MIT Rocket Team USLI 2011 Flight Readiness Review

Transcription

1 MIT Rocket Team USLI 2011 Flight Readiness Review March 21, 2011 Project Valhalla

2 REPORT AUTHORS Christian Valledor Project Manager Andrew Wimmer Team Safety Officer, Tripoli Rocket Association Level 3 Ben Corbin EHS Representative Ryan McLinko Launch Vehicle Lead Jonathan Allen Payload Lead Eric Peters Recovery Lead Ben Couchman Avionics Lead Anna Ho Outreach Coordinator Jake Bograd-Denton Michelle Burroughs Jason Elizalde Jedediah Storey Leo Tampkins

3 NAMING CONVENTIONS NASA and the space industry have typically been very creative with their acronyms for spacecraft and space instruments. From LEM to SRB, NASA has come up with the most recognized TLAs the -world has ever known, approaching even greater rhetorical strength with C.O.L.B.E.R.T.. We in the MIT Rocket Team also share such an affinity for acronyms, so we decided to put some extra thought into the naming of our rocket project. We also honor NASA s tradition of naming conventions based off mythological gods from ancient civilizations. However, where NASA chooses its naming inspiration, we wholeheartedly disagree; the Greek and Roman gods are too overused for us to name our rocket parts after. Instead, we turned to the Norse gods, whose deeds are so epic and intense that modern comic book heroes are based off of them. While most people are familiar with the god Odin, few are familiar with the one-handed god if single combat, victory, and heroic glory, for which the third day of our modern week is named after. Scandinavians would have called this god in his human form a yager, which is German for "hunter" or someone who tracks animals. This is appropriate considering the goal of the UAV payload is to track targets on the ground and autonomously report their location. For these reasons, we are naming the entirety of our rocket the Tactical Yager Rocket TYR One of the most famous Norse gods is the Norse god of fire. A brother to Odin and father of a wolf, this god has inspired many names including a company that produces rocket engines. We at the MIT Rocket Team couldn't pass up the opportunity to use his name for a piece of our rocket. This is why we have named the structural section of the rocket that houses the motor and spits fire the Ludicrously Overpowered Kinetic Impulser LOKI Vikings were inspired to be brave in their fights because only the righteous and brave who die in battle would enter the gates of Valhalla to sit with Odin and fight by his side in the end of times. One son of Loki, an eight-legged horse that travels across land, sea, and air, is responsible for delivering fallen warriors to their glorious fate. This god carries those valiant warriors who die in battle over land and sea across the Rainbow Bridge through the gates of Vahalla. For this reason, we have named our motor the Single- Loaded Electrically-Ignited Propulsively Numerous Integrated Rocket SLEIPNIR From the basis of these Norse naming inspirations, there was only one logical choice for building a machine that would reach the heavens, named after the home of Odin himself Project Valhalla However, not even the Norse gods were powerful enough by modern standards when it came to naming the crown jewel of our rocket: the UAV payload we are ejecting from the rocket to perform a search mission. The one series of magnificent beings that exists today from which we can draw truly incredible names from is none other than the Transformers series. While casual fans of the series may often bicker about who their favorite Transformer is, often arguing between Optimus Prime and Megatron, there is

4 only one Transformer that truly and unequivocably incites emotions of grandeur, terror, destruction, power, and might, and it first appeared in the 1986 Transformers theatrical release as the Transformer that was so powerful it literally ate planets. For this reason, we have given our unmanned aerial vehicle the name UNmanned Integrated Craft for Rescue with Onboard Navigation UNICRON

5 TABLE OF CONTENTS Naming Conventions... 3 List of Figures... 7 List of Tables FRR Summary Team Summary Launch Vehicle Summary Payload Summary Changes Since CDR Launch Vehicle Changes Payload Changes Activity Plan Changes Launch Vehicle Testing and Design of Vehicle Rocket Design and Subsystems Subsystem Requirements and Descriptions Component, Functional, and Static Testing Safety and Failure Analysis Recovery Subsystem Parachute Choice Test Results from Ejection Charges and Electronics Failure Modes and Safety Analysis Mission Performance Predictions Flight Profile Simulation Safety and Environment Identification of Safety Officers Saftey and Failure Modes Analysis Potential Hazards Environmental Concerns Payload Integration Feasibility Describe Integration Plan Payload Integration Plan Installation and Removal, Dimensions, Precision Fit Tasking & Integration Schedule Payload Experiment Concept Science Value Payload Science Objectives Payload Success Criteria Payload Design Summary Assembly and Testing Planned Component, Functional and Static Testing Test Measurement, Variable and Controls Safety and Environment Identification of Safety Officer... 74

6 4.5.2 Payload Failure Modes Potential Hazards Environmental Concerns Launch Operations Launch Operations and Procedures Checklists Recovery Preparation Motor Preparation Igniter Installation Setup on Launcher Troubleshooting Post Flight Inspection Safety and Quality Assurance Saftey Analysis Environmental Concerns Activity Plan Budget Timeline Educational Engagement Conclusion

7 LIST OF FIGURES Figure 3-1: Overall Rocket Figure 3-2: Tube Coupler Segment Figure 3-3: Nose Cone Figure 3-4: Motor Centering and Retention Figure 3-5: Recovery System Bulkhead Figure 3-6: Sabot Overview Figure 3-7: Payload Integration Stacking (NOT TO SCALE) Figure 3-8: Recovery Configuration Figure 3-9: Sabot Hard Point Figure 3-10: MAWD Flight Computer Figure 3-11: ARTS2 Flight Computer Figure 3-12: ARTS2 Telemetry Transmitter Figure 3-13: ARTS2 Telemetry Receiver Figure 3-14: ARTS GUI Figure 3-15: ARTS Data Analyzer Figure 3-16: Avionics Package Figure 3-17: Axial Case BCs Figure 3-18: Lateral Case BCs Figure 3-19: Axial Case Results Figure 3-20: Lateral Case Results Figure 3-21: Motor Retention BCs Figure 3-22: Motor Retention Displacement Figure 3-23: Motor Retention Stress Figure 3-24: Predicted CM and CP Locations Figure 3-25: Cesaroni L1115 Thrust Curve Figure 3-26: Predicted Acceleration and Velocity Profiles Figure 3-27: Simulated Altitude Profile Figure 3-28: Contrail of the full scale test flight showing coning during flight Figure 3-29: COMPARISON BETWEEN THE MODEL ANGULAR ACCELERATIONS OF THE TEST FLIGHT CONFIGURATION (UNBALLASTED, LEFT) AND THE COMPETITION CONFIGURATION (BALLASTED, RIGHT) Figure 3-30: Tube-Tube interface Figure 3-31: Main parachute/shock cord attached to eye bolt and recovery system bulkhead Figure 3-32: Integrated avionics assembly, main parachute, sabot and UAV assembly 56 Figure 4-1: UAV Figure 4-2: UAV (Stowed Configuration) Figure 4-3: CP distribution for Various Airfoils Figure 4-4: CL Distribution for NACA 4412 Airfoil Figure 4-5: Access Port Bulkheads Figure 4-6: Electronics Bay Access Ports Figure 4-7: Top View of Electronics Bay... 65

8 Figure 4-8: Wing Dihedral Figure 4-9: Wing Locking Mechanism Figure 4-10: Tail Section Figure 4-11: Tail Section (Stowed)... 69

9 LIST OF TABLES Table 3-1: Rocket Budget Summary Table 3-2: Hardware Specifications Table 3-3: Launch Loading Table 3-4: Recovery Shock Calculations Table 3-5: Carbon Fiber Properties Table 3-6: Axial Stress Calculations Table 3-7: Buckling Calculations Table 3-8: Bending Calculations Table 3-9: Payload Bulkhead Bolt Shear Calculations Table 3-10: Threaded Rod Sizing Table 3-11: Motor Retention Sizing Table 3-12: Stringer Sizing Table 3-13: Parachute Descent Rates Table 3-14: Potential Rocket Failure Modes Table 3-15: Tool Use Injury Potentials and Mitigations Table 3-16: Tasking and Integration Schedule Table 5-1: Possible Launch Failure Modes Table 5-2: Launch Operations Risk Assessment Table 6-1: Funding Sources Table 6-2: System Budget Summary... 98

10 MIT Rocket Team, Massachusetts Institute of Technology Cambridge, MA Dr. Paulo Lozano Faculty Advisor 1 FRR SUMMARY 1.1 TEAM SUMMARY Andrew Wimmer Safety Officer, Rocket Owner, TRA # 9725 Level 3 awimmer@mit.edu John Kane Local NAR Contact kane@mit.edu 1.2 LAUNCH VEHICLE SUMMARY The purpose of the launch vehicle is to reach an apogee of 1 mile and deploy the UAV payload after descending to an altitude of 2500 feet. Diagrams of the vehicle are provided below in the rocket section. The carbon fiber airframe will be 140 inches long, the inner diameter of the rocket tube is designed to be 6 inches, and the outer diameter of the fins is inches. Furthermore, the mass of the rocket is projected to be 51.3 lbs, including UAV and the ballast placed at the motor mount as necessary in order to reach an apogee of 5280 feet using a single commercial Cesaroni L1115 motor from a 1.5 launch rail. Payload deployment will be performed at 2500 feet using two sabot halves that will be pulled out of the tube by the drogue parachute and separated using the deployable UAV wings. 1.3 PAYLOAD SUMMARY The rocket payload will consist of a 3.75 ft long, 7.5 pound UAV that will be launched from the rocket at an altitude of 2500 ft. It will fit inside the rocket by means of folding wings, tail, and propeller. The UAV will have a TR 35-30A 1700kv Brushless Outrunner motor onboard but will function as a glider for the majority of its flight. The UAV will fly to GPS coordinates supplied by a human operator. The UAV will not require advanced airplane or flight knowledge, which will make it useful for search and rescue type missions as well as for scientific research. The UAV will carry GPS tracking, airspeed sensors, atmospheric sensors, an accelerometer, a video capture device, and an onboard computer.

11 2 CHANGES SINCE CDR 2.1 LAUNCH VEHICLE CHANGES A few changes have been made to the rocket design since CDR in order to solve problems and improve the design. The changes revolve primarily around review of the design to simplify fabrication and integration. Airframe The length of the rocket was increased by 6 inches to 140 inches to eliminate the need for the sabot to neck down into the nose cone. Body tubes are no longer vacuum-bagged to simply the fabrication process. The fins were increased by 30% to provide additional dynamic stability Recovery Drogue parachute deployment charge size increased from 2.1 grams to 3.5 grams. Number of shear pins decreased from four to two. Drogue parachute diameter increased to 5ft due to desire to use a tangle-free nylamesh parachute. Main parachute diameter decreased to 12ft due to prolonged wait time on 14ft parachute. Deployment The sabot will no longer neck down into the nose cone, so as to simplify the manufacturing of the sabot and hard points. The sabot shell is made using a phenolic tube rather than a layer of fiberglass. The sabot was lengthened by 6 to allow for longer UAV wings. Propulsion The motor retention system has been modified to allow for ballast to be attached to the bottom of the rocket. Avionics No changes have been made to the launch vehicle avionics. 2.2 PAYLOAD CHANGES I. Wings a. Wingspan was increased to 61.4in. b. A dihedral latching mechanism has been added.

12 II. Tail a. The horizontal tail average chord was increased to 4.6in to increase stability. b. The vertical tail average chord was increased to 7.1in to increase stability. c. The taper on both stabilizers was eliminated to increase stability d. Tail will now be made of plywood to simplify the manufacturing process. III. Fuselage a. Electronics assemblies have been added directly in front of and behind the wings. b. Plastic bolts and nuts are being utilized in various nonstructural applications to reduce weight. c. Wing rotation mechanism s movement constrained to reduce possibility of failure. d. Fuselage length increased to 50in to accommodate the increased wingspan. 2.3 ACTIVITY PLAN CHANGES Since the completion of the critical design review, the team has completed educational outreach events at the Boston Museum of Science, and on MIT campus. The event at the Museum of Science was held on Saturday February 5 th. We set up three tables in the Museum s Current Science and Technology Center: one and a half for visitors to make their own Alka-Seltzer rockets, half for visitors to make their own parachutes, and one for us to set up our rockets and composites, where we answered questions and discussed our interests with anyone interested. We delivered two 20-minute talks: one in the morning, on rocket technology and history, and one in the afternoon, on the space industry. Due to museum rules, we were unfortunately unable to set off the rockets or drop the parachutes, but were informed that if we return in the spring, we will be able to do so outside. Our outreach event at the MIT Museum, which will take place on Sunday, May 1, will follow a similar format. We also taught two classes on MIT campus on Saturday March 13 th, both on Rockets and Composites, for a total of 23 high school students. 3 LAUNCH VEHICLE 3.1 TESTING AND DESIGN OF VEHICLE ROCKET DESIGN AND SUBSYSTEMS As described in the summary section, the purpose of the rocket is to reach 1 mile and deploy the UAV at an altitude of 2500 feet. This will be accomplished with a Cesaroni

13 L1115 motor and a 140 inch long, 6 inch diameter airframe. The UAV will be contained within a sabot, which will be located just aft of the nosecone. The drogue parachute will be above the sabot, the main parachute below the sabot, and the avionics below the recovery system. The overall rocket can be seen in Figure 3-1. FIGURE 3-1: OVERALL ROCKET Furthermore, the rocket budget summary (for mass and cost) can be seen in Table 3-1. Note that the mass does not include the ballast used to tune the rocket s apogee. TABLE 3-1: ROCKET BUDGET SUMMARY System Mass (kg) Cost (USD) Propulsion Rocket Airframe-Body Airframe-Fairing Avionics/Comm Payload Support Equipment Recovery SUBTOTAL The subsystems, which will be described in greater detail below, are: Airframe Recovery Deployment Propulsion Avionics/Communications SUBSYSTEM REQUIREMENTS AND DESCRIPTIONS Airframe The airframe is comprised of the following components:

14 Body Tube Nose Cone Fins Motor Retention System Recovery System Bulkhead Each of these will be described in detail below. The body tube is a carbon fiber laminate tube of inner diameter 6. The laminate is a 2- ply layup of Soller Composites 14.5 oz/sqyd biaxial sleeve carbon fiber fabric and Aeropoxy 2032/3665 matrix. Carbon fiber was chosen as the material for the primary structure due to its high strength-to-weight ratio, toughness, and ease of manufacture to customized shapes and dimensions. The biaxial sleeve was chosen due to difficulties in fabricating wrinkle-free tubes. All layups for the rocket are done in-house using a custom oven in the rocket team lab. For fabrication and transportation reasons, it would be difficult to make the entire tube in one segment. As a result, the body tube is split into 2 segments, with a seam just below the base of the sabot, as seen in Figure 3-2. The two segment lengths are for the upper segment and 48 for the lower segment. The seam between the tubes is accomplished by adding use of a phenolic coupler that is epoxied to the upper segment. The use of an internal coupler was accomplished by lengthening the upper segment. FIGURE 3-2: TUBE COUPLER SEGMENT Additionally, the tube will have 2 pressure relief holes (of 0.25 diameter, unless otherwise specified) in each of the following locations:

15 Just above the fins in the propulsion section Avionics bay: the hole for the switches will double as a pressure equalization hole In the middle of the section between the avionics bay and the sabot In the nose cone The nose cone Is a commercial PML fiberglass nosecone. A piece of standard phenolic tubing has been attached to the shoulder to allow the nosecone to fit properly in the tube. A beeline tracker is also integrated into the bulkhead in the nosecone FIGURE 3-3: NOSE CONE The nose cone is mounted to the body tube using 2 nylon 2-56 bolts (MMC 97263A077) that are threaded into the phenolic tube, which will act as shear pins. Bolts are used because they can be easily threaded into the shoulder during integration and will fail at low loading since they are plastic. Four fins were chosen with the dimensions as shown in Appendix 3 for rocket stability reasons (see Section 3.4). The fins are a carbon fiber, 3/16 MDF, carbon fiber sandwich laminate to maximize stiffness with minimum mass. The fins are located in position and angle relative to the rocket using slots that are laser-cut into the motor centering rings. MDF was chosen because it is available in uniformly flat sheets allowing for easier fabrication. Oversized slits added to the body tube to allow the fins to pass through, but provide no DOF restrictions. Fabrication of the fins is as follows: Sand the laser-cut MDF core edges (not tabs) to a taper Laminate the MDF core with a ply of carbon fiber on each face using standard plate lamination techniques (see manufacturing plan section) Obtain body tube with motor tube and centering rings installed Affix fins to the centering rings/motor tube assembly using 5 minute epoxy and let cure Apply another layer of carbon fiber across and between the fins, i.e. Tip-to-Tip The motor mount consists of a commercial 75mm motor tube and laser-cut, plywood centering rings. There are four centering rings in total, one located at each end of the motor tube and two in the middle, with four ¼-20 steel threaded rod stringers running through all four centering rings. The farthest forward is made from 1/2 plywood. The

16 farthest aft centering ring, which is not glued in place, is made from two rings of 3/16 plywood sandwiched together; the OD of the forward ring is the ID of the body tube, and the OD of the aft ring is the OD of the body tube. This transfers some of the thrust load through compression of the aft centering ring, rather than through shear in the epoxy joints holding the motor mount in the body tube. The middle centering rings are made from ½ plywood, with four slots to accept the fin tabs. The fin tabs are also slotted at the centering ring locations to allow the fins to contact the motor tube for additional support. One is located near the forward edge of the fin tabs and the other near the aft edge of the fin tabs, close to the aft-most centering ring. Plywood is chosen because it is relatively cheap, strong, light, and able to withstand the high temperatures of the motor casing without deforming. The four steel stringers are used to stiffen the motor mount and as a way to mount steel ballast plates to the aft of the rocket. The steel plates are cut with a water-jet from varying thicknesses of steel in a pattern that is able to interface with the rear of the motor mount. This will enable us to easily adjust the rocket s ballast on launch day by simply sliding off the aft centering ring, sliding on (or sliding off) some of the steel ballast plates, and bolting the aft centering ring on to the rear of the rocket. Motor retention will be accomplished as follows. Two 8-32 T-nuts will be mounted to the aft-most centering ring, 180 apart at a radius of roughly screws will go through two small clearance holes in the motor retention plate and screw into the T-nuts to hold the plate in place. The motor retention plate will be a piece of 1/32 steel sheet that has a hole cut in it; this hole will be made large enough for the motor s nozzle to fit through, but small enough to keep the motor casing from falling out of the motor tube. There is a thrust ring on our 75mm hardware that prevents the motor casing from moving forward during burn. The mounting and retention system can be seen in 4.

17 FIGURE 3-4: MOTOR CENTERING AND RETENTION The recovery system bulkhead serves as a reaction point for lateral forces from the payload, which come from two sources: inertial force of payload during boost, and drogue drag force between drogue parachute deployment and main parachute deployment. Axial forces will be reacted through a threaded rod to the motor casing, for reaction via the motor mount assembly. The bulkhead must also be removable to enable removal of the avionics bay, which sits between the motor retention bulkhead and the recovery system bulkhead. The threaded rod will be used to affix the payload support bulkhead to the airframe. A single bolt will be used to fix the rotational position of the avionics bay. The bulkhead will need to attach to both the charge released locking mechanism and the quick link to the main parachute shock cord. As a result, the bulkhead needs to have an eye bolt that is capable of transferring the loads to the bulkhead, which will be done through via an eyenut (MMC 3274T41) and threaded rod (MMC 95412A652). The design of the bulkhead is shown in Figure 3-5. Recovery FIGURE 3-5: RECOVERY SYSTEM BULKHEAD A detailed description of the recovery process can be found in the Section 3.2.

18 Deployment Deployment of the UAV and parachutes is as follows. Initially, the stacking of the rocket above the recovery system bulkhead is as follows (as seen in Figure 3-6 and Figure 3-7): Payload Bulkhead attachment quick links Charge released locking mechanism Sabot ejection charge Main parachute Sabot base hardpoint Sabot halves (cradling UAV) Sabot top hardpoint Drogue parachute quick link Drogue parachute Nose cone ejection charge Note: There is a redundant charge in the nose cone and a redundant igniter in the charge released locking mechanism. FIGURE 3-6: SABOT OVERVIEW FIGURE 3-7: PAYLOAD INTEGRATION STACKING (NOT TO SCALE) The deployment then occurs as follows:

19 Just after apogee, nose cone ejection charge fires Nose cone separates, but remains attached to the drogue parachute Drogue parachute deploys Rocket descends to 2500 feet At 2500 feet, the charge released locking mechanism fires. Mechanism to be used is the FruityChutes L2 Tender Descender, additionally a pair of redundant 0.5 gram charges will be fired at this point to ensure the sabot leaves the tube The drogue parachute pulls the sabot out of the rocket tube As the sabot leaves the tube, the spring-loaded UAV wings (which are restrained prior to the sabot leaving the body tube) push the sabot halves apart The sabot pulls the main parachute bag out behind it Main parachute deploys and remains attached to the main body tube After deployment, the rocket will fall to the ground in two sections, as shown in Figure 3-8: Sabot and nose cone, which are attached to the drogue parachute via the upper hardpoint and a shock cord Main body tube, which is attached to the main parachute via the recovery system bulkhead and a shock cord FIGURE 3-8: RECOVERY CONFIGURATION Deployment into two pieces (rather than one) is performed in order to minimize the chance of contact between the sabot/uav and the body tube after separation. This will

20 enable the drogue parachute to pull the UAV/sabot away from the rocket to allow clean separation and minimize the chances of entanglement. As described above, the UAV is encased within the two sabot halves, which are made of PML 6 Phenolic tube cut in half. Force will be transferred between the hardpoints using a 4x 10-24threaded rods (MMC 94435A355), which will mount to the upper and lower hardpoints using clearance holes and nuts. Finally, hardpoints are glued to the upper and lower ends of the sabot halves. These hardpoints enable recovery and deployment system fixtures to be attached to the sabot. One of these hardpoint sets is shown in Figure As can be seen below, the hardpoint halves utilize a brace to ensure that the halves remain together during deployment. Adhesive is applied as shown in the figure. An eye bolt is threaded into the lower hardpoint half, which serves as the attachment points for: Lower hardpoint: the charge released locking mechanism Upper hardpoint: the drogue parachute and upper shock cord (attaches to nose cone) It should be noted that the upper hardpoint will require eye bolts in both hard point halves due to ensure both sabot halves remain attached to the drogue parachute. Propulsion FIGURE 3-9: SABOT HARD POINT The rocket will be powered by a Cesaroni L1115 solid rocket motor. This motor was chosen because it is commercially available and does not require any modifications in order to reach the flight altitude requirement of 5280 feet based off the mass estimates available this early in the design process. The motor is actually more powerful than required given the current mass estimates, but this will ensure that even with mass creep over multiple design iterations, the rocket mass can be optimized with ballast weight to come as close to 5280 feet as the models can predict. The Cesaroni L1115 is also reloadable and relatively inexpensive compared to its Aerotech counterparts. It does not require extensive ground support equipment

21 compared to hybrid motors, which were originally considered for propulsion. The L1115 is 75mm in diameter, 24.5 inches in length, and has a total impulse of 4908 Newtonseconds over a 4.49 second burn time. For the full-scale test, the Cesaroni K1085 solid rocket motor was used. The K1085 has enough power to launch the full system up to an altitude of 2300 feet and still has the same diameter as the L1115, so minimal changes will have to be made to the motor housing section for the full scale test launch. The K1085 is 75mm in diameter, 13.8 inches long, and provides 2486 Newton-seconds of thrust over a 2.1 second burn time Avionics/Communications The purpose of the rocket avionics is to control parachute deployment while collecting rocket flight data and relaying it to the ground station. The rocket avionics system is comprised of two flight computers (minialt/wd and ARTS2) and an ARTS2 transmitter. The minialt/wd flight computer serves as a backup altimeter that measures the rockets altitude during launch and stores in on the computer board and will fire a redundant igniter for the recovery charge after the ARTS is programmed to. This data can be retrieved after rocket recovery where the minialt/wd flight computer is connected to the ground station computer via a minialt/wd to PC Connect Data Transfer Kit. The ARTS2 flight computer handles primary parachute deployment as well as determining the rocket state variables and flight states. The ARTS2 Transmitter transmits the data from the ARTS2 to the ground station receiver. Rocket Flight data includes: State Variables: o Altitude o Maximum Altitude o Velocity o Acceleration Flight State: o On Pad o Thrust o Coast o Apogee o Descent o Drogue parachute Deployment o Main parachute Deployment Power Supply Four 9 volt batteries will provide power for the flight computers and transmitters. One of the batteries will be dedicated towards powering the minialt/wd while the other three will power the ARTS2 flight computer and telemetry system to create a power source redundancy in case one was to fail. On the ARTS2 board one battery powers the two

22 systems while the other powers the igniters. They will be located inside the removable rocket avionics section of the rocket, alongside the rest of the avionics system. Hardware Description MiniAlt/WD Logging Dual Event Altimeter (PerfectFlite) This flight computer measures the rocket s altitude by sampling the surrounding air pressure relative to the ground level pressure. The altitude above the launch platform is calculated every 50 milliseconds. After launch, the device continuously collects data until landing. Altitude readings are stored in nonvolatile memory and can be downloaded to a computer through a serial data I/O connector. The minialt/wd has two channels for parachute deployment; one for the main parachute and the other for drogue parachute. FIGURE 3-10: MAWD FLIGHT COMPUTER Altimeter Recording and Telemetry System (ARTS2 Flight Computer) (Ozark Aerospace) This flight computer calculates the rockets altitude by sampling the surrounding air pressure relative to the ground level pressure and measuring the rockets acceleration. The rate at which the altitude above the launch platform is calculated is adjustable and will be set at 200 samples per second with an overall recording time of 82 seconds. Altitude readings are sent to the ground station via the ARTS2 telemetry transmitter. Also the altitude and other flight data are stored in nonvolatile memory to be downloaded to a computer through a serial data I/O connector. The ARTS2 has two channels for parachute deployment; one for the main parachute and the other for drogue parachute. 1. Terminal Connector 2. GPS Connector

23 3. Programming header 4. Battery Configuration 5. Main Battery Connection 6. Power Switch Connector 7. 9V Pyro Battery Connection 8. Option Switches 9. Output Channel Terminals. Channel 1 Apogee, Channel 2 Main. FIGURE 3-11: ARTS2 FLIGHT COMPUTER ARTS2 Telemetry Transmitter (Ozark Aerospace) 100mW 900MHz spread spectrum transmitter Integrated wire antenna on the board is connect to a larger antenna on the rocket Works with the ARTS-TT2-W and ARTS-TT2-RPSMA ARTS flight computer gets connected directly to the transmitter board Transmits real time flight data to the ARTS telemetry receiver FIGURE 3-12: ARTS2 TELEMETRY TRANSMITTER ARTS2 Telemetry Receiver (Ozark Aerospace) Receives telemetry from the ARTS2 transmitter and sends it to the computer Connect to the ground station computer via a serial cable

24 FIGURE 3-13: ARTS2 TELEMETRY RECEIVER TABLE 3-2: HARDWARE SPECIFICATIONS Hardware MiniAlt/WD ARTS2 ARTS2 Transmitter Operating Voltage 6-10 volts 9-25 volts volts Minimum Current 10 milliamps Dimensions 0.90 W, 3.00 L, 0.75 T 1.40"W, 3.75"L, 0.75"T ~2.50"W, ~7.50"L, ~1.50"T Weight 20 grams ~20 grams ~200 grams Altitude Accuracy +/-.5% Operating Temperature Maximum Altitude 0C to 70C 25,000 feet 100,000 feet 100,000 feet Switches Each altimeter has a dedicated toggle switch to provide power to each altimeter. These will be activated on the launch pad in the vertical position. Software Telemetry software: Displays flight data in real time using text, 2-D, and 3-D graphical user interfaces.

25 ARTS Software V1.61 (Data Analyzer): FIGURE 3-14: ARTS GUI Used in analyzing the data collected by the ARTS2 and also configuring parachute deployment and sample rate settings. Parachute Deployment FIGURE 3-15: ARTS DATA ANALYZER Both the ARTS2 and the minialt/wd are programmed to deploy the drogue parachute at apogee, while the main parachute and the UAV are set to deploy after apogee is reached at an altitude of 2500 feet. The MAWD will be set to back-up the main deployment at 1700 feet. This creates system redundancy in case one of the flight computers fails. Transmission from Rocket to Ground Station Since the carbon fiber material of the rocket body tube disrupts RF signals, the wire on the ARTS2 transmitter will be extended out of the avionics bay. The 14 gauge copper

26 wire is connected to the transmitter antenna via a binding post. It then extends from the avionics bay and warps around the lower section of the rocket body. Each loop of the helix is spaced 33 centimeters apart to prevent destructive interference. Kapton tape is placed above and below the wire to prevent contact with the carbon fiber. Based on the second full scale test launch, the antenna will not have sufficient range to transmit over the entire range of the flight. However since the transmitter board does not hinder our operation, the telemetry system shall be left in and a loss in transition will be accepted. Mounting/Placement Placed in the avionics bay, which is in the lower segment of the rocket as described below. The flight computers will be mounted in such as way so that their pressure and acceleration readings are not disturbed. This means that the barometer on both the ARTS2 and miniatl/wd would have to have at least a 1 centimeter clearance from any closest surface parallel to it. Also, the ARTS2 will be mounted with its length parallel to the rocket s length in order for the accelerometer to record proper positive values. The avionics will be mounted into an avionics integration tube, which is shown below in Figure 3-18.

27 FIGURE 3-16: AVIONICS PACKAGE As can be seen in the figure, the boards and battery are mounted to a plate, which will be mounted vertically in the rocket frame. A series of L brackets will be used to mount this vertical avionics plate to the avionics tube, which can be integrated with the rocket in a preassembled form. Additionally a Beeline tracker will be mounted to the side of the avionics assembly. The hole in the top of the avionics package is for wires to reach into the upper portion of the rocket. The switches on the bottom image are used to arm each of the following just before launch: ARTS2/ARTS2 Transmitter Power Switch MAWD Power Switch The bottom avionics plate is grooved so that the phenolic tube packaging shell can be attached. The boards and batteries can be mounted to the avionics plate, which is mounted to the top plate using 4x 6-32 fasteners. Nutplates are glued to the insides of each of the mounting brackets, such that bolts can be used without standard nuts. After insertion into the tube, four additional 6-32 fasteners are bolted from the bottom plate into the bottom L brackets, the holes of which are threaded. After this is assembled, the whole avionics package may be inserted into the rocket as described in the payload integration plan. This design was chosen to make the avionics assembly as modular as possible, while still maintaining access just before flight and low mass/cost of the assembly. [See FRR-Drawings document] DRAWINGS ANALYSIS RESULTS In order to verify the design of the rocket, a battery of analysis was applied to the rocket airframe, bulkheads, and mechanisms. The order of analysis is as follows: Define loading conditions Design part Use hand calculations to size the part Validate hand calculations using finite element method Re-size as necessary Loading Conditions Determining loading conditions for a vehicle that must withstand a variety of largely unknown dynamic and static loading is a difficult task. Furthermore, the rocket airframe can be significantly overdesigned without applying too significant of a penalty to the mass budget. As a result, the loading conditions were often estimated using significant margin to account for uncertainty. Launch Loading is summarized in Table 3-3.

28 TABLE 3-3: LAUNCH LOADING Launch Loading Aero Loading 90 lbf Peak Thrust 385 lbf Payload Mass 15 lbm Max Accel 8 G Total Axial 595 lbf MUF 1.5 Design Axial lbf Total Lateral 150 lbf MUF 1.5 Design Lateral 225 lbf Aerodynamic loading is determined from the Rocksim model, peak thrust is determined from the Cesaroni L1115 Thrust Curve, payload mass is determined from the UAV design and sabot, and maximum acceleration is determined from the Rocksim model. Although many of these peak loads are applied independently from each other, to provide for a conservative calculation, the loads are summed to create a total load, which is then margined by a 1.5 model uncertainty factor, resulting in a design axial load of 890 lbf. Lateral loading is determined by summing half of the aerodynamic and payload forces and margining by a 1.5 model uncertainty factor. This is assumed to be highly margined since as much as half of each of these loads is unlikely to be applied in the lateral direction. Regardless, the design lateral load is therefore 225 lbf. Recovery shock calculations are determined by examining the change in momentum of the rocket due to deployment, as shown in Table 3-4. TABLE 3-4: RECOVERY SHOCK CALCULATIONS Recovery Shock Calculations Initial Rate 64 ft/s Final Rate 18 ft/s G 32 ft/s^2 T 0.1 s Accel 460 ft/s^2 Gs Rocksim Gs 8 MUF 2 Design Gs Design Force lbf Recovery calculations show the descent rates of the system prior to deployment (under the drogue) to be 64 ft/s and of the main rocket after deployment (under the main) to be 18 ft/s. Assuming a deployment time of 0.1s, this results in 14 Gs. Adding a model uncertainty factor of 2 to this results in 29 Gs, which (given the mass of the payload system) results in a design recovery shock force calculation of 430 lbf.

29 Body Tube Analysis As described in the design section, the body tube is made from Soller Composites Biaxial Weave carbon fiber. The modulus and strength are taken from Soller Composites and the resulting strain allowable is derived, as shown in Table TABLE 3-5: CARBON FIBER PROPERTIES Material Properties E psi V 0.3 E (claimed) 34 Msi E_lam (claimed) Msi Strength 110 ksi Strain strain FOS 3 strain w/mos µstrain Using these properties, hand calculations could be performed for axial compression, global buckling, and bending. These intermediate calculations as well as the resulting margins of safety can be seen in Table 3-6 through Table 3-8. It should be noted that, for the sake of being conservative, lateral loads are taken to be applied at the top of the rocket and restrained at the base. TABLE 3-6: AXIAL STRESS CALCULATIONS Axial Stress Calculations ID 6 in OD in Area In^2 Axial Stress psi Strain strain µstrain MOS TABLE 3-7: BUCKLING CALCULATIONS Buckling r/t Z Kc 4000 Fcr psi L/r Axial Load lbf MoS 2

30 Axial Load lbf Allowable MOS TABLE 3-8: BENDING CALCULATIONS Bending I in^4 Z 3.75 in M in-lbf Stress psi Strain strain µstrain MOS In order to verify these calculations, a finite element model was developed in Femap. The applied boundary conditions of the axial case are shown in Figure The boundary conditions in the lateral load case are shown in Figure FIGURE 3-17: AXIAL CASE BCS FIGURE 3-18: LATERAL CASE BCS This model was then solved using NEi Nastran, resulting in the ply 1 effective strain and displacement outputs for the axial case, as shown in Figure The results for the lateral case are shown in Figure 3-20.

31 FIGURE 3-19: AXIAL CASE RESULTS

32 Payload Bulkhead Analysis FIGURE 3-20: LATERAL CASE RESULTS As described in the design part of the document, 4x #6 bolts are used in order to fasten the tube segments together. From this, the effective strain may be determined and compared to allowables. This calculation takes into account the allowable shear area, using a 4 ply thick doubler and assuming that only two of the bolts are being used, as shown in Table 3-9. TABLE 3-9: PAYLOAD BULKHEAD BOLT SHEAR CALCULATIONS

33 P/L Bulkhead Bolt Hole Shear N 4 n "used" 2 hole dia in T in shear area in^2 Stress psi Strain strain µstrain MoS µstrain The 3/8 threaded rod is used to connect the parachute shock loads to the motor retention system, as shown in Table Motor Retention Analysis TABLE 3-10: THREADED ROD SIZING Threaded Rod Sizing rod dia in A_tot in^2 Stress psi allowable psi stress MoS Since the threaded rod transfers shock load to the motor retention system, the motor retention system must be able to react the load of the deployment. It must also, however, be able to react the original axial load. The initial sizing calculations are as shown in Table TABLE 3-11: MOTOR RETENTION SIZING Motor Retention Sizing # Used 3 Thk 0.5 in ID 3 in OD 6 in avg D 4.5 in "b" in I in^4 M in-lbf Z 0.25 in Stress psi failure 1500 psi stress MoS This was validated using Ansys Workbench. The loading conditions are shown in Figure 3-21.

34 FIGURE 3-21: MOTOR RETENTION BCS The resulting displacement is shown in Figure 3-22 and the stress is shown in Figure 3-23 FIGURE 3-22: MOTOR RETENTION DISPLACEMENT

35 FIGURE 3-23: MOTOR RETENTION STRESS Payload Support Equipment Sabot Stringer Sizing As described in the design portion of the document, the stringers are 4x #10 rods that connect the top and bottom parts of the sabot together. The size of these aluminumthreaded rods can be verified as shown in Table TABLE 3-12: STRINGER SIZING Stringer Sizing stringer dia in # 4 A_tot in^2 Stress psi Rho lb/in^3 Length 50 in sabot dia 6.5 in Mass lbm allowable psi stress MoS COMPONENT, FUNCTIONAL, AND STATIC TESTING Functional tests to be performed are several deployment and recovery tests, which are completed after UAV prototype completion. Deployment altitude will be verified using barometric testing. The team has constructed a small vacuum chamber, which is capable of roughly simulating

36 ambient pressure. As a result, the avionics package was placed into the vacuum chamber to ensure that it sends charge ignition commands at the right times. In order to verify the failure force of the shear pins, the first version of the rocket airframe is used with a representative nose cone, with the shear pins mounted in flight orientation. The black powder charge will be ignited at the closed end to validate the mass of black powder to be used. This testing is described further in Section UAV deployment will also require testing, which is performed in a couple of phases: (1) the force of the drogue parachute on the sabot is simulated to ensure that the sabot separates from the tube and the UAV deploys and (2) integrated deployment tests from a balloon platform. This test will be described further in Section 4.4. A series of avionics tests will also be performed. A summary of the tests is provided below. Greater detail can be found in Section 4.4. The emergency locator beacons (transmitters and receiver) operation has been checked, by searching for the beacons in a representative location and for recovering the test rocket. Each computer is also checked to ensure that they downlink properly to the ground station. This is performed on the ground in a field and then on a balloon platform using a representative ground station and rocket. Finally, these tests culminate in a representative scaled test launch, which verifies the functionality of all systems, including the UAV. Discussion of Full Scale Test Flight Results February 20, 2011: The team travelled to Vermont to launch with the Champlain Region Model Rocket Club. A K1085 was used for the flight. Integration took longer than expected due to the cold and the tight fit and perfect alignment needed to install screws. The rocket weighed 40.1 pounds on the pad and was flown with a 7 pound UAV stand in that was to deploy a parachute. Winds were 7-10mph and the temperature was 25 degrees F. The K1085 took approximately 4 seconds to ignite, and lifted off, arcing slightly into the wind. Apogee was at 2344 according to the MAWD, 2190 according to the barometric sensor on the ARTS2 and 2070 according to the accelerometer on the ARTS2. At apogee, the nose cone was ejected and the drogue parachute deployed. The rocket fell through the set 1300 and 1000 backup altitude without a main parachute. The rocket impacted the field at 51 feet/second and did not sustain any damage. Upon post-mortem inspection, it was found that the coupler/shoulder on the nosecone came unglued from the nose cone itself leaving the shoulder still shear-pinned to the body tube. This prevented the sabot from sliding out upon main deployment.

37 March 20, 2011 The team intended to return to Vermont, however, muddy field conditions caused the launch to be postponed, requiring the team to travel to an MDRA launch in Maryland. The motor for this flight was the L1115 we intend fly in Huntsville. Liftoff weight was 43.8 pounds, and the UAV airframe was flown with stand-in wings and a parachute again. Winds were 10-13mph and the temperature was 50 degrees F. The motor took approximately 2 seconds to ignite, which was longer than expected given the improved ignition system. The rocket took off with no visible arcing. About feet off the ground, the rocket started coning, and continued to until it had damped itself out shortly after the motor burned out. The rocket reached 5604 according to the MAWD, 5591 according to the barometer on the ARTS2 and 5998 according to the accelerometer on the ARTS. At 1300, the release mechanism fired as planned, however, the sabot did not deploy at this time. A.5 gram charge, along with a backup igniter in the release mechanism fired at 1000 and helped push/pull the sabot out. The sabot cleanly deployed the UAV, its parachute and the 12 R12 main parachute. The UAV wings unfolded properly as well. The pieces were later recovered, with the sabot, nosecone and drogue within 200 of the main body second. The sabot, nose cone and drogue landed a few seconds prior to the main second, but their descent rate was well within the defined limits. With the exception of the coning on ascent and a slight crack in the phenolic making up the sabot that occurred on landing, the flight went exactly as planned. We plan to use composites to reinforce the inside of the sabot to avoid damaging it upon landing again. More simulations will be performed, however, it seems at the moment that in order to reduce the effects of coning, either larger fins are needed or a lower moment of inertia around the x and y axis. Given the difficulty of reducing the moment of inertia, larger fins will likely be added. Additionally, we will install a backup.5 gram charge connected to the backup altimeter to push the sabot out. We only flew with one charge on this flight in order to test if it was actually needed. Overall, the test flights proved that the rocket configuration we are planning to fly in Huntsville works as designed and within safe limits of high powered rocketry. 3.2 SAFETY AND FAILURE ANALYSIS See Section for a detailed table of launch vehicle and recovery system failure modes. 3.3 RECOVERY SUBSYSTEM PARACHUTE CHOICE

38 When the drogue parachute is deployed at apogee, it will need to support a total system mass of 24 kg. A 5ft diameter parachute will be used to achieve a descent rate of 57 ft/s. Once an altitude of 2500 ft AGL is reached, the tether securing the sabot inside the rocket will release, allowing the drogue parachute to pull the sabot and the main parachute out of the rocket. At this point, the rocket body will separate from the sabot/nose/drogue section and free fall as the main parachute deploys. This will allow for a considerable gap between the rocket body and the sabot, decreasing the risk of the deployed UAV colliding with the rocket or becoming entangled in the main parachute. With the UAV deployed and the sabot separated from the rocket body, the remaining structure has a mass of 17 kg. With a 12ft diameter parachute, a final descent rate of 21 ft/s can be achieved. Under the 4ft parachute, the nose cone and sabot will have a final descent rate of 21.5 ft/s. TABLE 3-13: PARACHUTE DESCENT RATES Final Descent Rate System Under Drogue Nose/Sabot Final Descent Rate Rocket Body Under Main 54 ft/s 21.5 ft/s 21 ft/s The drogue parachute and nose cone are directly connected to the sabot. This assembly is initially connected to the recovery system bulkhead via the explosive tether. The main parachute is also secured directly to the recovery system bulkhead (not by the tether). Its deployment is constrained by the sabot. The charge release mechanism will contain 0.2 grams of black powder. This number is recommended by the manufacturer. 1 The calculations for the amount of black powder required to successfully separate the nose cone from the body tube can be found below. The drogue deployment charge must provide ample force to break the shear pins, accelerate the nose cone away from the rocket body, and accelerate the drogue parachute out of the nose cone. Four #2-56 nylon screws (MMC 94735A177) will be used as shear pins to retain the nose cone. Nylon 6/6 has a shear strength of 10ksi. 2 With this, the maximum shear force can then be calculated by the following equation:, 1 Tender Descender User s Guide, 2

39 where A is the cross-sectional area of the bolt, and τ is the shear strength. For a #2-56 screw, the minimum pitch diameter is in. 3 This leads to a shear force of 40 lbf. With four pins, the charge will have to provide a minimum force of 120 lbf. Adding 25% margin, the charge will need to provide a total force of 150 lbf. This leads to a required black powder mass of 2.1 g TEST RESULTS FROM EJECTION CHARGES AND ELECTRONICS While preliminary static recovery testing indicated the calculated amount of black powder was sufficient to overcome the restraint force of the shear pins, results from our first full-scale test launch and subsequent static testing using flight hardware indicated otherwise. Additional rounds of testing were performed with different combinations of charge size and number of shear pins. A charge containing 4.2 grams of black powder proved sufficient to overcome the force of four shear pins, but required a relatively large container and was deemed unsafe. A charge of 2.1 grams was tested with three shear pins, but was still unable to separate the nose cone from the body tube. After additional testing, it was decided that a charge containing 3.5 grams of black powder and two shear pins would be used. Testing of the Tender Descender has confirmed that the manufacturer s recommended charge size of 0.2 grams is sufficient to operate the device FAILURE MODES AND SAFETY ANALYSIS See Section for a detailed table of launch vehicle and recovery system failure modes. Flight Profile Predictions 3.4 MISSION PERFORMANCE PREDICTIONS FLIGHT PROFILE SIMULATION For the Flight Readiness Review flight profile simulations, the RockSim 8 model used in previous reviews was updated due to changes in the structure, specifically the overall lengthening of the body tube and the associated mass difference. The RockSim model mass was verified against the Solidworks model and the actual rocket used in the scale test. Parachute descent rates were verified against the MATLAB parachute sizing model. The RockSim model agreed with the Solidworks model mass to within 0.05 kg Black Powder Pressure-Force Calculator: Pressure_Force_Calculator_Ver2.xls

40 and with the MATLAB model descent rates to within 3 feet per second. Figure 3-24 shows the RockSim model. FIGURE 3-24: PREDICTED CM AND CP LOCATIONS A battery of simulations was run, taking into account the approximate location and altitude of the launch site and average temperature, pressure, and humidity conditions. It was known that the Cesaroni L1115 would be more powerful than necessary and propel the rocket higher than the target altitude. Figure 3-25 shows the thrust curve of the L1115. FIGURE 3-25: CESARONI L1115 THRUST CURVE With no added ballast or winds, the rocket flew over 2000 feet above the target altitude. This was expected and desired, especially considering the mass margin of the payload and other components, the masses of which have only been measured up to this point. Initially, the RockSim model had a mass of kg and an initial stability margin of Groups of ten simulations were run to find an optimal mass of the ballast that needed to be added. Each simulation had variable light winds (3-7 mph), and the ballast mass and launch rail angle were varied until the desired apogee and landing range distance were achieved. The optimal value for the ballast weight added to the bottom motor bulkhead mount is 5.05kg, giving the rocket a total wet mass of kg. This gives an average

41 altitude over 10 simulations of 5266 feet (maximum 5280 feet, minimum 5255 feet) and a distance at landing of no more than 500 feet from the launch location, with an average distance of 100 feet. The center of gravity (CG) and the center of pressure (CP) in this configuration are and inches, respectively, from the tip of the nosecone. The reason a lower altitude was chosen as the target value instead of 5280 feet exactly was because the variance in altitude in the simulations will cost more if the rocket overshoots the target altitude. 2 points are taken off the final score for every foot above the target altitude the rocket achieves but only 1 point is taken off for every foot below the target the rocket flies. At t = 0, the Cesaroni L1115 is ignited. Burnout occurs at 4.49s, and apogee occurs at approximately 19.4 seconds. At this time, the first charge is ignited to eject the nosecone and deploy the drogue chute, which pulls the sabot out of the rocket. At an altitude of 2500 feet, the second charge is ignited. This charge releases the UAV from the sabot, separates the nosecone, drogue chute, and sabot from the rest of the rocket body tube, and deploys the main parachute. Figure 3-20 shows the acceleration and velocity of the rocket during the first 30 seconds of flight (the remaining flight time was omitted for clarity). The maximum speed occurs near burnout, and does not exceed Mach 0.5. The maximum predicted acceleration occurs at the parachute deployment, as expected. While the magnitude of the maximum acceleration is high compared to what was expected, this is still within the range that the carbon fiber structure of the rocket can stand. An initial concern was that the parachute cords could rip the body tube apart during high-speed deployment. Future modeling will try to reconcile the nearly instantaneous parachute deployment featured in RockSim and the expected unraveling time of the chute to prevent such high accelerations in simulations.

42 FIGURE 3-26: PREDICTED ACCELERATION AND VELOCITY PROFILES Figure 3-26 shows the simulated altitude profile of the rocket. Burnout and apogee are shown with red and blue dotted lines, respectively, and the main parachute deployment can be seen as the kink in the altitude line near 50 s.

43 FIGURE 3-27: SIMULATED ALTITUDE PROFILE More flight profile modeling at the time of launch will more accurately define the launch conditions, including launch pad altitude, actual weather conditions (relative humidity, average wind speed, etc.), and competition settings. Immediately before the flight, these conditions will be taken into account and the mass of the ballast will be adjusted according to on-site measurements of the center of gravity to have as inputs to the simulations to achieve the predicted altitude. Altitude Sensitivity Analysis The Cesaroni L1115 was chosen because the final altitude would be less sensitive to errors in the added ballast mass compared to smaller motors. However, the final altitude is still susceptible to small changes in the ballast mass. Two additional 10-flight simulations were run with a +/ kg change in the ballast mass. The average was -19/+23 feet lower/higher than the baseline ballast (5.05 kg), suggesting that near the baseline ballast, the dual variable for change in altitude with

44 respect to ballast mass is 420 feet per kilogram for small perturbations in the ballast mass. This suggests that the team must carefully weigh the rocket and determine the mass immediately before flight so that the ballast can be estimated to the highest degree of accuracy to achieve the optimal flight. Flight Test Data Versus Simulations The scale motor flight simulated with the K1085 had simulation results that closely match the actual flight parameters seen after flight. Analysis of the full scale launch with the L1115 on March 20 th has not yet been completed due to the complexity of attempting to simulate the coning seen during ascent and its effects on the final altitude; however, some basic analysis has been done comparing the test flight to the test RockSim model. The launch configuration used in the full-scale test flight did not have any ballast added to the back end of the rocket, which increased the static margin to This is highly overstable and makes the rocket susceptible to weather-cocking even in light winds. According to RockSim model, the rocket should have reached nearly 7400 feet. However, the rocket only reached an altitude 5604 feet. The most significant difference between the test flight hardware and the competition flight hardware was the construction of the lower body tube. The phenolic tubing used to construct the test flight body tube was not long enough, so the final product was bent considerably. Because of this, the thrust error angle causing coning during flight. This would have reduced the final altitude significantly, which agrees with the ~1800 ft difference in altitude between the test flight and the RockSim model. Figure 3-28 shows the contrail of the test flight.

45 FIGURE 3-28: CONTRAIL OF THE FULL SCALE TEST FLIGHT SHOWING CONING DURING FLIGHT Further analysis on the flight profile models between the test flight and the competition flight show that the overstable configuration would lead to higher angular accelerations due to winds. Figure 3-29 shows a plot of the angular accelerations of the unballasted and ballasted launch configurations. The unballasted configuration experiences a maximum angular acceleration that is over 5 times as much as the ballasted configuration, and the period of oscillation is smaller. This effect, coupled with the high error in the thrust angle, caused the full-scale test rocket to oscillate during flight and achieve a much lower altitude that predicted. FIGURE 3-29: COMPARISON BETWEEN THE MODEL ANGULAR ACCELERATIONS OF THE TEST FLIGHT CONFIGURATION (UNBALLASTED, LEFT) AND THE COMPETITION CONFIGURATION (BALLASTED, RIGHT) Even though the static stability of the ballasted configuration is less than the unballasted configuration, the dynamic stability is higher 3.5 SAFETY AND ENVIRONMENT IDENTIFICATION OF SAFETY OFFICERS Andrew Wimmer will be the primary rocket safety officer for the team. Ben Corbin is the team s MIT EHS representative and is the assistant safety officer and is in charge of safety issues not directly related to the rocket. Both team members have considerable experience in their respective areas SAFTEY AND FAILURE MODES ANALYSIS

46 The following table provides an updated analysis of the failure modes of the proposed launch vehicle and payload integration. See Section 5.2 for Safety and Failure Modes Analysis for Launch Operations. TABLE 3-14: POTENTIAL ROCKET FAILURE MODES Risk Likelihood Effects Risk Mitigation Plan Failure to integrate vehicle in allotted 4 hour time period Sabot does not slide freely into rocket at time of integration The parachute bag is unable to slide in/out of rocket freely at time of integration Lost essential hardware (e.g. shear pins, screws, batteries, etc,) General payload integration failure Motor fails to ignite Catastrophe at Take-Off Motor failure Low Low Low Medium Low Low Low Low Loss of launch opportunity Loss of integration time Loss of integration time Loss of integration time while looking for the missing parts; possible mission failure Mission failure Unable to launch; minor mission setback Total mission failure; possibly spectator injury Total mission failure; possible spectator injury Practice integration under time constraints. Test fit before launch day. Have sandpaper on hand at launch. Repack parachute or find what the bag is handing-up on and fix Bring spare hardware Practice integration before launch day Replace igniter To mitigate this risk, we have detailed setup, integration, and launch procedures. We will conduct safety checks at every stage to ensure adherence to all safety guidelines. Follow proper launch safety distances. Store and assemble motor in accordance with

47 Structural failure Unstable launch vehicle Low Low Total mission failure; possible spectator injury Possible mission failure; possible spectator injury Birdstrike Low Flight path altered Lack of failure of shear pins General recovery system deployment failure Sabot not deploying Low Low Medium No parachute deployment; catastrophic failure; possible spectator injury Partial to total mission failure; possible spectator injury Payload not deployed; main manufacturer s instructions. Large safety factors accounted for during the design process reduce the impact that launch loads will have on weaker structural areas. Use standard construction procedures for LII-LIII rockets. Conduct structural testing. Spectators should always keep visual on rocket. Confirm vehicle stability before launch. Ensure actual CG position is acceptable relative to calculated CP. Spectators should always keep visual on rocket. Follow all NAR launch rules, holding launch if any wildlife overhead. Extensive deployment testing will be conducted to validate the amount of black powder being used for deployment is sufficient to break pins. Spectators should always keep visual on rocket. Extensive testing of all recovery subsystems. Follow extensive launch ops check lists. Make sure to arm altimeters. Spectators should always keep visual on rocket. Extensive testing. The wing release locking

48 Drogue parachute not deploying Entanglement of main parachute Recovery System Structural Failure (bulkheads, attachment points, shockcords, etc) Low Medium Medium parachute not deployed; possible spectator injury No force available to pull sabot and main parachute from rocket body; catastrophic failure; possible spectator injury Partial mission failure. Payload deployment still viable; main rocket body damage; possible spectator injury Partial mission failure. Payload deployment still viable; main rocket body damage; possible spectator injury mechanism will keep the UAV wings locked until sabot exits body tube. This will prevent premature opening of the sabot, decreasing the possibility of the sabot binding inside the rocket body. Spectators should always keep visual on rocket. Extensive deployment testing will be conducted to find optimal packing method for drogue parachute. Spectators should always keep visual on rocket. Parachute will be properly packed in deployment bag. Spectators should always keep visual on rocket sections. Extensive testing of recovery system structures to ensure their ability to meet strength requirements. Spectators should always keep visual on rocket. Extensive testing to ensure wing rotator UAV unable to locking mechanism Sabot fails to deploy; upper disengages after sabot open after section decent rate Low exits rocket body, and ejection from outside of allowable that spring force of the rocket window; mission wing rotator is sufficient failure to separate sabot halves. Failure of Low Loss of points and Use multiple simulation

49 vehicle to reach desired altitude Recovery device deployment on ground Brush Fire Low Low potential loss of science value Property damage; partial mission failure; possible injury Fire damage; possible spectator injury programs and data from actual flight tests to fine tune rocket mass and motor selection Avoid placing body in path of parts if electronics are armed. Wear safety glasses when necessary. Shunt charges until they are attached to recovery electronics. Do not move the rocket with armed electronics. Have fire protection equipment (and personnel trained in its use) onsite, as per NAR regulations. Follow NFPA table for dry brush around pad area POTENTIAL HAZARDS A listing of personnel hazards and evidence of understanding of safety hazards is provided in the sections below. TABLE 3-15: TOOL USE INJURY POTENTIALS AND MITIGATIONS Tool: Injury Potential: Risk mitigation procedure: Electric Handheld Sander Burns, cuts, skin abrasion Avoid loose clothing Rotary Cutter/Dremel Soldering Iron Handsaw Cuts, skin abrasion, eye damage from flying debris, respiratory damage from dust Burns Cuts, splinters, skin abrasion Always wear a mask, gloves, and safety goggles when operating Exhibit care not to come in contact with hot element Wear proper safety gear (gloves and goggles)

50 Table Saw Wood Lathe Table Router Drill Press Miter Saw Band Saw Belt Sander CNC Water Cutter CNC Laser Cutter Mill Metal Lathe Cuts, Limb/appendage removal Cuts, broken appendages Cuts, Limb/appendage removal Cuts, abrasion, loss of limbs/ appendages Cuts, Limb/appendage removal Cuts, loss of limbs/appendages Burns, skin abrasion Cuts, loss of limbs/appendages Burns, eye damage, respiratory issues, poisonous off-gassing Loss of limbs, scarring, eye damage from flying chips Loss of limbs, cuts, eye damage from flying chips Avoid loose clothing, follow safety procedures found in instruction manual. Avoid loose clothing, use proper tools and safety equipment Use proper protective gear. Use proper protective gear, hold down work with clamps Avoid loose clothing, follow safety procedures found in instruction manual. Use proper protective gear. No loose clothes, wear proper protective gear (gloves) Only trained personnel use this tool. Oversight by shop manager A training course is required by MIT to use the laser cutters on campus. Many safety measures exist, including ventilation and failsafe switches. A training course is required by all shop managers to use mills. Safety goggles are always worn. A training course is required by all shop managers to use lathes. Safety goggles are always worn. Safety Codes

51 The Tripoli Rocketry Association and the National Association of Rocketry have adopted NFPA 1127 as their safety code for all rocket operations. A general knowledge of these codes is needed and will be required by all team members. These codes are found in the Appendix document. Hazards Recognition The Hazards Recognition Briefing PowerPoint Presentation was given prior to commencing rocket construction, and will be given multiple times throughout the next few weeks. It covers accident avoidance and hazard recognition techniques, as well as general safety. 1) General a) Always ask a knowledgeable member of the team if unsure about: i) Equipment ii) Tools iii) Procedures iv) Materials Handling v) Other concerns b) Be cognizant of your own actions and those of others i) Point out risks and mitigate them ii) Review procedures and relevant MSDS before commencing potentially hazardous actions c) Safety Equipment i) Only close-toed shoes may be worn in lab ii) Always wear goggles where applicable iii) Always use breathing equipment, i.e. face masks, respirators, etc, where applicable iv) Always wear gloves where applicable, e.g. when handling epoxy and other chemicals 2) Chemicals a) The following are risks of chemical handling: i) Irritation of skin, eyes, and respiratory system from contact and/or inhalation of hazardous fumes. ii) Secondary exposure from chemical spills iii) Destruction of lab space b) Ways to mitigate these risks: i) Whenever using chemicals, refer to MSDS sheets for proper handling ii) Always wear appropriate safety gear iii) Keep work stations clean iv) Keep ventilation pathways clear v) Always wear appropriate clothing 3) Equipment and Tools a) The following are risks of equipment and tool handling: i) Cuts

52 ii) Burning iii) General injury b) Ways to mitigate these risks: i) Always wear appropriate clothing, e.g. closed-toed shoes. ii) Always wear appropriate safety equipment iii) Always ask if unsure iv) Err on the side of caution 4) Composites Safety a) Carbon fiber, fiberglass, epoxy, and other composite materials require special care when handling. b) The following are risks composites handling: i) Respiratory irritation ii) Skin irritation iii) Eye irritation iv) Splinters v) Secondary exposure c) Ways to mitigate these risks: i) Always wear face masks/respirators when sanding, cutting, grinding, etc., layups. ii) Always wear gloves when handling pre-cured composites iii) Always wear puncture-resistant gloves when handling potentially sharp composites iv) A dust-room has been constructed, as per MIT EHS guidelines, specifically for the handling of composite materials. d) No team member will handle carbon fiber until properly trained ENVIRONMENTAL CONCERNS All waste materials will be disposed of using proper trash receptacles Non-polluting recovery system heat protection will be used Solid rocket motor manufacturers instructions will be followed when disposing of any rocket motor parts Consideration of environmental ramifications will be made regarding applicable activities Proper blast shields on the launch pad will be used to prevent direct infringement of rocket motor exhaust on the ground Waste receptacles (trash bags) will be available for use around the prep area to encourage proper disposal of waste from rocket prep activities The following list of materials have been identified as potentially hazardous: o Aeropoxy 2032 Epoxy Resin o Aeropoxy 3660 Hardener o Ammonium Perchlorate Composite Propellant o Black Powder

53 See FRR-MSDS document for complete MSDS specifications on these materials. Integrating the UAV into the sabot: 3.6 PAYLOAD INTEGRATION FEASIBILITY DESCRIBE INTEGRATION PLAN The following steps should be completed after the parts of the Pre-Flight Checklist before this step have been completed. In other words, the UAV should be completely ready to be deployed before integrating it into the sabot. Separate the sabot halves and lay them next to each other. Set the UAV on a sturdy working surface While bracing the body, carefully fold the UAV s wings back, making sure to depress the rotating mechanism such that one wing folds under the other. Fold the three spring-hinged tail surfaces. Place the UAV in one sabot half. Insert the activation pin that pulls out at time of deployment. Perform one final look-over to make sure everything seems ok. While holding the tail surfaces so they do not spring out, place the UAV in one of the sabot halves. Carefully close the sabot with the other half. Use rubber bands to hold the sabot shut until it is ready to be integrated into the rocket. 3.7 PAYLOAD INTEGRATION PLAN INSTALLATION AND REMOVAL, DIMENSIONS, PRECISION FIT 1) Integrate Avionics Bay* a) Integrate the altimeters and tracker b) Integrate 4 New Batteries c) Test electronics (turn on and off) d) Turn on tracker e) Attach avionics plate to bottom cap with L-brackets and 6-32 bolts, which are inserted, from the outside, through nuts on the inside f) Wire ejection charge wires through upper avionics plate g) Attach avionics plate to top cap with L-brackets and 6-32 bolts, which are inserted from the outside through nutplates on the insides h) Insert threaded rod through avionics bay i) Slide assembly into tube j) Check all connections

54 k) Check pressure holes l) Install motor into motor mount tube and screw into threaded rod to hold avionics bay in FIGURE 3-22: AVIONICS ASSEMBLY ON PLATE INSIDE TUBE, ATTACHED WITH L-BRACKETS 2) Make Black Powder Ejection Charges and assemble Tender Descender (Safety Officer will oversee this step) 3) Integrate antenna* a) Antenna is pre-attached to main body b) Connect the antenna (14 gauge insulated copper wire) to the connector on the avionics bay FIGURE 3-30: TUBE-TUBE INTERFACE

55 4) Integrate rocket body with sabot/uav assembly by sliding the sabot into the upper tube, ensuring the proper end is up 5) Recovery* a) Attach Tender Descender to ejection charge wires from avionics bay b) Attach Tender Descender shock cord to the upper Tender Descender quick link c) Attach the lower Tender Descender quick link to the avionics bay eye-nut d) Attach main parachute shock cord to avionics bay eye-nut e) Place main parachute deployment bag in lower tube f) Attach deployment bag line to Tender Descender shock cord 6) Nose Cone a) Turn on and install tracker* b) Attach parachute to shock cord with a Girth Hitch* c) Attach shock cord to nose cone with bowline knot* d) Attach shock cord to the top of the sabot hardpoints with water knot e) Attach ejection charges to wires on sabot f) Place ejection charges in nose g) Fold and pack parachute in nose h) Install nose on upper tube i) Install shear pins 7) Tube integration a) Ensure that the quick link and ejection charge wires in the lower tube are easily accessible b) Slide upper segment on to lower segment c) Install 4x 6-32 alignment bolts d) Connect the two (2) ejection charge wires to the sabot e) Connect the quick link to the Tender Descender and the deployment bag to the lower sabot hard point f) Install detachable door with 12x 6-32 bolts *These items may be completed prior to arrival at the launch site to reduce prep time

56 FIGURE 3-31: MAIN PARACHUTE/SHOCK CORD ATTACHED TO EYE BOLT AND RECOVERY SYSTEM BULKHEAD FIGURE 3-32: INTEGRATED AVIONICS ASSEMBLY, MAIN PARACHUTE, SABOT AND UAV ASSEMBLY TASKING & INTEGRATION SCHEDULE TABLE 3-16: TASKING AND INTEGRATION SCHEDULE

57 Overall Task Number of People* Time Integrate Avionics Assembly and motor 3 15 minutes Assemble UAV 2 15 minutes Integrate main parachute 2 5 minutes Integrate UAV assembly and nose cone 3 10 minutes Integrate two tubes together 2 5 minutes Total time: Approximately 60 minutes *This includes one person with the checklist who will be supervising 4 PAYLOAD 4.1 EXPERIMENT CONCEPT The idea of a deploying an Unmanned Aerial Vehicle (UAV) with a rocket is not an entirely original idea; however, the end goal of producing a simplified flight control interface is a new idea. Current UAV technology requires a classically trained pilot to remotely fly the craft, or at the very least requires operators to undergo a large amount of training in the operation of remote controlled equipment. The control interface that the MIT Rocket Team is developing aims to reduce the amount of training required to successfully complete a UAV mission, opening this class of technology to a greater range of potential users. As such, a simple point-and-click flight system is being developed to easily translate user operations into functional flight controls, allowing for successful operation with little or no operator training. Furthermore, by choosing a rocket deployment, and keeping to a $5000 budget, it further allows for this technology to be applied to situations where time and budget are controlling factors. This quick deployment and relative low cost of operation would ideally suit the needs of search and rescue operations, reconnaissance missions, and even rapid scientific data gathering missions. By completing NASA s Science Mission Directorate, the MIT Rocket Team is further proving the range of applications the UAV is capable of completing. The requirements of the SMD do not explicitly require the complexities of a UAV. However, by using the UAV form factor, the MIT Rocket Team will be able to complete all required tasks of the SMD mission in a more precise manner over several missions, and will also have the capability to loiter in any airspace, being limited only by the charge on the batteries. payload. For example, during flight, a UAV will generally maintain the same orientation with respect to the horizon, allowing for all images taken during and after the flight to keep the sky and ground in the same location with a very low chance of error. Furthermore, the use of a controllable payload allows for the investigation of specific areas, allowing for the gathering of data of greater importance while limiting the need for secondary missions. Finally, the use of a UAV allows for a greater amount of data to be

58 collected due to the extended flight time of a UAV platform compared to other payload options. 4.2 SCIENCE VALUE PAYLOAD SCIENCE OBJECTIVES There are two different aspects to the payload, each with their own objectives; the SMD payload requirement and decreased complexity in UAV flight. The payload objectives relating to the SMD payload are to log atmospheric pressure, temperature and humidity along with solar intensity and UVI data at 5-second intervals as well as taking at least two still images during flight and three after landing. The payload objective relating to decreasing the complexity in flying a UAV is to complete the flight and mission (visually locating the rocket and landing) solely using the software provided at the ground station without reverting to back up manual control PAYLOAD SUCCESS CRITERIA The data logging and sensors shall be deemed successful if the payload obtains and logs atmospheric pressure, temperature and humidity along with solar intensity and UVI data at 5-second intervals as well as taking at least two stills during flight and three after landing. It shall be deemed a success regardless if the data is collected by the main or the back-up sensor package. Fulfilling the SMD payload requirement successfully shall also demonstrate the flexibility in the UAV design. If the UAV operator successfully visually locates the rocket and lands in a state fit for reusability, without resorting to use of the back-up manual flight control then this will demonstrate successful reduction in complexity of UAV control. 4.3 PAYLOAD DESIGN SUMMARY

59 FIGURE 4-1: UAV FIGURE 4-2: UAV (STOWED CONFIGURATION) Airfoil Note: All airfoil estimates are made using 2-dimensional flow assumptions. The first design parameter to be chosen was the airfoil. A chord of 5 in. was chosen for the airfoil because it very nearly the largest chord that can be used on the airplane without making it impossible to fit inside the rocket. Using this chord we then plotted coefficients of lift and drag for various angles of attack using Xfoil (Xfoil is CFD command line program developed at MIT for designing and analyzing airfoils).

60 Atmospheric data was found using a java applet by DesktopAeronautics 5. Standard cruising conditions were considered to be the following: Velocity = 25m/s ( ft/s) Altitude = 2500 ft Reynolds number = 204,000 Using this Reynolds number, nine different airfoils were analyzed. The 6 airfoils with the highest C L are shown in Figure

61 FIGURE 4-3: CP DISTRIBUTION FOR VARIOUS AIRFOILS The leftmost graph shows the coefficient of lift plotted against the coefficient of drag. The center graph shows C L vs. α (the steeper graph) as well as the pitching moment (C m ) vs. α (the flatter graph). After some consideration, the NACA 4412 was chosen because it has a very high ratio of C L /c d and it has one of the highest C L around an angle of attack of 3, which will be the UAV wing s angle of attack while cruising. The 3 angle of attack was chosen because it provides significantly more lift but doesn t produce too much drag. Also, higher angles of attack would make it more difficult for the UAV to fit inside the rocket. The NACA 4412 airfoil is a thinner airfoil which makes it more difficult to place structures inside it (most importantly the dihedral lock mechanism

62 which will be discussed in a future section). This problem is more than made up for by the reduction in drag and the reduction in airfoil thickness. The reduced airfoil thickness means the wings take up less space inside the rocket which leaves more space for a larger fuselage. Also, due to the mechanics of the wing rotation mechanism, the mechanism s height is directly proportional to the amount it must compress which would be increased by a thicker wing. This would further complicate its design due to the limited height of the interior of the fuselage (the wing rotation mechanism will also be discussed in detail in a future section). At a 3 angle of attack and the cruising conditions outlined above, the NACA 4412 airfoil has a 2-D coefficient of lift of The stall angle of attack at 25 m/s was determined to be roughly 14. The 2-D glide ratio and 2-D glide angle at 3 are 0.08 and 4.6 respectively. FIGURE 4-4: CL DISTRIBUTION FOR NACA 4412 AIRFOIL

63 Fuselage Design Since the CDR, the fuselage been extended to a new length of 50 in. to increase the wingspan. The fuselage is rotationally symmetric except for two flat sections: one at the nose to accommodate a camera window and one on the bottom for a Kevlar-epoxy skid plat. The UAV does not possess landing gear due to the limited space inside the rocket so the Kevlar-epoxy skid plate will protect the fuselage when the UAV lands. The fuselage will be made of fiberglass and its structural integrity will be further ensured by means of bulkheads inside the fuselage. The bulkheads also provide sturdy attachment points for the wings and tail. There are two access ports in the fuselage, one extending from the nose to the wings and the other extending a short distance behind the wings. The access ports will provide easy access to all of the avionics equipment, sensors, camera, and battery. The access ports will be attached by means of interlocking bulkhead halves made of plywood (shown in Figure 4-5). FIGURE 4-5: ACCESS PORT BULKHEADS

64 FIGURE 4-6: ELECTRONICS BAY ACCESS PORTS The seam between the access port and fuselage will then be sealed by tape to avoid any disturbances in the airflow. The electronics housed inside the UAV mostly reside in a polycarbonate frame located behind the wings. The camera is located behind a scratch resistant acrylic window which will provide a wide range of view. The batteries are located in their own polycarbonate case in front of the wings in order to bring the center of gravity of the UAV forward to the desired location.

65 FIGURE 4-7: TOP VIEW OF ELECTRONICS BAY A folding propeller is used to ensure that the UAV fits inside the rocket. The motor and a folding propeller are located at the rear of the aircraft which allow an unobstructed view for the camera at the front of the aircraft and allows the propeller to automatically fold back when it stops spinning. The propeller will fold up while gliding to reduce drag and upon landing to avoid incurring damage. Manufacturing The fuselage for the UAV is made with a wooden mold that is made on a wood lathe. Using this positive mold, the body is made using three plies of fiberglass. A standard layup process is used in the fuselage s fabrication, and the composite body is allowed to cure in the team s oven. Due to the cylindrical shape of the fuselage, only one half can be made at a time. During the layup process, wrinkles form around the nose of the fuselage. These wrinkles are sanded down and any holes or imperfections that arise during this process are fixed by putting the fuselage over the mold, applying strips of fiberglass to the necessary areas and wetting the fabric with thirty-minute epoxy. The nose section is then vacuum bagged while it is still on the mold and allowed to cure. Shapes are cut out of the fuselage halves to allow for the wing rotation mechanism, the camera window, the solar sensors window, and the tail frame. The camera window is then installed in the bottom half of the fuselage. The bulkheads for the electronics structures, wing rotation mechanism, and tail are glued to the fuselage and any

66 remaining internal components are installed. The two halves of the fuselage are then joined together with a thin strip of fiberglass and epoxy. Afterwards, the two access ports are cut out. Wings The total wingspan was increased slightly to reduce wing loading but unfortunately 13% of the total wingspan is not covered by an airfoil is thus nonlifting. The lifting wingspan is therefore only 53.5 in. which results in a lifting wing area of in 2 and a wing loading of lbs/in 2. The wings have a dihedral of 5 which begins 2.75 in. out from the fuselage on each wing to improve stability. The wings are unable to fit inside the rocket while in their 5 dihedral so 2.75 in. out from the fuselage there is a steel hinge on either wing which allows the wing to fold flat. The wings are naturally pushed into their 5 dihedral once outside the rocket by the air pushing on them and are held in place by spring loaded latches located inside each wing.

67 FIGURE 4-8: WING DIHEDRAL The wings are attached to the fuselage by means of a rotating mechanism so that they can be rotated back while the UAV is stowed inside the rocket. The mechanism stores energy in one torsion and one compression spring which unfold the wings once the UAV is released from the rocket. The wings are kept from unfolding inside the rocket by means of a rope that is wrapped around them and held together with a pin that is pulled out after the UAV/sabot leaves the rocket. In order to fold up, one wing folds on top of the other. When unfolding, the upper wing fully rotates out before it falls down into its final position. Shoulder bolts keep the wings from rotating out too far and cap screw heads on the lower wing fit into holes in the upper wing to keep them aligned with each other. A polycarbonate glide plate prevents the upper wing from falling into place too early which reduces the chances of failure. Also, two platforms in the mechanism force the upper wing to fully rotate into position before it can fall down which keeps the wings from rubbing against each other when unfolding. FIGURE 4-9: WING LOCKING MECHANISM

68 The plates attaching the wings to the rotator are now made of FRP fiberglass sheets to reduce the mechanism s weight. The wing rotation mechanism is made of T6 aluminum. Manufacturing The first step in the wing manufacturing process is to cut precision foam cores on CNC foam cutter. These foam cores are then glued to the outer aluminum airfoil rib on the mounting plates attached to the dihedral hinges. The wings are then laminated with a layer of 2.0 oz fiberglass and a 2.5 strip of carbon fiber weave between 0.1- to 0.6- chord on both the top and bottom surfaces. Mylar is used on both surfaces to make them perfectly smooth. After curing, the wing tips and trailing edges are trimmed square. The next step is in-laying the aileron servos into the wings. We chose to use servos mounted inside the wings instead of a pushrod system because of the risk involved of the rotating wings shearing the pushrods. At this point, the wings are attached to the wing rotating mechanism. Monokote is applied to the exposed aluminum rib sections (the sections of the wings without a foam core) and to cover the servos. Tail FIGURE 4-10: TAIL SECTION

69 FIGURE 4-11: TAIL SECTION (STOWED) Final Design Choices Both the horizontal and vertical stabilizer areas have increased since CDR to provide more stability for the UAV. The increase in area was achieved by eliminating the tapers on both stabilizers and increasing the horizontal and vertical stabilizer average chords to 4.6 in. and 7.13 in. respectively. The corners of the stabilizers are now rounded to mitigate some of the increase in drag caused by eliminating the taper. Stability is important because the UAV is flown by an autonomous computer program. The UAV will mostly fly in a straight line and quick maneuvers will not be required during its flight so an over stable aircraft will not be problematic. The new estimated vertical and horizontal tail volume coefficients are shown below along with their percent increase from the preliminary design. Manufacturing Plywood is now used in place of foam/fiberglass for the tail. This change was made to simplify the tail manufacturing process (The tail sections were very small and thus very difficult to cut on a foam cutter). The plywood is cut on a computerized laser cutter using *.dxf files exported from SolidWorks. The elevator and rudder are cut separately on the laser cutter. They are then attached to the tail by means of hinge tape. The two hinge surfaces are also rounded to provide a smooth hinge. Finally the leading and trailing edges of the tail surfaces are rounded somewhat to provide a more streamlined surface and reduce drag.

70 Stability The center of gravity was initially chosen to be located 25% of the fuselage length from the nose (11.25 in.). The fuselage length has since been increased to 50 in. without moving the absolute position of the center of gravity which means the X cg is now 22.5% of the fuselage length from the nose. The placement of the center of gravity will be accomplished by putting the heavier avionics equipment such as the batteries in front of the wings. Because the wings must fold up, the center of gravity, and consequently the wings, are placed close to the front of the aircraft in order to maximize the (wing span)/(fuselage length) ratio. After placing the center of gravity, the horizontal and vertical stabilizers were sized. An initial vertical tail volume coefficient (V v ) and a horizontal tail volume coefficient (V h ) of 0.05 and 0.5 respectively were used. The following are the two equations used to estimate tail area using the desired volume coefficients: Since the PDR though, the tail volume coefficients have been increased to improve stability by removing the tapers, increasing the average chord, and increasing the tail moment arms. The new tail volume coefficients are shown below. V h % Increase V h % Increase Because the UAV is flown by autopilot, we chose a positive static margin of 0.1 (S.M.). The distance from the wing leading edge to the neutral point and the center of gravity of the aircraft were then estimated using the following equations: x np c AR AR h 4 V h AR 2 The results were then used to calculate the distance from the nose to the wing s leading edge to be 8.0 in. 4.4 ASSEMBLY AND TESTING

71 4.4.1 PLANNED COMPONENT, FUNCTIONAL AND STATIC TESTING The team has performed preliminary testing on the structural components of the payload as well as expected performance of the payload components. Testing progress is as follows: To test whether the wing rotator will deploy wings from a folded configuration the rotator was outfitted with prototype wings and folded into stowed configuration with wings held in place. Wings were then released and rotator was able to unfold wings and lock into flight configuration. A prototype UAV was launched during a full scale rocket launch to test the wings capability of separating the sabot halves upon deployment from rocket tube, ensuring that the spring components of the wing rotator mechanism are capable of separating sabot. The sabot was successfully able to slide out of the rocket body and the wings were able to separate the sabot halves. Further testing is necessary to ensure reliable performance of the payload. The tests to be performed are as follows: Wings will be tested further with variable mass, using small weights. As verified by initial testing, wings will not reach full load capacity, failing prematurely at the point of contact between the wing and the edge of the table holding the wings. Horizontal stabilizer will be tested in the same manner as the wings. Test the wing rotator several times in the following manners: o A symmetric variable mass on attached wings, and wings are expected to fail before the aluminum wing rotator. Will apply variable mass until wings break off rotator. o Drop test UAV from tethered weather balloon. Fuselage will be tested in a manner of ways: o Crush test from nose to tail o Impact test from the top of the body and the bottom of the body. This will ensure that the fuselage will be able to withstand the impact force expected upon landing. Will also verify the connecting points between the two halves of the fuselage will withstand forces upon deployment and landing. o Abrasive test on the Kevlar landing strip, testing the resilience of the Kevlar for the skidding expected upon landing. Several flight tests to ensure that control gains are suitable for the final UAV design and prototype. Drop test from tethered weather balloon, testing avionics system and prototype UAV work under mission conditions.

72 4.4.2 TEST MEASUREMENT, VARIABLE AND CONTROLS Testing and verification of the avionics occurs in three distinct phases: ground testing, on a test aircraft and lastly on the final UAV, thus enabling ground testing shall consist of validating the correct operation of all hardware and sensors in a non-critical environment. The testing on the test aircraft serves to verify that the subsystems within the avionics systems work as expected in flight case and to validate changes made to the flight computer hardware and software for the purposes of the competition. The flight testing on the UAV is to demonstrate the avionics system is able to function correctly in its intended flight configuration and importantly, that it is capable of recovery after deployment from the rocket. Phase One Ground Testing The flight computer, GPS/IMU, and telemetry boards have all been connected and shown to be functional. The control servos were connected to the avionics system and actuated through a program to ensure that the servos respond as expected. An R/C receiver has been attached to the avionics system and the system has been confirmed to be responsive to commands from and R/C handset. Furthermore, the telemetry system has been tested by noting the accuracy of the GPS. It was found that the accuracy in the position reported by the GPS is satisfactory horizontally, but unsatisfactory in the reported vertical position. To obtain greater accuracy in the vehicle s altitude, pressure altitude data and the GPS altitude data will be mixed. The time for the GPS to get position lock at start-up was found to be within ten seconds, and the GPS was found to be able to maintain a good lock when moving at a reasonable speed. Back-up boards are in the process of being constructed and tested. These boards will undergo tests to ensure that they function properly under a variety of conditions, such as varying altitude and weather conditions. The real-time video system has been tested and we were unable to reach the maximum line of sight range at which the system was no longer able to adequately transmit video data. It can thus be concluded that for the purposes of this rocket launch, the system s performance is more than acceptable. Phase Two- Test Aircraft Initial flight-testing has occurred on a previously built UAV owned by the rocket team. The test aircraft was tested in a manual configuration, i.e. the R/C aircraft was controlled solely by an operator using a standard R/C controller to verify that the flight computer and control servos were setup correctly. Stability tests were then performed to determine the control gains of the flight computer necessary to achieve stable flight. It is worth noting that these gains will not necessarily be those required for the final UAV, but the autonomous flying ability gained from this is essential for further flight testing.

73 Further testing is necessary, specifically to determine if the UAV is able to navigate to the coordinates of waypoints uploaded to the avionics system in-flight. For this testing, the flight computer hardware will be wired to the primary sensors and the flight computer software modified to log the sensor data on the internal volatile memory and transmit the logged data post landing. This functionality will then be tested in multiple flights to ensure correct operation. Furthermore, the back-up sensor board, real-time video transmission and still capturing systems shall be integrated into the R/C aircraft, tested and refined as necessary. Phase Three UAV Testing The avionics system will then need to be tested with the final version of the UAV. The first flight testing shall be to determine the control gains required for stable flight of the UAV. For the purposes of these tests, the equipment that is not essential for flying (i.e. everything but flight computer, telemetry link and GPS/IMU) shall be replaced by appropriate ballasting to minimize the risk of damage to components. Once adequate control gains have been determined, a series of flight tests shall be undertaken to ensure that the sensor systems and data logging systems, as well as the imaging systems, still function as desired. These flights will also determine if the propulsion system s duration and thrust are sufficient to maintain steady-level flight for at least 30 minutes. Further testing representative of flight scenarios shall also be undertaken, including point-to-point flying based on user inputs at a ground station. Drop tests from a tethered weather balloon shall also be used to simulate UAV deployment to ensure the UAV/Avionics is capable of recovering from the postdeployment dive. The UAV will be unpowered (propulsion system off) due to safety reasons for these tests; the lithium polymer propulsion battery will be replaced by ballast to mitigate the risk of the lithium polymer battery exploding due to damage if the UAV were to crash. Gliding should be sufficient to test all avionics. A test section of the rocket body tube will be hung from a balloon platform attached to the weather balloon. The UAV will be packed into the sabot, and the sabot will be placed in the body tube and connected to a radio controlled Tender Descender. The balloon will be tethered and raised to an altitude sufficient enough such that the UAV and sabot will be falling at speeds identical to those of launch conditions when approaching an altitude of 200 ft, releasing from the rocket tube at 200 ft; this altitude should be sufficient for full UAV deployment, while restricting the safety radius needed to be cleared of personnel on the ground to a reasonable value. The sabot will be dropped under drogue parachute, and the UAV will deploy. These tests shall be performed with ballast instead of non-essential electronic components. This ballast will be placed in such proportions and arrangements to maintain the center of mass of the UAV, providing sufficiently accurate mission conditions for the UAV. 4.5 SAFETY AND ENVIRONMENT

74 4.5.1 IDENTIFICATION OF SAFETY OFFICER Andrew Wimmer will be the primary rocket safety officer for the team. Ben Corbin is the team s MIT EHS representative and is the assistant safety officer and is in charge of safety issues not directly related to the rocket. Both team members have considerable experience in their respective areas UAV is damaged before launch weekend PAYLOAD FAILURE MODES Failure Mode Description Consequence Mitigation Failure of a The UAV needs to critical system be repaired or during testing or replaced quickly to improper allow for testing to handling continue. UAV is damaged on launch weekend before launch The UAV is damaged during deployment The UAV s autopilot fails Recontact sabot Contact rocket body with with Drifting of gyroscopes in the inertial navigation system during launch queue The degree of damage needs to be assessed. Quick repair or complete replacement with a backup UAV needs to be done. The UAV may not be controllable in autopilot or manual mode, and part or the entire mission may fail. The backup pilot engages manual control and the UAV is flown like a normal R/C aircraft. All data can still be collected and transmitted. Avionics testing in a commercial R/C aircraft and ground testing of the UAV will be performed before the UAV is flown. Extensive flight testing with the final UAV design will be performed, and an experienced backup pilot will be standing by during any flight test with the R/C controller to mitigate chances of a crash landing. Extreme care will be taken when packing and handling the UAV. A backup UAV will be taken to the launch site to prevent total mission failure in the event of the primary UAV being damaged. High strength materials, including fiberglass, carbon fiber, polycarbonate, aluminum, and steel, are used in the construction of the UAV to prevent total failure and part separation. A backup pilot with manual R/C controller and a spotter with binoculars will be on hand at all times, ready to take over control of the UAV in case of any autopilot failure.

75 The UAV becomes entangled in shock cords or shroud lines upon exiting the sabot. Control surface(s) break off during initial dive pull-out Wing breaks off immediately after deployment High-wind conditions on launch day The UAV will not be able to fly. The pilot will immediately go to manual mode to prevent the UAV from trying to correct its flight and potentially damage the rocket. The UAV will still be able to collect and stream data. Rocket parachutes may not deploy correctly. Depending on the control surface, the UAV may no longer be maneuverable and will likely crash land. Data can still be collected. The UAV will no longer be maneuverable and will crash. Depending on the damage sustained, data may or may not be able to be collected. The UAV could be damaged by wind gusts. The UAV experiences turbulent flying conditions, making control difficult. The location of the parachutes and the length of the shock cords are such that the UAV should avoid entanglement. Testing will drive adjustments to the rocket s recovery system to mitigate UAV entanglement. Balloon drop tests will verify that the UAV can pull out of a dive and fly without sustaining damage. The UAV s wing rotator and wings have been over-engineered to sustain much higher loading than could be experienced throughout the entire flight. Balloon drop tests will verify that the UAV can pull out of a dive without sustaining damage. The UAV structure is designed to handle high-loading situations, such as those experienced by strong winds. The UAV autonomous controls will be able to stabilize the flight of the UAV in high winds. The UAV design will not need to take

76 The propulsion system fails The UAV is out of sight at time of deployment The UAV makes it out of the sabot, but does not right itself (The autopilot and GPS require the UAV to be upright.) The UAV s propeller Video may be shaky. The UAV will not be able to sustain or gain altitude, decreasing mission time. The UAV will still be controllable. No knowledge of whether or not the UAV has successfully pulled out of its initial dive. Backup pilot cannot engage manual control without sight of UAV. The backup pilot will engage manual control and hopefully be able to steady the UAV. If control is not gained, the UAV will crash land, but still be able to collect and transmit data. Damage to the rocket and UAV into account flying in very high winds (15mph+), because it is likely that the entire launch will be postponed if high-wind conditions are present due to the danger of launching high power rockets in high-winds. Flight testing and pre-flight checks will ensure that all UAV systems are working properly before final integration. The avionics are designed to not need a propulsion system to function properly, and will still be capable of landing the plane safely. Not having sight of the UAV should not be a problem, as the autopilot in the UAV will undergo extensive testing before launch day, and the tablet pilot interface should work as planned. A spotter with binoculars, whose main responsibility is to track the UAV by sight, will be with the backup pilot at all times. The live video feed and GPS coordinates transmitted from the UAV should also give information pertaining to the status of the UAV. Upon release from the sabot, the UAV will immediately and automatically set its control surfaces to a position that should pull it out of the initial dive in an upright position. Wing dihedral should make being upright a much more favorable position. If all this fails, the backup pilot has a long time (2500 ft) to steady the UAV with manual control. The autopilot will be programmed not to engage the motor until a set

77 contacts part of the rocket while under throttle An internal component in the UAV moves during flight, shifting the UAV s CG The UAV deploys correctly, but flies over the crowd or out of range of the ground station. UAV contact with ground station is lost The UAV requires an emergency or forced landing due to one of the above or other risks may occur. The UAV may be put into a spin due to such contact. The UAV becomes difficult to maneuver. Autonomous flight capability may be lost and manual control may have to be engaged. Manual control is engaged and the UAV is piloted out of the no fly zone. Contact will try to be regained and/or manual control will be engaged. Manual control is engaged and the backup pilot brings the UAV down as quickly and safely as possible. If that is not an option, an emergency kill time has passed from deployment. If this fails, the backup pilot can engage manual control and set the motor throttle to zero. All components will be securely mounted. Balloon drop tests will verify all component mounts are secure enough. A no fly zone will be coded into the tablet pilot system disallowing the addition of waypoints above the crowd or outside the range of the ground station. The rocket will be launched far away from the crowd, as per NAR regulations, so the chance of UAV deployment over crowd is minimal. The range of the ground station will be large enough to cover the UAV at any possible deployment location. The UAV will be programmed with a loiter-mode that will be activated if ground station signal is lost. The UAV automatically lands at a set location after a set amount of time out of contact with the ground station. In loiter-mode, the UAV will still avoid the no fly zones. Low cruising/reconnaissance altitude allows for minimal landing time in case of an emergency. The mass of the UAV has been minimized to decrease damage in case of an emergency. The addition of an emergency kill switch in the programming (controlled by the

78 GPS Failure Telemetry Failure Video Camera Failure Still Camera Failure Flight Computer Sensors Failure Backup Board Failure Flight Computer Failure GPS hardware failure Satellite lock failure Telemetry module hardware failure Antenna failure Camera hardware failure Antenna failure Servo failure actuation Sensor hardware failure Arduino board failure Sensor hardware failure SD card failure SD card writter failure Flight computer hardware failure Flight computer software freezes switch is engaged that puts the plane in a steep downward spiral. No navigation data to run UI Loss of communication between UAV and Groundstation Loss of first person view video stream Loss of capability to take still photos No sensor data collected by the flight computer No sensor data transmitted to ground station No sensor data collected by the backup board Loss of communication between UAV and Groundstation backup pilot) is a last resort to quickly land the UAV. Testing to ensure antenna function and adequate GPS lock prior to launch day Flight is continued under manual mode using visual data Testing to ensure antenna and telemetry module functions adequately prior to launch day Flight is continued under manual mode using visual data and sensor data logged by backup board Testing to ensure antenna and video hardware functions correctly Flight is continued using pilot and spotter team Stills taken from the video stream during flight and landing If combined with a video camera failure then results in a loss of image gathering capability Backup board logs sensor data using an independent set of sensors to an SD card Flight computer logs sensor data and transmits it to ground station If combined with flight computer sensor failure then results in a lack of sensor data Manual control and a pilot spotter teams glide the UAV down safely

79 Gyroscopic Sensor Failure Accelerometer Failure Loss of Motor Power Loss of Avionics Power Primary Groundstation Computer Failure Secondary Groundstation Computer Failure Flight computer software crashes Sensor hardware failure Sensor hardware failure Motor dedicated battery pack runs out Motor hardware failure Avionics dedicated battery pack runs out Computer system crash Application crash Laptop battery runs out Computer system crash Application crash Laptop battery runs out No sensor data collected by the flight computer Motor throttled to idle Loss of orientation data Little consequence Motor throttled to idle Loss of communication between UAV and Groundstation No sensor data collected by the flight computer Motor throttled to idle Loss of telemetry link Loss of first person video feed Can be flown under manual control if control loops become unstable Can be flown under manual control if necessary UAV can be glide to landing via the groundstation UAV performs uncontrolled glide to ground Multiple backup computers capable of running the groundstation software Employ manual control until software restart Multiple backup computers capable of displaying video feed Use a spotter in place of video feed until software restart Note, in all circumstances where control of the UAV is lost, with the exception of loss of avionics power, the UAV will be able to collect and transmit data while it is falling POTENTIAL HAZARDS For manufacturing hazards, see Hazards Recognition and Tool Use in Section

80 In order to assure safe and successful operations concerning the payload, a checklist must be followed. In order to reduce personnel hazards the following precautions must be taken: Avoid standing in the plane of the propeller when UAV propulsion system is on. Do not try to catch the UAV during landing. Make sure all relevant testing (reference checklist) has been completed prior to attempting a flight test. Make sure the checklist is followed and all steps are completed properly in a thorough, workmanlike manner to assure mission success. Lithium Polymer Battery Hazards and Procedures: Always charge lithium polymer batteries with a balancer. Out of balance packs can explode. Never over-discharge a lithium polymer battery (below 2.7V per series cell). Always use an electronic speed controller (ESC) with a low voltage cut off feature. Never attempt to charge a lithium polymer battery if it looks bloated, damaged, over discharged (below 2.7V per series cell). Damaged packs can explode. Never leave a lithium polymer battery unattended while charging. Always charge lithium polymer batteries on a non-flammable surface and away from flammables. Take extreme caution around the UAV in the case of a crash. The battery packs may explode if damaged. Never discharge a lithium polymer battery at more than the published discharge rate. The pack may explode if discharged too quickly ENVIRONMENTAL CONCERNS All waste materials will be disposed of using proper trash receptacles Consideration of environmental ramifications will be made regarding applicable activities The following materials have been identified as potentially hazardous: Aeropoxy 2032 Epoxy Resin Aeropoxy 3660 Hardener Lithium Polymer Batteries See CDR-MSDS document for complete MSDS specifications on these and other materials 5 LAUNCH OPERATIONS 5.1 LAUNCH OPERATIONS AND PROCEDURES

81 5.1.1 CHECKLISTS Caution Statement Recall the Hazards Recognition Briefing. Always wear proper clothing and safety gear. Always review procedures and relevant MSDS before commencing potentially hazardous work. Always ask a knowledgeable member of the team if unsure about equipment, tools, procedures, material handling, and/or other concerns. Be cognizant of your and others actions. Keep work station as clutter-free as possible. Equipment Packing Checklist: 1. Support Equipment and Tools a. Safety Gear i. Goggles ii. Rubber Gloves iii. Leather/Work Gloves iv. Face Masks v. All Safety Documents and References b. Furniture i. Tent (1x) ii. Tables (2x) iii. Chairs (6x) iv. Rocket assembly benches c. Generator i. Gas ii. Power Strip(s) (3x) iii. Extension Cord(s) (3x) d. Tools i. Corded Drill ii. Cordless Drill 1. Cordless Drill Batteries 2. Charger iii. Drill Bit Index(s) iv. Wrench Set v. Pliers vi. Screwdriver Set vii. Hex Keys Set viii. Files ix. Sandpaper x. Knives xi. Flashlight

82 xii. Soldering Iron 1. Solder 2. Solder Wick 3. Sponge xiii. Wire Cutter/Stripper(s) xiv. Extra Wire (Black and Red) xv. Pocket Scale e. Adhesive i. 5-minute Epoxy (2 part) ii. CA and Accelerant iii. Aeropoxy (2 part) iv. Epoxy Mixing Cups v. Popsicle Sticks vi. Foam (2-part) vii. Foam (solid) f. Other supplies i. Tape 1. Duct Tape 2. Scotch Tape 3. Vacuum Tape 4. Electrical Tape 5. Masking Tape 6. Gaffer s Tape ii. Trash Bags iii. UAV Camera Port Cleaner iv. Isopropyl Alcohol (general clean up) v. Water Bottle vi. Camera Lens Cleaning Supplies vii. Paper Towels viii. Wipes ix. Spare Hardware x. Lithium/Silicon Grease (for building reload; other) xi. Zip-ties xii. Talcum Powder (for parachutes) 2. Ground Station a. Antennas i. Rocket (1) ii. UAV (3) iii. Antenna Mounts b. Emergency Locator Transponder (ELT) (UAV and Rocket) (3x)

83 c. Emergency Locator Receiver d. UAV Main Pilot Computer e. UAV Secondary Computer f. Rocket Ground Station Computer g. UAV Manual R/C Controller h. Binoculars i. Monitors j. Power Adapters for all Computers k. Mice (3x) l. Cables i. Antennas ii. Monitors iii. Other m. Miniature Weather Station (wind speed/direction, temperature) 3. Launching Equipment a. Launch Pad b. Launch Rail c. Stakes for Pad d. Angle Measuring Tool e. Electronic Launch System (ELS) i. Battery ii. Battery Charger iii. Controller iv. Leads 4. Rocket a. Body i. Lower Tube Section ii. Upper Tube Section iii. Nose Cone iv. Ballast v. Shear Pins (10x) b. Recovery i. Parachutes 1. Drogue (2x) 2. Main (2x) 3. Nomex Parachute Protectors (3x) ii. Shock Cord iii. Ejection Charges 1. Black Powder 2. Charge Holders (4x)

84 3. Igniters (4x) iv. Charge Released Locking Mechanism (2x) v. Quick links (10x) vi. Duffel bag for parachutes during recovery of rocket c. Motor i. Casing ii. Reload (2x) iii. Retention 1. Retention Plate 2. Retention Hardware d. Avionics i. Avionics Bay ii. Altimeters 1. ARTS2 (1x) 2. ARTS2 Transmitter Board (1x) 3. MAWD (1x) iii. Antenna (attached to outside of rocket body) iv. 9V Batteries (10x) v. ELTs (one in Bay, one in nose cone) (3x) vi. Hardware x1 bolts (10x) locknuts (6x) 5. UAV a. UAV b. Motor (2x) c. UAV Propeller (3x) d. UAV Lithium Polymer Batteries (2x) and Spare Batteries (3x) e. Lithium Polymer Battery Charger/Balancer f. Spare Servos (3x) g. Spare Control Linkages h. Sabot i. Avionics i. Flight Computer ii. Back up Sensor Logging Board iii. Sensors iv. Flight Digital Still Camera v. Video Board and Video Camera vi. Manual Control Receiver (Back Up: 72MHz) vii. Antennas (72MHz, 900MHz, 2.4GHz) viii. ELT

85 6. Miscellaneous a. Digital Camera b. Video Camera c. Extra Batteries d. Two-Way Radios e. Two-Way Radio Chargers f. Manuals for all Equipment and Gear Pre-Flight/Final Assembly Checklist: 1. Ground Station a. Furniture Set Up b. Generator i. Full Tank ii. Extra Gas iii. Connect Extension Cord(s)/Power Strip(s) c. Computers i. Set Up ii. Plug in Power Adapters iii. Mice iv. Set Up Monitors v. Power Up d. Antennas i. Mount and Set Up GHz MHz ii. Connect to Computers e. Set Up ELT Receivers i. Test on each of 3 channels 2. UAV a. Mechanical i. Inspect Fuselage (follow detailed checklist) 1. Internal Structure 2. External Structure 3. All Electronics/Avionics Mounts 4. Motor Mounted Securely 5. Kevlar Skid Plate ii. Inspect Wing and Wing Folding Mechanism iii. Test Wing Folding Mechanism 1. Fold and let Unfold at least twice 2. Adjust as necessary

86 iv. Inspect all Hinges v. Test All Folding Hinges 1. Fold and let Unfold 2. Adjust as necessary vi. Unfold Everything vii. Inspect All Control Surfaces 1. All should be free and clear to rotate 2. Inspect and Move All Hinges 3. Inspect Control Linkages and Servos viii. Inspect Camera Dome 1. Clean Dome if necessary 2. Check Connection to Fuselage 3. Check Camera Mount ix. Inspect UV Sensor Window 1. Clean if necessary b. Power Systems i. Inspect Motor ii. Check if Propeller Secure iii. Give Motor a Test Spin (by hand) iv. Inspect Motor Controller v. Make sure all electronics are Switched Off vi. Connect and Secure Charged Lithium Polymer Batteries c. Avionics i. Install Flight Computer ii. Install Back up Sensor Logging Board iii. Install Video Board and Video Camera iv. Install Digital Camera v. Install Manual Control 72MHz Receiver (Back Up) vi. Inspect All Sensors vii. Install ELT viii. Connect Everything ix. Set No-Fly Zones x. Set Loiter-mode Landing Location d. Communication/Controls i. All servos connected to proper channels ii. All Avionics Connected iii. Power On iv. Test All Control Surfaces (using standard/manual R/C 72MHz transmitter) 1. Trim

87 2. Actuate one direction 3. Actuate other direction v. Test Motor (using standard/manual R/C 72MHz transmitter) 1. Clear objects/people from the plane of the propeller 2. Throttle Up 3. Throttle Down vi. Power Motor/Motor Controller Off vii. Test Flight Computer 1. Communicating with Ground Station viii. Test Data Feeds (turn UAV avionics on) 1. Temperature 2. Humidity 3. Solar Irradiation 4. UV Irradiation 5. Pressure ix. Test IMU/GPS 1. Transmitting Telemetry x. Test Autopilot (Make sure control surfaces respond correctly) 1. Pitch UAV Up 2. Pitch UAV Down 3. Yaw UAV Right 4. Yaw UAV Left 5. Roll UAV Left 6. Roll UAV Right xi. Test Data Logging 1. Digital Camera Still Shot Recorder 2. Back Up Sensor Data Logging xii. Test Video Feed 1. Receiving Video xiii. Test ELT 1. Receiving ELT signal xiv. Power Up Motor/Motor Controller xv. Flight Test with Manual R/C Control (no autopilot) 1. Receiving All Data 2. Proper Control Responses xvi. Ground Test of Point-and-Click Control (with autopilot) 1. Receiving All Data 2. Proper Control Responses xvii. Aerial Test of Point-and-Click Control 1. Trim control surfaces before flight

88 2. Back Up with Manual R/C Control e. Switch out Lithium Polymer Batteries f. Final Overall Inspection g. Install UAV into Sabot. See Payload Integration Plan. 3. Rocket a. Lay-out rocket sections in order b. Check Body Antenna c. Install Ballast into appropriate sections of sabot and body tube d. Refer to Payload Integration Plan i. Follow, then continue with this checklist e. Install all shear pins f. Prepare Motor Reload i. Safety Officer will oversee this step g. Slide motor casing into rocket h. Screw on motor retention i. Make sure the tube-tube and tube-nose cone interfaces are secure j. Inspect rail guides k. Do a pre-launch briefing Launch Checklist: 1. Get approval from event administration to set up pad, ELS, and rocket 2. Set up pad 3. Tip pad over and install rail 4. Check all tube interfaces 5. Slide rocket onto rail down to stop 6. Tip up launch pad 7. Stake pad to ground 8. Arm Electronics a. Have manuals on-hand b. Listen for proper beeps 9. Put igniter into motor and secure it 10. Connect launch clips 11. Connect ELS to battery 12. Clear launch area/back up appropriate distance 13. Make sure Ground Station and Pilots are ready 14. Get approval from event administration for launch The following depend on procedures outlined by event administration: 15. Cameraman ready 16. Check to see if range and skies are clear

89 17. Insert key into ELS check continuity 18. Countdown from Launch 20. Remove key from ELS 21. Disconnect ELS from battery 22. Recover Rocket and UAV RECOVERY PREPARATION Using a short length of nylon webbing, attach the inside of the deployment bag to the loop at the top of the main parachute. Secure the top of the deployment bag to a table leg or other hard point. Stretch out the parachute and untangle the lines. Place a weight on the outstretched, untangled lines, to hold them in place. Next, flake out the canopy.

90 Next, fold the canopy width-wise so it can fit inside the deployment bag. Fold the leader connecting the deployment bag and the parachute in a figure 8 and secure with a rubber band. Place inside deployment bag. Begin placing the canopy inside the bag, folding it over itself in an S pattern.

91

92 After the parachute is in the bag, begin folding the shroud lines, again, in an S-pattern. Tuck the folded lines into the bottom of the deployment bag. Secure the Velcro flap of the bag MOTOR PREPARATION One of the team s L2 members will supervise motor assembly. All fire hazards, e.g. people smoking, lighters, potential ignition sources, will be removed from the immediate surroundings during motor preparation. See Appendix document for official and detailed Pro75 motor preparation instructions. The assembled motor will be slid into the motor tube, and motor retention will be screwed on IGNITER INSTALLATION Once the rocket is on the pad, tipped vertical, and all electronics are armed, the motor igniter will be installed. Care will be taken to fully insert the igniter into the motor. The igniter will be held in with tape and a 1/8 dowel, which will be easily pushed out when the motor lights. Launch lead clips will be securely attached to the igniter leads at the appropriate time.