping hazard in lab.

Persons in lab may trip over objects.

Unclean lab space could injure persons attempting to leave lab.

3B Maintain properly cleared exits. Keep Lab space trash free.

3D

Safety Officer Unavailable.

No identifiable person is in charge of ensuring that adequate safety operations are followed.

Possible increased likelihood of other risks.

2B Numerous students involved in the club are Level 1 certified by TRA and NAR. All participants briefed on safety operations.

2D

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4.5.4 Launch Vehicle Safety Analysis The following table shows launch vehicle safety risks, possible causes, risk ratings, and mitigation strategies.

Table 4-9: Launch Vehicle Safety Risk Analysis

Risk Cause Overall Effect Risk Rating

Mitigation Strategy Post Mitigation Rating

Parachute deploys too early or too late.

Altimeter malfunction. Improperly installed or damaged shear pins.

Rocket does not land with desired kinetic energy. Unexpected stress may damage recovery hardware or vehicle body.

2B Redundant altimeters. Rigorous ground testing of altimeters and other flight electronics.

2E

Parachute does not deploy.

Altimeter malfunction. Improperly installed shear pins.

Rocket does not land with desired kinetic energy. Damage to vehicle body. Endangerment of personnel near the launch site.

2D Redundant altimeters. Test altimeters. Run OpenRocket simulations to ensure rocket lands safely even if one parachute does not deploy.

2E

Attachment hardware failure (eyebolts, bulkheads, shock cords)

Large, unexpected stresses greater than load limit. Improperly installed hardware.

Rocket does not land within kinetic energy limit. Damage to vehicle body. Endangerment of personnel near the launch site .

2D Ensure all components can handle stresses (stress analysis). Proper manufacturing and assembly processes followed.

2E

Cracks form around blade holes.

Stress concentration around blade holes causes damage to outer material.

Crack formation limits number of possible flights. Potential for catastrophic failure.

2C Holes designed to have minimal street concentration. Fiberglass and epoxy reinforcement along hole perimeters

2E

Centering ring failure. Extreme stresses causes centering ring failure. Centering rings are not installed properly.

Displacement or dislocation of rocket motor during flight.

1C Ensure that epoxy is rated for maximum expected stresses. Have experienced team members verify that build is proceeding correctly.

1E

Bulkhead failure. Extreme or unexpected stresses.

Loss of eyebolts, payload, or motor. Possible damage to payload.

2C Rigorous ground testing of bulkhead materials to ensure strength. Analytical verification before building.

2E

Premature independent section separation.

Altimeter malfunction.

Extreme trajectory change Near certain mission failure

2D Ground testing of separation points with ejection charges. Rigorous testing of recovery electronics

2E

Shock cord breaks Extreme stresses cause shock cord to snap. Shock cord separates from eyebolt.

Independent sections of rocket may exceed kinetic energy calculations, be destroyed or damaged at landing, or lost on site.

1C Ensure that shock cord is rated for maximum expected stresses. Verify that eyebolt is correctly installed.

1E

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Risk Cause Overall Effect Risk Rating

Mitigation Strategy Post Mitigation Rating

Fins become unaligned or shear off during flight

Poor connection to motor mount assembly. Damage upon arrival or during vehicle construction.

Unexpected trajectory. Danger to persons near the launch site. Potential mission failure.

1D Careful inspection of the fins before and after mounting and between each flight

1E

Motor mount assembly fails to hold motor during burn

Centering rings have imperfect fit Centering rings have poor connection to mount or inner wall of body tube

Unexpected trajectory Further failures within the rocket Hazardous falling debris

1D Careful inspection of mount and centering before attachment and between flights Examination of connections pre-flight

1E

Premature motor ignition

The motor is exposed to open flame or an electrical arc

High risk of injury to local persons Potential damage to the launch vehicle

2D Store the motor safely and in a closed case throughout transport

2E

Fractures in body tubes

Careless transport Harshly working the material

High risk of launch day failure if unnoticed

1D Transport the vehicle in sections Slow speed cutting and sanding

1E

Fractures in a coupler Harsh workmanship Poor handling

High risk of launch failure 2D Careful planning of work done on couplers

2E

Fatigue failure of any and all components

Improper loading analysis

Limits flight repetitions 2C Ensure that components can withstand high cycle counts under expected stress amplitudes

2E

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4.5.5 Payload Safety Analysis Table 4-10 shows payload safety risks, possible causes, risk ratings, and mitigation strategies.

Table 4-10: Payload Safety Risk Analysis

Risk Cause Overall Effect Risk Rating

Mitigation Strategy Post Mitigation Rating

Blades deploy too early. Blades deploy too quickly

Altimeter malfunction. Controller error. Software error.

Roll and target altitude may not be achieved. Cam system may be damaged or inoperable.

2C Vigorously test altimeters and controllers. Use redundant components. Ensure Cam system can withstand stresses.

2E

Blades deploy too late Altimeter malfunction. Controller error. Software error.

Roll and target altitude may not be achieved

2C Ground testing using sample input to observe resultant deployment. Integration ground testing.

2E

Blades deploy asymmetrically.

Cam system failure. Trajectory altered, stress may cause structural failure.

1D Ground testing and integration testing. Ensure proper manufacturing of Cam plate.

1E

Blade structural failure.

Unexpected stresses. Improper handling or assembly of payload.

Trajectory altered, stress may cause structural failure. Destruction of payload.

1C Buckling, bending, and torsion load calculations predict success. Proper storage, handling, and assembly of payload procedures followed.

1E

Payload instrumentation failure.

Erroneous data fed into microprocessor. Failure of payload.

Flight goals not met. 2C Redundancies in instrumentation and error checks in code.

3D

Algorithm / Misc software failure.

Error in code. Failure to account for every scenario.

Flight goals not met.

2C Error checks in code. Rigorous ground testing of software. Simulation and verification of software.

3D

Internal structural failure.

Unexpected stresses. Improper handling or assembly of payload

Trajectory altered. Payload system inoperable. Failure may propagate to vehicle body.

1C Rigorous ground testing and simulation. Proper storage, handling, and assembly procedures followed.

1E

Roller jam in cam. Roller failure.

Unexpected and rapid deployment of blades. Damage to rollers or cam.

Incorrect deployment of one or more drag blades. Trajectory altered. Flight goals not met

2C Ground test of deployment system, add lubricant if necessary. Proper storage, handling, and assembly procedures followed.

2D

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4.5.6 Environment Safety Analysis The following table shows the Environmental Safety risks, possible causes, risk ratings, and mitigation strategies.

Table 4-11: Environmental Safety Risk Analysis

Risk Cause Overall Environmental Impact

Risk Rating

Mitigation Strategy Post Mitigation Rating

Rocket harms wildlife on takeoff or landing.

Animal presence at landing site / launch pad

Harming an animal is detrimental to the ecosystem.

1D The team will visually ensure no animals are present at launch site or nearby areas.

1E

In flight Rapid Unplanned Disassembly (RUD) of rocket.

Failure of vehicle or payload structure.

Rocket debris is detrimental to the ecosystem. Falling debris may also present a risk to nearby personnel.

1C Ensure a high factor of safety. Use eco-friendly parts. Use a reliable motor.

1E

Parachute deployment failure.

Charges fail to deploy parachute. Parachute is packed incorrectly.

The rocket will descend at a high velocity, scattering debris at crash site. This debris is detrimental to the ecosystem. Additionally, the incoming rocket may present a risk to nearby personnel.

1C Rigorous ground testing of recovery systems. Parachute packed carefully and checked.

1E

Launch pad fire. Rocket engine ignition causes nearby vegetation to catch fire.

A wildfire is detrimental to the local ecosystem. Fire may present a risk to nearby personnel.

1C Launch pad will be cleared of any easily combustible materials (eg. tree branches). Team will have fire extinguisher on hand.

3E

Carbon Dioxide released during launch.

Rocket motor releases gas detrimental to the environment.

Gases released contribute towards global warming.

3B Team will arrange donation to plant trees that scrub CO2.

4B

Trash left at launch site.

Trash is improperly disposed of at launch site.

Litter is detrimental to the ecosystem.

4A Team will bring trash bags to dispose of trash. All parts will be documented and accounted for before team leaves launch site.

4D

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Table 4-12: Environmental Impact on Rocket Risk Analysis Risk Overall Effect Risk

Rating Mitigation Strategy Post

Mitigation Rating

Animal interferes with rocket / components / personnel.

Potential damage / loss of rocket or components.

3D Team will not leave rocket components in areas near local wildlife. Team members will not approach local wildlife.

3E

High humidity. Warping of rocket components.

2B Rocket components will be made of materials resistant to warp.

4B

High temperatures. Warping of components. Overheating of electronics.

2B Rocket components will be made of materials resistant to temperature change. Electronics will be well cooled.

4B

Low temperatures. Warping of components. Plastic may become brittle.

2D Rocket components will be made of materials resistant to temperature change.

4D

High winds. Rocket trajectory may be affected. Parachute may drag rocket and cause damage on landing.

2C Rocket stability will be designed so that the wind impacts trajectory as minimally as possible. Launches will be planned in advance according to weather forecasts to ensure timely completion.

3D

Wet conditions / Rain / Mud.

Electronics may be damaged by rain. Body structure may be damaged.

2C All electromechanical and electrical parts will be shielded. Launch will be postponed in case of adverse conditions.

4C

Vehicle Cannot be Found.

GPS unit failure. High winds cause extreme drift.

1C Rocket will use additional GPS units for redundancy. Vehicle will have identifiable, bright paint scheme.

1E

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4.5.7 RF Safety Analysis The team rocket will have wires running to both black powder charges. The longest wire will be no longer than 20 inches. The rocket payload will also carry a RF transmitter. The altimeters create a voltage discharge across the black powder charges to detonate them at the correct altitudes.

It is possible for the RF transmitter to induce a voltage across the wires which may prematurely detonate the black powder charges.To test this, 80 inches of wire [details] was placed next to the RF transmitter. A 10kohm resistor was used in place of a black powder charge, and an oscilloscope measured voltage across the wires. The RF transmitter was powered on, and any voltage change was noted. The RF transmitter was powered on for twenty minutes.

The testing revealed no voltage change due to the RF transmitter. While further testing will be done, the team believes the black powder charges will be safe. The transmitter is a low power transmitter, and the altimeters put 5 volts across the charges to detonate them. Additionally the wires in the rocket are a quarter the length of the tested wire.

Nonetheless the team has developed risk mitigations for this issue. The electronics will be shielded as mentioned in requirement 2.12.2. The electronics will not be turned on until the rocket is on the Launchpad as to minimize any RF induction.

4.5.8 MSDS All Material Safety Data Sheets are available on the team website http://rrs.union.rpi.edu/

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5. Payload Criteria The objective of the payload system is to satisfy Experiment Requirement 3.3 as well as Vehicle Requirement 1.1. Both of these requirements will be met using the RRS team’s Roll-Blade design, which will be explained in detail in 5.2.

5.1 Payload Subsystem Timeline Table 5-1 below describes the payload subsystem timeline during the launch vehicle ascent, after motor burnout has completed

Table 5-1: Payload Phase Timeline

Phase Action

Zeroing Phase Post motor-burnout. Gyroscopes will detect residual roll and deploy blades to eliminate remaining angular velocity.

Roll Phase First set of blades deployed. Angular position and velocity tracked by gyroscopes and magnetometer while deployment length is adjusted accordingly.

Counter-Roll / Rezeroing Phase

First set of blades is retracted while second set is deployed in order to counter the induced roll. Angular position and velocity are again tracked and the deployment length is adjusted accordingly.

Active Drag Phase Following Rezeroing. Both sets of blades are deployed at equal lengths in order to actively decrease vertical velocity of the rocket. The microprocessor actively monitors altitude and velocity in order to match an idealized altitude model.

5.2 Design of Payload Equipment Verification of this design can be found in Section 7.

5.2.1 Mechanical Design

5.2.1.1 Roll-Blade Subsystem Vehicle Requirement 1.1 states that the rocket must deliver the payload to an altitude of 5,280 feet. The rocket motor has been selected such that if the rocket were to fly without active drag control, the final altitude would be slightly above this height. The drag control will be dynamically controlled in flight to decrease the vertical velocity of the rocket and adjust the maximum height of the rocket to exactly 5,280 feet at apogee.

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The roll-blade subsystem will also fulfill Option 2 of Payload Requirement 3.3, “Roll Induction and Counter Roll.” Two sets of blades will induce both the roll and counter roll. The blade system is, in reality, two sets of an identical system. Each system consists of the 3 blades, a “straight plate,” a “cam plate,” a stepper motor, and a plate to secure the stepper. One of the two systems is flipped and placed on top of the other, creating the complete “Roll-Blade Subsystem” pictured below. Having the two sets on the same plane diminishes turbulent effects that would propagate if the sets were separated vertically along the rocket. This will allow for easier CFD analysis.

Figure 5-1: CAD Model of Assembled Payload Subsystem

The blades are have the airfoil shape of a NASA NLF15 airfoil and will serve to impart a roll moment about the rocket in order to accomplish Option 2 of Payload Requirement 3.3. The airfoils will be attached to threaded rods which will be passed through the tracks of two plates - the “cam plate” and the “straight plate” either above or below them. The threaded rods will be fitted with rollers and nuts. The rollers will be in place to reduce resistance in the tracks as the actuation occurs; and the nuts will be in place at the end of the rods in order to constrain them. A third set of plates will be utilized in order to mount the motors that will drive the actuation. An exploded view of the subsystem can be seen below.

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Figure 5-2: Linear Plate (left) and Cam Plate (right) from Top View

Figure 5-3: Exploded View of Payload Subsystem Components

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5.2.1.2 Integration Plan The Roll-Blade subsystem will be located in a body tube directly above that which contains the fin can. The target location can be seen below in the rendered CAD model.

Figure 5-4: Target Location of Payload Subsystem

As mentioned in section 3.2, the basic integration plan is to secure the payload can to the vehicle by using 12 screws that will fasten the L-brackets to the body tube inner wall. Along the bottom 6 mating locations, a body tube coupler will come between the outer wall and the payload can, while the top 6 mating locations will not have this extra layer. This has been noted and the solution to this will be to add spacers to the top 6 locations to ensure equal spacing between the L-bracket and the body tube.

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The blades will extend from the body through rectangular slots cut into the airframe, described at the end of section 4.1.3.1. As mentioned, the stress concentrations caused by the removal of material is substantial, and will require reinforcement. These reinforcements will be interior and made with payload integration in mind in such a way to not interfere with integration. 5.2.1.3 Computational Fluid Dynamics Results Using ANSYS Fluent as a design tool, work has been done to analyze the effect of the blades. The motivation behind the analysis is to be a means of verification of the system, to provide early indications of the effectiveness of the system without the need of wind tunnel testing. Simulations were run in Mach 0.5 flow, to model the rocket shortly into its coast phase. With one set of blades fully deployed, the model was simulated using several different angular velocities, to find where the torque produced by the blades would equal zero. This angular velocity would correspond to the maximum angular velocity the rocket could reach. These simulations suggest that the the maximum roll rate the blade system can induce on the rocket is 25 revolutions per second. This number represents the maximum angular velocity that can be expected with a set of blades at full deployment. An actual flight will not require the blades to deploy fully, resulting in a more manageable angular velocity to complete the objective.

5.2.1.4 Defense of Current Design The current design is the leading alternative as presented in the PDR. This roll-blade design is favored over alternatives (e.g., reaction wheel) as it does not add as much mass to the system and the motors do not bear a direct load (as it does in the geared blades design). Additionally, there are fewer critical moving parts for the cam mechanism than in the alternatives (rack and pinion, or crank mechanisms would require many moving parts to redirect inputted motion for each individual blade). Furthermore, the natural cylindrical envelope of the cam mechanism makes efficient use of the natural cylindrical interior of the rocket body.

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5.2.2 Electrical Design Payload Electronics The electronic payload is designed to control the stepper motors driving the cam system based on data from various sensors. Excluding the microcontroller, wires, and other miscellaneous parts, the electronic components are listed in the table below.

Table 5-2: Electrical Components

Name Pin Usage Retailer URL

H-Bridge

5-6 GPIO Digikey http://www.digikey.com/product-detail/en/texas-instruments/SN754410NE/296-9911-5-ND/380180

GPS 2 Serial Adafruit https://www.adafruit.com/products/746

Radio Level Shifter 0 Sparkfun https://www.sparkfun.com/products/12009

Radio Socket 2 Serial Sparkfun https://www.sparkfun.com/products/11373

SD 4 SPI Adafruit https://www.adafruit.com/product/254

Data 2 Serial Adafruit https://www.adafruit.com/product/1604

Regulator 0 Digikey http://www.digikey.com/product-detail/en/texas-instruments/LM2575N-5.0-NOPB/LM2575N-5.0-NOPB-ND/212628

An Arduino Zero microcontroller controls the electronic components. The Arduino Zero features an ARM Cortex M0+ microprocessor that provides 32 bit arithmetic at a clock rate of 48Mhz. Additionally, the Arduino Zero contains all the necessary pins required to operate the components listed above. Typically, the Arduino Zero does not support having 3 different serial connections running simultaneously. However, the processor does, and with some minor changes to the Arduino IDE code, additional hardware serial connections can be made. Two Lithium Ion 18650 batteries will power the electronic components. Connected in series, the batteries produce ~7.4V. The 18650 batteries can support a far higher current consumption than expected (<500mA with 2.0A peak) and have enough capacity to last for well over the entire flight. The 7.4V signal will be regulated down to 5V through two switching voltage regulators. The first regulator will be solely dedicated to powering the H-Bridge and motors, while the second regulator will provide 5V or 3.3V power (via the Arduino Zero) to the rest of the components. Except for the Xbee, all the components will

43

communicate using 3.3V logic. The Xbee will communicate to the Arduino through a level shifter, to correct for its 5V logic. All the parts for the electronics are chosen to be compatible with a 2.54mm breadboard. The final wiring will be converted into an EAGLE schematic, which will produce a PCB to mount to components. The usage of a PCB will minimize the size of the electronics (by increasing density) while increasing the reliability of the electrical connections. Additionally, removing the wires will reduce time necessary for assembly and troubleshooting while helping to prevent human error. The electronics may be turned on using either a hall effect switch or a screw switch (currently undecided). The hall effect switch will turn the power for the circuit on or off depending upon the strength of the magnetic field around the hall effect sensor. This allows a neodymium magnet to be used to turn on the electronics by touching the magnet to the side of the rocket where the switch is fixed internally. This makes assembly extremely easy, as no holes need to be drilled to enable access to a screw switch. The status of the radio connection can determine the power state of the electronics. We believe that the hall effect switch will be an ideal way to turn on and off the rocket, although knowing the risks involved, we plan to test the hall effect switch extremely rigorously before we rule out using the tried and true screw switch. The software design will depend upon the actual performance of the parts ordered. For example, a noisy magnetometer would cause us to remove the derivative component of a PID controller because of the huge derivative values resulting from noisy readings. If a sensor tends to drift, we might need to account for that using a measurement from another sensor. In general the code will function as follows:

- Begin reading the GPS, Radio, and the Altimeter - Transmit periodic status messages to the radio while waiting for a go

reply - Prime the system to activate when the Altimeter altitude changes

significantly for multiple reads. - Wait until X successive reads above Y altitude AND after Z seconds while

constantly transmitting data. - Using a PID controller, rotate the rocket as required using the stepper

motors. - Retract the flaps and shut off power to the motors. - Transmit and log data until landing - Stop logging data when landed. (determined by X successive reads with

small drift in altitude). - Continue sending GPS data until the batteries run out or the circuit is

turned off.

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5.3 Risk Assessment Table 5-3: Payload Risk Assessment

Risk Likelihood Impact Risk Reduction Plan

Blades deploy early / too quickly

Medium Change of trajectory and possible failure of payload mission

Redundant checks in code to assure motor burn has ceased

Blades do not deploy / Unable to deploy

Low Overshoot target height, no rolling occurs

Ground test using sample input to observe resultant deployment. Integration ground testing.

Blades deploy late / too slowly

Medium Possible failure of payload mission, possible overshoot

Increase requirement on imparted torque to add safety factor

Blades deploy asymmetrically

Low Asymmetric blade deployment, trajectory altered

Test 01 in section 7.1

Blade structural failure Low Asymmetric blade deployment, trajectory altered

Buckling, bending, and torsion load calculations done beforehand

Payload instrumentation failure

High Erroneous data fed into microprocessor, failure of payload

Redundancies in instrumentation and error checks in code

Algorithm / Misc software failure

Medium Possible mission failure, trajectory altered

Error checks in code, ground testing of final code

Internal structural failure Low Possible mission failure, center of gravity altered

Bearing loads calculated and factor of safety added

Roller jam in cam system Low Possible failure of payload mission, trajectory/apogee altered

Ground test of deployment system, add lubricant if necessary

Structural failure of body tube at payload slots

Low Possible mission failure, altered trajectory/apogee

Complete structural analysis with scrutiny near payload sections

Structural failure of body tube due to workmanship error

Medium Possible mission failure, altered trajectory/apogee

Planning out all workmanship processes in advance and analyzing potential risks

Fins change alignment in flight

Low Possible mission failure, altered trajectory/apogee

Redundant reinforcement of fins along the motor mount

Structural failure of the motor mount

Low Possible mission failure, altered trajectory/apogee, potential major damage to all internal components

Complete analysis of centering rings and motor mount tube

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6. Launch Operations Procedures Table 6-1 below shows detailed launch operations procedures. Preconditions list the requirements before attempting a step, as well as any personal protective equipment or personnel necessary to complete the step. Postconditions provide indicators of successful completion of a step. Possible hazards introduced from an error in the step are also listed.

Table 6-1: Launch Operations Procedures

Stage Preconditions Step Postconditions Possible Resulting Hazards

General Preparation

Components gathered

Ensure all parts, materials, and tools are accessible to the team

All parts, material, and tools are accounted for.

N/A

Recovery Preparation

Shroud Lines attached to Main and Drogue Parachutes

Tightly fold parachutes into packing shape. Wrap shroud lines around each chute to secure shape. Store in individual bags

Parachutes are folded in preparation for packing.

Incorrect folding may hinder the packing process or lead to bad parachute deployment.

Recovery Preparation

Shock cord is attached to recovery hardware mounting points

Tie a knot in each length of shock cord to form a loop in the middle. Attach a quick link to the loop, and parachute shroud lines to the quick link

Shock cords are fitted with quick links and parachutes are attached. Ensure parachutes are attached to the correct shock cord.

Incorrect tying of knot may lead recovery system failure. Incorrect attachment of parachutes may lead to recovery system failure

Payload Preparation

Static testing of piece-parts and sizing checks

Assemble payload Assure proper function of payload out of the rocket body

Improper assembly can result in payload failing to function properly

Payload Preparation

Premeasure dimensions to ensure fit. All components are counted for before insertion

Insert payload Payload can is completely in place and secured.

Improper installation can result in unsecured mass and improper function of payload

Payload Preparation

Premeasure dimensions and alter if necessary

Test drag blade clearance

Perform close clearance test with payload can inside rocket body and fully mounted

If blades cannot clear slots, payload will not be able to fully function

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Stage Preconditions Step Postconditions Possible Resulting Hazards

Motor Preparation All motor pieces are accounted for. Club mentor present

Prepare motor by inspecting all pieces for any signs of damage, and greasing necessary components

All motor components are accounted for and free of any signs of damage. Approval from mentor to continue with motor assembly.

Failure to detect signs of damage may lead to catastrophic motor failure

Motor Preparation Motor components inspected. Club mentor present

Assemble motor according to the instructions included in Appendix C

Motor correctly assembled under the supervision of the club mentor.

Failure to exactly follow instructions may lead to catastrophic motor failure

Motor Preparation Motor correctly assembled. Club mentor present

Carefully insert motor into motor mount and secure with motor retainer cap.

Motor inserted into vehicle and properly secured.

Failure to properly insert motor could lead to catastrophic launch failure

Recovery Preparation

All sources of heat and flame are removed from vicinity. Rubber gloves doned for those handling black powder

Prepare black powder ejection charges by carefully measuring out for both upper and lower separation.

Ejection charges ready for use.

Using the incorrect amount of black powder could result in recovery system failure or vehicle damage

Recovery Preparation

Ejection charges prepared, avionics sled assembled, all electronics turned off

Apply sealant to rim of avionics bay and assemble. Attach ejection charge wires to appropriate terminal blocks

Avionics bay is appropriately sealed, and ejection charges are correctly attached.

Improper seal of avionics bay could lead to altimeter malfunction. Improper ejection charge connection could lead to recovery system failure

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Stage Preconditions Step Postconditions Possible Resulting Hazards

Recovery Preparation

Drogue Parachute folded and correctly attached to shock cord

Pack thermal insulation into lower airframe. Roll any unprotected shock cord into folded drogue parachute, wrap chute in nomex blanket, then pack into lower airframe

Drogue parachute is packed; chute and extra shock cord is protected from motor and ejection charges

Unprotected parachute or shock cord may be damaged by motor or ejection charge detonation, which may lead to recovery system failure

Recovery Preparation

Main Parachute folded and correctly attached to shock cord

Pack insulation into upper airframe. Roll any unprotected shock cord into folded main parachute, wrap chute in nomex blanket, then back into upper airframe

Main parachute is packed; chute and extra shock cord is protected from ejection charges

Unprotected parachute or shock cord may be damaged by ejection charge detonation, which may lead to recover system failure

Vehicle Preparation

Drogue Parachute correctly packed inside lower airframe

Conjoin upper airframe with lower airframe, then assemble with shear pins, taking care not to damage pins

Upper airframe is securely attached to lower airframe via shear pins

Improperly securing the separation point may lead to recovery system failure. Damage to shear pins may lead to premature recovery deployment

Vehicle Preparation

Main Parachute correctly packed inside upper airframe

Attach nose cone to upper airframe, then secure in place with shear pins, taking care not to damage pins

Launch vehicle is completely assembled

Improperly securing the separation point may lead to recovery system failure. Damage to shear pins may lead to premature recovery deployment

General Preparation

Launch vehicle is completely assembled. Club mentor is present.

Double check final assembly with club mentor, paying close attention to shear pins

Launch vehicle assembly is verified by experienced club mentor

N/A

Vehicle Preparation

Launch vehicle is completely assembled and verified by club mentor

Mark the Center of Pressure according to simulation profiles, and Center of Gravity according to balance

CP and CG are marked on vehicle, ready for RSO inspection

Vehicle will not be permitted to launch unless RSO’s minimum stability margins are satisfied

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Stage Preconditions Step Postconditions Possible Resulting Hazards

General Preparation

Launch vehicle is completely assembled, verified, and CP and CG are marked

If the launch venue requires flight cards to launch, fill one out

Launch vehicle information is given on flight card

Vehicle will not be permitted to launch without a filled out flight card

Recovery /Payload Preparation

Launch vehicle is completely assembled and arming switches are accessible

Double check all electronics by toggling arming switches. Turn all electronics back off prior to proceeding

Functionality of electronics has been verified

Launching without verifying functionality of on board electronics may lead to payload or recovery system failures

General Preparation

Launch vehicle is completely assembled, CP and CG are marked, assembly and electronics are verified

Consult the RSO to be assigned a launch pad, wait for appropriate time.

RSO is notified that vehicle is ready to launch

N/A

General Preparation

Launch vehicle is assembled, marked, and verified; RSO has given all clear to enter launch area

Safely transport rocket to designated launch pad

Rocket safely moved to designated launch pad

Mishandling rocket could damage the launch vehicle. Special care should be taken to keep the rocket level and avoid disturbing shear pins

General Preparation

Launch vehicle has been transported safely to designated launch pad

Carefully slide launch vehicle onto launch rail, then lift rail into upright position

Rocket correctly mounted on the launch rail, which has been moved to the upright position

Rail buttons could be damaged if rocket is carelessly mounted on launch rail

Motor Preparation Rocket is mounted upright on the launch rail

Insert igniters all the way into the motor, and secure into place by lightly taping to launch vehicle. Short out igniter circuit.

Igniters are inserted into the motor, and ignition circuit is shorted out

Improper insertion of igniters may lead to faulty motor ignition. Failure to short ignition circuit may lead to premature motor ignition

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Stage Preconditions Step Postconditions Possible Resulting Hazards

Recovery/ Payload Preparation

Launch vehicle is mounted upright on launch rail with igniters inserted

Turn on electronics for the recovery and payload subsystems. Wait for appropriate feedback from altimeters

Electronic systems are engaged, and launch vehicle is completely prepared

Failure to turn on electronics will result in payload or recovery system failure.

Motor Preparation Launch vehicle is mounted upright on launch rail with electronics engaged and igniters inserted

Unshort the igniter circuit, and attach the alligator clips from the ignition system

Igniters are wired up to the ignition system

Failure to correctly wire ignition system may lead to faulty ignition, or dud ignition

General Preparation

Launch vehicle is completely prepared and ignition system is correctly wired

Return to a safe distance, and submit flight card to LCO or RSO

Rocket is ready to launch, registered with LCO, and awaiting ignition

Failure to notify correct personnel may delay launch

Launch Rocket is ready and registered to launch

Launch Launch vehicle has launched

See section 4.5

Post-Launch Recovery

Vehicle has launched, RSO has given all clear signal for recovery

Ensure all energetics are inert. Inspect launch vehicle for initial damage. Bring all components back to preparation area

Launch vehicle has been retrieved and inspected for initial damage

Care should be taken around any lingering energetics, and around the motor components which will retain heat for some time after launch

Post-Launch Recovery

Launch vehicle has been returned to staging area

Clean reusable motor components according to manufacturer specifications

Motor is cleaned and ready for additional use

Failure to thoroughly clean motor components may damage them over time, or hinder future uses.

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6.1 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 to be tested on the ground before a launch to verify that it is programmed and wired correctly. 6.2 Motor Preparation The motor will be mounted within the casing and the integrity of the interface with the rocket body will be checked. The retention cap is screwed on. The igniter charge is installed securely before launch. 6.3 Launcher Setup The rocket will be slid onto the launch rod horizontally and then raised to a vertical position. 6.4 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 shortened such that static discharges do not trigger the igniters. The igniter will be fully inserted and stopped off so that it is firmly held in place. The igniter will then be connected to the launch console. 6.5 Troubleshooting A static test of the flight computer telemetry will be conducted on the pad to ensure that the radio link between the RRS team and the rocket is operational. The flight computer will be checked for full functionality and sensor readings will be checked to ensure that they are connected/functioning properly. 6.6 Post-Flight 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.

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7. Project Plan 7.1 Testing

Table 7-1: Testing Plan

Test # Subsystem Objective Success Criteria Testing Variable Methodology

01 Payload Static deployment test with 3 blades, cam/straight plates, and stepper motor.

Equal deployment of blades in reasonable time frame

Deployment length and time of deployment

- Assemble Roll-Blade subsystem - Deploy blades, observing symmetry and deployment time

02 Payload / Integration

Clearance test for blade-slot interface

No interference Distance between slots and blades

- Assemble payload can inside body tube, aligning blades with slots - Deploy blades, checking for close clearances

03 Payload Deployment given sample conditions

Actively adjusting deployment given sample code

Deployment length and time of deployment

- Assemble Roll-Blade subsystem - Feed sample code into microprocessor - Observe deployment

03 Recovery Ejection charge ground test

Ejection charges separate desired section without damaging vehicle

Strength of ejection charge

- Assemble recovery system - Activate charge(s) - Observe deployment

04 Recovery Parachute deployment test

Rocket lands without damage

Deployment altitude

- Assemble recovery system - Observe parachute deployment

05 Recovery Recovery system stress testing

No components (eyebolts, bulkheads, etc…) fail and rocket lands without damage

Forces on recovery system

- Assemble recovery system - Observe parachute deployment

06 Vehicle General assembly testing

All components assemble as expected by the vehicle design

Assembly time, ease of assembly, tools required

- Assemble vehicle - Observe any difficulties or impediments

07 Vehicle Body tubing stress testing

Test pieces of body tubing do not fail within expected stress ranges

Forces on airframe

- Create piece of body tubing suitable for testing - Use Instron machine for stress testing - Observe tolerable stress ranges

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7.2 Requirement Verification Table 7-2: Requirement Verification

Req # Requirement Definition

Design Feature

Verification Plan Testing Steps

Status Referenced Sections

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

Rocket Mass, Rocket Motor, Design

OpenRocket Simulations Payload Mass Rocket material selection Motor Selection Stay as close as possible to original Mass Estimates

Read apogee at several launches in different weather conditions

Simulations - run Mass - calculated Materials - selected Error in mass estimates - rocket construction is not complete

2.2 4.1.4 4.1.5 4.3.2

1.4 Reusable Body Strength (fins, airframe, parachutes, etc.)

Stress Analysis of Vulnerable Components Use of Large Factors of Safety in Critical Components Use High-Strength Epoxies Store all components Safely Design Followed Components inspected upon order arrival show no signs of damage

Launch in appropriate weather multiple times

Analysis - complete Factors of safety - calculated Epoxies - ordered Components - on order

4.1.7 4.5

1.5 Four or fewer Independent Sections

General Design

Design four or fewer independent sections for the launch vehicle Follow design

Rocket is launched and recovered at most four independent sections

Design - Complete Construction - Underway

4.1.4

1.6 Single Stage General Design, Motor Design

Design only a single stage rocket Follow design

Rocket is launched with only a single stage motor

Design - Complete Construction - Underway

4.1.4

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Req # Requirement Definition

Design Feature

Verification Plan Testing Steps

Status Referenced Sections

1.11, 1.13

Commercially available solid motor propulsion system not exceeding Class L

Motor Ensure Mass and General Rocket Design allow for a motor not exceeding Class L approved by the TRA and NAR Follow design Stay as close as possible to original mass estimates

Motor not exceeding Class L is used at several launches in different weather conditions

Design - Complete Construction - Underway

4.1.4

1.9 Capable of Launch by 12 V DC firing system

Motor, Motor Retainer

Select a motor retainer that allows for access to motor Select a motor able to be launched with 12V DC firing system Follow design Safely store motors

Launch rocket on standard 12 V DC firing system multiple times

Design - Complete Motor - Selected/Ordered Construction - Underway

4.1.3 4.1.4

1.14 Minimum static stability margin of 2 Cal at rail exit

Motor, Vehicle Design

Use simulation data to ensure placement of vehicle CG and CP creates a static stability margin of at least 2 Cal Follow design Stay as close as possible to original mass estimates Update simulation profiles to reflect any change

Launch rocket several times to ensure stability off the launch rail

Simulations - Complete Construction - Underway

4.4.2

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Req # Requirement Definition

Design Feature

Verification Plan Testing Steps

Status Referenced Sections

1.15 Minimum velocity of 52 f/s at rail exit

Motor, Vehicle Design

Use simulation data to select a motor with sufficient thrust to propel the rocket off the rail with a velocity of at least 52 f/s Follow design Stay as close as possible to original mas estimates Update simulation profiles to reflect any change

Launch rockets several times, use flight data to ensure minimum exit velocity

Motor - Selected Construction - Underway

4.1.3 4.1.4 4.4.2

2.1 Electronic Dual Deploy

General Recovery Design (parachutes, shock cord, ejection charges, etc.)

General Design is Dual-Deploy Use of a Drogue and Main Parachute Design for Multiple Separation Points Design Followed

Rocket launched with appropriate recovery system, Drogue and Main parachutes deploy at different points in flight

Design - Complete Construction - Underway

4.1.3 4.1.4 4.3

2.1 Drogue Deploys at Apogee, Main Parachute Deploys at much lower altitude

Parachutes, ejection charges, General Rocket Design, Altimeters

General Design has Drogue Parachute at First Separation Point and Main Parachute at another Simulations run with required deployment locations Follow design Set Altimeters to required ejection charge deployments (and keep track of which side is deployed when)

Rocket launched multiple times with parachute deployment at required times

Design - Complete Simulations - Run Construction - Underway

4.1.3 4.1.4 4.3

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Req # Requirement Definition

Design Feature

Verification Plan Testing Steps

Status Referenced Sections

2.10 Shear Pins hold rocket sections together until Parachute Deployment

Shear Pins, Ejection Charges

Shear Pin Strength Accurately Calculated and inserted into Recovery Design Ejection Charge Strength Accurately Calculated Follow design Shear Pins show no signs of damage before install

Rocket launched multiple times with parachute deployment at required times

Shear Pin Analysis - Underway

4.3.3.2

2.3 Independent Sections have less than 75 ft-lb of KE at Landing

General Rocket Design (Mass), Parachutes

Kinetic Energy of Each Independent Section Analyzed Accurately Follow design Stay as close as possible to original mass estimates Parachutes and Recovery system components show no signs of damage upon order arrival

No damage or hard landing evident in any section after multiple test launches in varied flight conditions

Analysis - Complete Construction - Underway Parachutes - Ordered

4.1.3 4.1.4 4.3 4.4

2.5 Redundant, Safe Altimeters

Altimeters, Supporting Recovery Electronics

Select at least 2 commercially available altimeters Design has an independent power supply to each Altimeter Follow design Altimeters stored safely All components of supporting electronics and altimeters tested before installation

Recovery System Operates as Expected, Altimeters report similar apogees in all test launches

Altimeters - Selected Design - Complete Construction - Underway Testing - Incomplete

4.1.3 4.1.4 7.1

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Req # Requirement Definition

Design Feature

Verification Plan Testing Steps

Status Referenced Sections

3.3.1. Teams shall design a system capable of controlling launch vehicle roll post motor burnout.

Roll-Blade subsystem

Roll-Blade subsystem is designed to control roll post burnout

See Tests 01-03

Design completed. Starting manufacturing

5.2-5.3

3.3.1.1.

The systems shall first induce at least two rotations around the roll axis of the launch vehicle.

Roll-Blade subsystem

CFD analysis confirms blades will impart enough torque to roll the rocket two full rotations post burnout.

CFD, Tests 01-03

CFD completed, awaiting materials for manufacturing

5.2.1.1 and 5.2.2

3.3.1.2.

After the system has induced two rotations, it must induce a counter rolling moment to halt all rolling motion for the remainder of launch vehicle ascent.

Roll-Blade subsystem

The opposing set of blades will halt rotations. They are able to impart an equal force in the opposite direction.

CFD, Tests 01-03

CFD completed, awaiting materials for manufacturing

5.2.1.1 and 5.2.2

3.3.1.3.

Teams shall provide proof of controlled roll and successful counter roll.

Roll-Blade subsystem

Redundant gyroscopes will log rotational data throughout the launch.

Test 03 Preliminary algorithms being written. Testing to follow.

5.2.1.1 and 5.2.2

3.3.2. Teams shall not intentionally design a launch vehicle with a fixed geometry that can create a passive roll effect.

Fins Fins will be aligned as they normally would - parallel to rocket body in order to not induce a roll passively.

N/A Awaiting rocket body materials

4.1.4

3.3.3. Teams shall only use mechanical devices for rolling procedures.

Roll-Blade subsystem

No compressed gas system or any other non-mechanical device will be used to impart a roll.

N/A Design completed. 5.2

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7.3 Team Derived Requirement Verification

7.3.1 Statement of Team Derived Requirements Table 7-3: Team Derived Requirements

Req # Requirement Definition Required Value

1A1 Max friction forces (set of 3 blades)

1N (static), 0.5N (kinetic)

1A2 Minimum angular velocity 2 rad/s imparted by each set

1B1 Mechanical subsystem fully assembled outside of rocket body before installation

N/A

1B2 Total weight must not exceed 6 pounds

6 pounds

1B3 Center of gravity must lie within one inch of axial centerline

One inch

2A1 Stepper motor starting torque

27Ncm

2A2 Maximum battery weight 100g

2B1 Instrument response time <1ms

2B2 Error rate of instruments below datasheet maximum

Varies

3A1 Max forces on attachment hardware, parachutes, etc...

Ultimate strength dependent on material of component

3A2 Line attachments do not tangle

N/A

3B1 Min black powder required for successful ejection charges

1.7 g

3C1 Both parachutes deploy N/A

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7.3.2 Vehicle Requirement Verification

Table 7-4: Vehicle Verification

Req # Requirement Definition Verification Method Current Status

3A1 Max forces on attachment hardware, parachutes, etc...

Stress analysis Initial stress calculations show all components are within factor of safety of 1.7 or above

3A2 Line attachments do not tangle

Multiple tests before final launch

Awaiting materials; swivel joint designed to prevent tangle

3B1 Min black powder required for successful ejection charges

Calculation Initial calculations completed with value of 1.7 g for both charges

3C1 Both parachutes deploy Multiple tests before final launch

Awaiting materials

7.3.3 Payload Requirement Verification

Table 7-5 - Payload Verification

Req # Requirement Definition Verification Method Current Status

1A1 Max friction forces (set of 3 blades)

Simple torque test Awaiting materials

1A2 Minimum angular velocity of 2 rad/s imparted by each set

Computational Fluid Dynamics

Testing Indicates max angular velocity of 150 rad/s

1B1 Mechanical subsystem fully assembled outside of rocket body before installation

In-lab full assembly Awaiting materials. CAD model indicates that this should be possible.

1B2 Total weight must not exceed 6 pounds

Weight measurement

Awaiting materials. Initial mass estimate of mechanical components of ~4.5lbs.

1B3 Center of gravity must lie within one inch of axial centerline

CAD verification Awaiting the modeling of electrical components. Mechanical components meet specifications.

2A1 Stepper motor starting torque Test starting

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torque

2A2 Maximum battery weight Weight measurement

2B1 Instrument response time Software timing

2B2 Error rate of instruments below datasheet maximum

Test with known expected results

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7.4 Budget Plan Table 7-6 provides a summary of important sections, as well as our projected bottom lines. Appendix C contains the fully detailed budget.

Table 7-6: RRS Budget Summary

Budget Summary

Vehicle Design Team Expenses $1,218.96

Recovery Team Expenses $398.95

Payload Team Expenses $798.16

Travel Costs $3,150.00

Income $9,825.00

Total Expenditures $5,566.07

Total Budget $4,258.93

7.5 Funding Plan At this point, the RRS has sufficient funding for the construction and testing of the rocket as well as travel to and from Huntsville for our team. This funding came from club membership dues, the RPI School of Engineering, and the RPI Mechanical, Aerospace, and Nuclear Engineering (MANE) Department.

In addition, we have also raised funds through a program at RPI called WeR Gold. This program sponsors student-run projects by reaching out to alumni and friends of Rensselaer. The information the RRS has submitted to WeR Gold has been 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. We expect to raise an additional $2500 this year to help our team continue to compete in coming years.

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7.6 Timeline The RRS has updated the Gantt chart schedule that has been used for planning during NSL. Despite delays reported in the PDR, the subscale launch vehicle was launched on schedule. Unfortunately, the ordering process through RPI’s School of Engineering Purchasing department took slightly longer than expected. Due to these delays, parts were not received and thus could not be inspected before the fall semester ended. Notification has been received that the majority of the full-scale vehicle parts have arrived over winter break. This delay only slightly changes the construction schedule. These changes have been reflected in the update made to the Gantt chart. As before, the updated Gantt chart cannot be feasibly included in this report, and will be instead hosted here[1]

7.7 Educational Engagement Since the PDR in November, the RRS has not held any new educational engagement events. The RRS had originally planned to have a rocket building event and launch with a local Girl Scouts troupe, but unfortunately the troupe had to reschedule. The team is now hoping to hold this event in the early spring. In addition to this, the RRS is currently reaching out to local elementary schools with the aim of working with math and science teachers to help conduct hands on activities for the students.

[1] http://rrs.union.rpi.edu/nasa2017.html

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Appendix A: Milestone Review Flysheet

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64

65

Appendix B: Vehicle Design Assembly Drawings

66

67

68

69

70

71

Appendix C: Cesaroni L910 Assembly Manual

72

73

74

75

76

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Appendix D: Detailed Budget Income

Income in Accounts (start Fall 2016) $4,445.00

Membership Dues $1,280.00

School Funding $1,600.00

Fundraising Goal $2,500.00

Total $9,825.00

Vehicle Design Team

Item Unit Cost Quantity Total Cost

Fiberglass Nose Cone $104.99 1 $104.99

Phenolic Airframe Tubing (48") $41.99 2 $83.98

Airframe slotting $2.50 3 $7.50

Motor Mount $18.99 1 $18.99

Motor Retainer $44.00 1 $44.00

Centering Ring $7.29 2 $14.58

Airframe Fins $83.56 1 $83.56

Phenolic Coupler Tube (12") $14.99 3 $44.97

Vehicle Motor (Cesaroni 2653-L585-P) $132.95 3 $398.85

Motor Casing (Cesaroni P75-2G-HS) $297.00 1 $297.00

Epoxy $40.00 1 $40.00

Bulkplate $6.89 4 $27.56

Estes Pro Series II Kit (subscale vehicle) $24.99 1 $24.99

Subscale Vehicle Motor (Aerotech G80) $27.99 1 $27.99

Total $1,218.96

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Recovery Design Team

Item Unit Cost Quantity Total Cost

PerfectFlite Stratologger SL100 $54.95 1 $54.95

Raven3 Altimeter $115.00 1 $115.00

Rocketman Standard Parachute (2’) $40.00 1 $40.00

Skyangle Classic II (60”) $139.00 1 $139.00

Miscellaneous hardware $50.00

Total $398.95

Payload Design Team

Item Unit Cost Quantity Total Cost

Threaded rod, 6' section $2.55 1 $2.55

Airfoil Blades 6 $100.00

Cam Plates $11.74 2 $23.48

Straight Guide Plates $11.74 2 $23.48

Motor Mount Plates $11.74 2 $23.48

Servo $20.00 2 $40.00

Pin Rollers $8.88 12 $106.56

Threaded Rods for cam assemblies 6 $37.00

Motor Controllers $35.00 2 $70.00

Bulkhead $9.00 2 $18.00

Payload L-brackets $0.84 24 $20.16

Arduino $50.00 1 $50.00

Rollers - Exit Guides $5.00 12 $60.00

Manufacturing Costs $100.00

Threaded Rods for Constraints 3 $23.55

Triple Axis Accelerometer + Magnetometer $14.95 2 $29.90

GPS Receiver $20.00 1 $20.00

Xbee wireless transmitter $35.00 1 $35.00

SD module $15.00 1 $15.00

Total $798.16

79

Travel

Item Unit Cost Quantity Total Cost

Hotel Rooms $160.00 15 $2,400.00

Gas, tolls (per car, round trip) $250.00 3 $750.00

Total $3,150.00

Income $9,825.00

Spending $5,566.07

Difference $4,258.93

80