USLI Flight Readiness Review

Transcription

1 UNIVERSITY OF MINNESOTA TWIN CITIES USLI Flight Readiness Review University Of Minnesota Team Artemis 3/26/2012 Flight Readiness Report prepared by University of Minnesota Team Artemis for NASA University Student Launch Initiative.

2 University of Minnesota USLI Team Mark A Senior - Aerospace Engineering, Team Lead Aboto001@umn.edu Chris H Senior - Mechanical Engineering, Payload Team Eric W Senior - Mechanical Engineering, Vehicle Team Derek L Senior - Aerospace Engineering, Vehicle Team Peter L Senior - Aerospace Engineering, Vehicle Team William Zach M Senior - Aerospace Engineering, Payload/Vehicle Team Takuhiro S Senior - Aerospace Engineering, Payload/Vehicle Team Travis S Senior - Aerospace Engineering, Payload Team Daniel V Senior - Aerospace Engineering, Payload Team University of Minnesota USLI FRR 2012 Page 2

3 Table of Contents 1. FLIGHT READINESS REPORT SUMMARY TEAM SUMMARY LAUNCH VEHICLE SUMMARY PAYLOAD SUMMARY CHANGES MADE SINCE CRITICAL DESIGN REVIEW CHANGES TO VEHICLE CRITERIA CHANGES TO PAYLOAD CRITERIA CHANGES TO ACTIVITY PLAN VEHICLE CRITERIA DESIGN AND CONSTRUCTION OF LAUNCH VEHICLE Overview Flight Reliability Confidence Component and Assembly Testing Workmanship Safety and Failure Analysis Full Scale Launch Test Results Mass Statement RECOVERY SUBSYSTEM Robustness of Recovery System Parachutes and Attachment Schemes Safety and Failure Analysis MISSION PERFORMANCE PREDICTIONS Mission Performance Criteria Flight Profile Simulations Validity of Analysis Stability Margins Kinetic Energy Management University of Minnesota USLI FRR 2012 Page 3

4 3.3.6 Altitude and Range with Variable Wind Conditions VEHICLE VERIFICATION Requirement Verification SAFETY AND ENVIRONMENT (VEHICLE) Safety and Mission Assurance Analysis Personnel Hazards Environmental Concerns PAYLOAD INTEGRATION Overview Compatibility Housing Integrity PAYLOAD CRITERIA EXPERIMENT CONCEPT SCIENCE VALUE Payload Objectives Payload Success Criteria Experiment Approach Experimental Test and Measurements, Variable and Controls Relevance of Expected Data and Accuracy/Uncertainty Preliminary Experiment Process Procedures PAYLOAD DESIGN Payload Design Precision of Instrumentation Workmanship Test and Verification Program PAYLOAD VERIFICATION SAFETY AND ENVIRONMENT (PAYLOAD) LAUNCH OPERATIONS PROCEDURES CHECKLISTS University of Minnesota USLI FRR 2012 Page 4

5 6. ACTIVITY PLAN BUDGET TIMELINE PARACHUTE EJECTION TEST FRIDAY, MARCH 30 TH, PAYLOAD EJECTION TEST SATURDAY, MARCH 31 ST, FULL SCALE FLIGHT SUNDAY, APRIL 1 ST, ANALYSIS OF FULL SCALE FLIGHT DATA TUESDAY, APRIL 3 RD, FRR ADDENDUM WRITTEN AND GIVEN TO USLI REPRESENTATIVES SUNDAY, APRIL 8 TH, LEAVE FOR HUNTSVILLE, AL WEDNESDAY, APRIL 18 TH, EDUCATIONAL ENGAGEMENT CONCLUSION APPENDICES APPENDIX I FLIGHT READINESS REVIEW FLYSHEET APPENDIX II DETAILED BUDGET TABLES University of Minnesota USLI FRR 2012 Page 5

6 List of Tables: Table 1: Electrical components for avionics bays Table 2: Featherweight Raven2 specifications Table 3: Entecore Electronics AIM USB altimeter specifications Table 4: PerfectFlite StratoLogger Specifications Table 1: Safety Analysis Payload Integration Table 2: Safety Analysis Airbrake Table 3: Safety Analysis Motor Table 4: Safety Analysis Fin System Table 5: Mass Balance Statement Table 6: RF Tracker Specifications for PT-1B Table 7: Parachute descent rates Table 8: Iris Ultra 144 specifications Table 10: Recovery Systems Failure Analysis - Electronics Table 9: Recovery Systems Failure Analysis - Hardware Table 11: Kinetic energy calculations Table 12: Kinetic energy calculations with maximum windspeed Table 13: Drift range with various wind speed Table 14: Safety and Mission Assurance Analysis... Error! Bookmark not defined. Table 15: Personnel hazards... Error! Bookmark not defined. Table 16: Control System Specifications Table 17: Data Collection system Specifications Table 18: BVGM-1 Specifications Table 19: RC-100X Specifications Table 20: S6020 Specifications Table 21 PT-1B Specifications Table 22: Temp/RH Probe Specifications... 88

7 Table 23: Silicon Pyanometer Specifications Table 24: Summary of Payload Failure Modes Table 25: Funding Sources and Amounts Table 26: Total Expenditures Table 27: On the pad vehicle costs Table 28: Half Scale Expenditures Table 29: Full Scale Expenditures Table 30: Payload Expenditures Table 31: Testing and Supplies Cost Table 32: Travel Expenditures University of Minnesota USLI FRR 2012 Page 7

8 List of Figures: Figure 1: CAD Image of Complete Vehicle Figure 2: Nosecone, payload piston and payload airframe CAD drawing Figure 3: Front Avionics Bay CAD drawing Figure 4: Upper booster tube and transition CAD drawing Figure 5: Coupler Section CAD drawing Figure 6: Lower booster section and removable motor mount tube CAD drawing Figure 7: Rear Avionics Bay CAD drawing Figure 8: Lower Booster complete assembly CAD drawing Figure 9: Airbrake isometric and top views in SolidWorks Figure 10: Firgelli L16 Linear Actuator Figure 11: Airbrake wiring schematic Figure 12:Load and Current curves for the L16 actuator provided by Firgelli Figure 13: Latching Relay Specifications Figure 14: Altitude profile found for the rocket without airbrakes (blue), with variable airbrake opening velocity (green), and instantaneous airbrake opening velocity (red).. 24 Figure 15: Velocity profile found for the rocket without airbrakes (blue), with variable airbrake opening velocity (green), and instantaneous airbrake opening velocity (red).. 25 Figure 16: This shows the force on each airbrake door for instantaneous brake deployment (red), and variable speed brake deployment (green) Figure 17: This shows the rocket deceleration without brakes (blue), with variable opening speed brakes (green), and instantaneously opening brakes (red) Figure 18: Copper ejection canister holder Figure 19: Aluminum ejection canister holder Figure 20: Aluminum ejection canister holder Figure 21: Main parachute U-bolt CAD drawing Figure 22: Rear Avionics Bay U-bolt CAD drawing Figure 23: Flight simulation plots... Error! Bookmark not defined.

9 Figure 24: Rocksim estimated altitudes... Error! Bookmark not defined. Figure 25: Motor thrust curve from Rocksim... Error! Bookmark not defined. Figure 26: Rocksim 2-D image Figure 27: Conceptual Drawing of Control System and Ground Station Figure 28: Automatic Orientation Correction System Figure 29: Rover Drawing With Control Component Locations Figure 30: Rover Drawing Showing Data Collection System Figure 31: Solidworks model of assembled rover Figure 32: As-built rover claw assembly. Note that the drive axle in this picture has not been trimmed and is longer than it will be in the final rover Figure 33: Deployed outrigger arm Figure 34: Wireless Camera Figure 35: Electronic Switch Figure 36: S6020 Servo Figure 37: SPMB4500NM Batter Pack Figure 38: SPMB4500NM Batter Pack Figure 39: PT-1B Transmiter Figure 40: Picture of the LRS Pro (Silver) mounted to RC Controller Figure 41: LRS Pro RC Receiver (Green) with Antennae Connected Figure 42: Electrical Schematic of Payload Control System Figure 43: Electrical Schematic of Data Collection System Figure 44: Electrical Schematic of Ejection/Camera Circuit Figure 45: HOBO Data Logger Figure 46: Data Logger Specifications Figure 47: HOBO Temp/RH Probe Figure 48: Solar Radiation Sensor Figure 49: A test of the HOBO data logging system University of Minnesota USLI FRR 2012 Page 9

10 University of Minnesota USLI FRR 2012 Page 10

11 1. Flight Readiness Report Summary 1.1 Team Summary Team Name: Project Name: Team Location: Team Officials: University of Minnesota Team Artemis Atmospheric Rover and Rocket Delivery System Minneapolis, Minnesota Dr. William Garrard - Faculty Adviser Gary Stroick - NAR/TRA Mentor 1.2 Launch Vehicle Summary Vehicle Dimensions Length: Diameter: Empty Vehicle Weight: Loaded Weight: Motor: Recovery: Rail Size: 126 inches inches payload section inches booster section 31.4 pounds 46.8 pounds (with motor and payload) CTI Pro75 L2375 Dual Deployment, redundant black powder ejection One and one-half inch button See Appendix for complete FRR flysheet. 1.3 Payload Summary The payload objective is to deploy a small, remotely controlled rover equipped with an array of sensors to collect temperature, relative humidity, and light intensity readings. The rover will also be equipped with a wirelessly transmitting camera, allowing the pilot to control the rover without visual contact. The purpose of this project is to simulate and explore the possibility of deploying small, inexpensive probes to extraterrestrial bodies in order to scout potential landing zones for more complex, large-scale missions.

12 2. Changes Made Since Critical Design Review 2.1 Changes to Vehicle Criteria A. Motor Changes The motor was changed back to the Cessaroni Technologies Pro75 L2375 motor. There were issues with our rail exit velocity and thrust-to-weight ratio with the motor selection stated in the Critical Design Review (CTI L1115), and this justified our change. Coincidentally, the two motors have the same dimensions and only differ in burn time and total impulse. It must be noted that the latest simulations with the L2375 and our current vehicle give us an apogee altitude of approximately 5800 feet AGL. Our airbrake system will be employed to slow our vehicle to apogee at 5280 feet. B. Airframe Changes The positioning of some of the bulkheads in the vehicle has changed. The nosecone bulkhead will be recessed further into the shoulder, shortening the length of the payload tube required. The construction of the bulkhead also has changed slightly. The piston payload bulkhead construction has also changed slightly. The rear avionics bulkhead and the front body tube centering ring have been combined eliminating the space between them and shortening the outer airframe length of the lower booster section. Extra support for bulkheads in the lower booster have been added in the form of coupler tubes epoxied to the airframe tube, between the bulkheads and centering rings. 2.2 Changes to Payload Criteria A. Circuit Construction As both the wirelessly transmitting camera and the electronic switch operate with a 9V battery, the circuit presented in CDR has been reconfigured to power both the camera and ignite the black powder charges with the same battery. Upon triggering the switch to close, the black powder charges will fire and eject the rover, while simultaneously powering up the camera initiating the video feed. This configuration saves on space and battery life, see Section 4.3 for more details. B. RC Transmitter Since CDR, range testing of the DX5e transmitter has shown the controller to be operable only within a range of 800 ft., far below the required range. To boost the signal, a new transmitter has been selected, the LRS Pro, see Section 4. The new transmitter works as slave to the DX5e and transmits commands at 0.5W at a frequency of 433MHz, significantly boosting the rover s operable range. University of Minnesota USLI FRR 2012 Page 12

13 C. Rover Drive Mechanism The rover wheels described in the CDR were determined to be insufficient to handle the difficult terrain expected upon landing, so the wheels have been replaced by an unfolding claw mechanism that will offer superior manoeuvrability in difficult terrain. The mechanism will be spring-loaded for automatic expansion upon ejection. Section C contains further details on the replacement drive mechanism. 2.3 Changes to Activity Plan A. Budget Changes The budget has changed little since the Critical Design Review. A few components on the payload and vehicle were upgraded since our on the pad costs were less than the maximum requirement of $5000. B. Timeline Changes The project timeline has changed since the Critical Design Review. We have currently pushed our scheduled full scale test flights back to April 1 st. This leaves zero margin for test flight failure. C. Educational Engagement Changes Since our last report we have done three outreach events. Two of them were science fairs at Galtier Middle School and the other at Lincoln Elementary School. The third event was an AIAA showcase event. With these three events we reached 282 people. University of Minnesota USLI FRR 2012 Page 13

14 3. Vehicle Criteria 3.1 Design and Construction of Launch Vehicle Overview The launch vehicle has not changed much since Critical Design Review. This report will touch briefly upon all vehicle systems, focusing on the structural and electrical elements. For more details about each system, please refer to the Critical Design Report. A. Vehicle Overview The CAD image below shows the complete assembled vehicle. It consists of three main sections tethered together for flight configuration. The vehicle itself consists of eleven modular sections that will be assembled to form the complete flight vehicle. The modular design was selected for ease of replacement of parts as well as for vehicle versatility. Figure 1: CAD Image of Complete Vehicle For flight configuration, the vehicle will consist of the forward payload assembly, the coupler section and the booster/fin can assembly which also houses the airbrake system.

15 B. Forward Payload Assembly The forward payload assembly consists of seven independent sections, not including the payload itself. It extends from the nosecone to the coupler. It contains the nosecone assembly, the payload airframe tube, the payload ejection piston assembly, the front avionics bay, the airframe transition, the upper booster tube section, and the main parachute ejection piston. Nosecone The nosecone is constructed of fiberglass and purchased from Public Missiles Limited. It is hollow and will house one of the vehicle RF trackers. It has a bulkhead epoxied within the shoulder that will be used to contain the payload within the airframe tube. The nosecone will be shear pinned into place to survive the rigors of flight, but must be ejectable to allow the payload to exit the vehicle when deemed safe. Figure 2: Nosecone, payload piston and payload airframe CAD drawing Payload Airframe The payload airframe is constructed of a fiberglass wrapped phenolic tube. The fiberglass wrap was selected to provide extra rigidity for payload contained within, as well as provide strength to withstand the large ejection charges that must be triggered to push the payload out of the vehicle. To separate the payload system from the rest of the vehicle, there is a separation bulkhead constructed of birch plywood epoxied at the rear of the payload airframe tube. University of Minnesota USLI FRR 2012 Page 15

16 Payload Ejection Piston Within the payload airframe, the payload will sit between the nosecone and a payload ejection piston which is tethered to a bulkhead. As with a typical parachute ejection system using a piston to protect the parachute from the black powder, we will use a piston to eject the payload from our vehicle. The piston is made of a phenolic tube, with a custom bulkhead epoxied in place. Front Avionics Bay The front avionics bay sits opposite the payload ejection piston on the backside of the separation bulkhead. It sits inside an inner sleeve that is also used to join the smaller diameter booster tubes to the larger diameter payload tube. The sleeve that the front avionics sled sits in is epoxied into a channel in the separation bulkhead. Two aluminum rods extend outward from the separation bulkhead along the length of the sleeve and are used to attach the avionics bulkhead to the sleeve, sealing the bay. Figure 3: Front Avionics Bay CAD drawing Upper Booster The upper booster is a 6 inch diameter phenolic tube, and slides over top of the avionics sleeve, once it is completely assembled and ready for flight. It is riveted in place when the front end of the booster is flush against the separation bulkhead. This will ensure the loads are transmitted over the airframe and not the sleeve, which is used to provide support against shear and torsion. University of Minnesota USLI FRR 2012 Page 16

17 Inside the booster will sit the main parachute ejection piston and the main parachute and shock cord. The main piston will be attached to the avionics bulkhead by a shock cord and quick link. Figure 4: Upper booster tube and transition CAD drawing Airframe Transition The airframe transition is constructed of fiberglass, and is actually a fiberglass nosecone which has been cut and modified to serve as an external transition for airflow from the larger diameter tube to the smaller diameter tube. It is fully a cosmetic/airflow device and does not carry any loads. It is designed to slide over the upper booster, and will be riveted (by a small shoulder) to the payload airframe. C. Coupler Section The coupler section is a smaller assembly joining the upper and lower booster tubes. The coupler section also contains the bulkhead responsible for separating the main and drogue parachute bays. It is one piece, consisting of a phenolic coupler tube with a one inch airframe band epoxied to the exterior of the coupler. It has a single birch plywood bulkhead epoxied to the forward end of the tube closest to the main parachute. On each side of the bulkhead u-bolts are attached to allow main and drogue shockcords to be attached. University of Minnesota USLI FRR 2012 Page 17

18 Figure 5: Coupler Section CAD drawing The upper and lower boosters will be shear pinned to the coupler tube when the airframes are butted up flush against the airframe band. The shear pins will prevent the boosters from separating prematurely in flight (and contain the parachutes until the proper time to deploy). D. Lower Booster Assembly The lower booster assembly consists of four independent sections. The main section is the booster airframe tube itself. Attached to this airframe are the rear avionics bay, the removable fins, and the removable motor mount system (which contains a removable boat tail section which contains the vehicles second RF tracker). Airframe Tube The lower airframe is constructed of phenolic tubing. It is heavily supported with inner coupler tubes epoxied along the length. University of Minnesota USLI FRR 2012 Page 18

19 Figure 6: Lower booster section and removable motor mount tube CAD drawing Rear Avionics Bay It is conveniently located on a phenolic tube the same diameter as the motor mount tube (instead of on a sled). This allows the vehicle to fly with a higher grain motor if deemed necessary, with the motor extending forward out the motor mount tube, under the rear avionics. Figure 7: Rear Avionics Bay CAD drawing University of Minnesota USLI FRR 2012 Page 19

20 Removable Fin System The fins are constructed of 3/16 inch G10 fiberglass. They are attached via pins to the fin mount system. The fin mount system is also constructed of G10 fiberglass, and is bolted to the airframe. Figure 8: Lower Booster complete assembly CAD drawing Removable Motor Mount System The removable motor mount system was designed and built with multiple purposes. It allows the vehicle to fly with a larger diameter motor than the one selected for this competition, giving the vehicle versatility and flexibility in flight parameters and missions. The removable motor mount system also allows access to the airbrake system for repairs and troubleshooting if needed. Since it is removable from the airframe, it also allows access to the fin mount system from inside of the airframe. Airbrake System The airbrake system is designed to increase the control of the vehicle between launch and apogee and is located aft of the fins. It is composed of three linear actuators connected to three doors which open at a user set altitude to increase drag and slow the airframe down to reach one mile. At apogee, the doors will close to provide a more robust orientation on landing. A drawing showing the motor mount tube and the attached air brakes can be seen below. University of Minnesota USLI FRR 2012 Page 20

21 Construction and Sizing Figure 9: Airbrake isometric and top views in SolidWorks The airbrake doors are composed of phenolic tube reinforced by layers of fiberglass on either side and close flush with the airframe when not in use. The hinges to attach the doors to the bulkhead will be machined out of 6061 T6 aluminum. We used west systems epoxy with an adhesive strength of 1500 psi for the aluminum hinges. Door sizing is set to 6 inches. Nylon cord will be used to connect each door with their respective linear actuator. The actuators will be mounted to the motor mount tube, and the actuator arm length is 140mm using Firgelli L16 linear actuators. A picture of the Firgelli L16 can be seen below. A gear ratio of 63:1 will be used giving a maximum lifting force of 100 N. University of Minnesota USLI FRR 2012 Page 21

22 Wiring and Component Specification Figure 10: Firgelli L16 Linear Actuator The three linear actuators will be wired in parallel and connected to a power supply separate from the avionics bay. Limit switches within the actuators will be used to mitigate damage to the actuators and prolong actuator life. A double coil DPDT latching relay will keep the actuators closed until a signal from an altimeter in the aft avionics bay is applied to the DPDT latching relay at a user set altitude which will open the actuators. A second signal will be sent to the opposing coil of the relay which will cause the actuators to close again. The circuit schematic can be seen below. The load and current curves for the actuator and the specifications of the latching relay can also be seen. University of Minnesota USLI FRR 2012 Page 22

23 Figure 11: Airbrake wiring schematic Figure 12:Load and Current curves for the L16 actuator provided by Firgelli Relay Part Number TX2-LT-9V-TH Relay Type Latching, Dual Coil Coil Current 15.5 ma Coil Voltage 9V Contact Form DPDT Contact Rating 2A Operate Time 4 ms Release Time 4 ms Figure 13: Latching Relay Specifications Airbrake Analysis University of Minnesota USLI FRR 2012 Page 23

24 Altitude (ft) Our airbrake analysis was done using MATLAB and estimating the drag forces on the airbrake doors using a flat plate assumption for drag. This estimation has been previously verified by our team lead on a rocket employing a similar airbrake door. Figures below show the projected altitude, velocity, force per airbrake door, and rocket deceleration Altitude Profile Time (sec) Figure 14: Altitude profile found for the rocket without airbrakes (blue), with variable airbrake opening velocity (green), and instantaneous airbrake opening velocity (red) The figure above shows three profiles for altitude of our rocket. The highest altitude, shown in blue, shows our final predicted altitude of 5800 ft without airbrake deployment. The lowest altitude, shown in red, is the predicted altitude of the rocket if the airbrakes opened instantaneously while the green line shows the most accurate estimation encoding for the variable opening speed given by the L16 linear actuators. These profiles were calculated using a 6 inch door length and a deployment altitude of 4350 ft. While a linear actuator speed of 20 mm/s was used in calculating the green line, the actual opening speed will be faster with the applied load from drag. We plan to test the actual opening speed of the linear actuators with the estimated applied loads when we order them. For now, an accurate estimation of final altitude would be found in between the red and green lines above giving a projected final altitude with airbrake deployment between 5100 ft and 5300 ft. The projected rocket velocity versus altitude is shown below. University of Minnesota USLI FRR 2012 Page 24

25 Velocity 350 Velocity Profile Time (sec) Figure 15: Velocity profile found for the rocket without airbrakes (blue), with variable airbrake opening velocity (green), and instantaneous airbrake opening velocity (red) University of Minnesota USLI FRR 2012 Page 25

26 Force (lbf) 50 Force per Brake Time (sec) Figure 16: This shows the force on each airbrake door for instantaneous brake deployment (red), and variable speed brake deployment (green) The figure above shows a range of possible forces applied to the airbrake doors in different scenarios. The red line shows the maximum force applied to a brake door if it were to instantaneously open when deployed. This scenario could occur if the actuator or mount broke at the beginning of its deployment. The forces applied to the airbrake doors when the speed of the actuators is applied gives a maximum force of 17 lbf while the linear actuators are capable of actuating 22.5 lbf. It must also be noted that this is the maximum force which the actuators are capable of actuating but not their maximum breaking load. The actuators are capable of loads greater than this without damage when the load is applied in the direction of motion. University of Minnesota USLI FRR 2012 Page 26

27 Deceleration (g) 5 Deceleration Time (sec) Figure 17: This shows the rocket deceleration without brakes (blue), with variable opening speed brakes (green), and instantaneously opening brakes (red) This analysis is conclusive that the decelerations from the airbrakes will not cause separation of the airframe due to shear pins breaking. AVIONICS SYSTEM Overview There are two avionics bays in the rocket design. The front avionics bay will provide signals for main parachute deployment as well as payload deployment. The aft avionics bay will provide signals to deploy the drogue parachute and to control the opening and closing of the actuator arms for the vehicle s airbrake system. For the recovery system each avionics bay will contain two altimeters, two power sources, and two on-off switches (Table 20) in compliance with the USLI guidelines for redundancy. However, the rear avionics system will carry an extra Raven2 altimeter to signal the airbrakes. University of Minnesota USLI FRR 2012 Page 27

28 Components Quantity Featherweight Raven2 Altimeter 2 PerfectFlite StratoLogger Altimeter 1 G-Wiz LCX Flight Computer 1 9 Volt Battery 4 Screw Switch 4 Table 1: Electrical components for avionics bay Figure 18: Featherweight Raven2 Altimeter Manufacturer Featherweight Altimeters LLC Axial Accel range and frequency 70 Gs, 400 Hz Lateral Accel frequency 35 Gs, 200 Hz Download Interface USB Power Supply 9 volts DC Max output 9 amps Size 0.8 x 1.8 x 0.55 Table 2: Featherweight Raven2 specifications University of Minnesota USLI FRR 2012 Page 28

29 Figure 19: Entecore Electronics AM USB altimeter Manufacturer Outputs Power Supply Output Current Dimensions Entecore Electronics Apogee and Main programmable 6V to 14V 4 amps 95mm x 25mm x 15mm Table 3: Entecore Electronics AIM USB altimeter specifications Figure 20: PerfectFlite Statologger Altimeter Manufacturer PerfectFlite Outputs Apogee and Main programmable Power Supply 4V to 16V, 9V nominal Max Output 9 amps Main Deployment 100 AGL to 9,999 AGL Dimensions 2.75 x 0.9 x 0.5 Table 4: PerfectFlite StratoLogger Specifications University of Minnesota USLI FRR 2012 Page 29

30 Orientation and Wiring The altimeters will be oriented axially to ensure accurate altitude results and the batteries will be oriented with the terminals facing the ground. Magnetic switches will be implemented to be turned on from the exterior of the airframe at launch. The wiring schematics are shown below for each avionics bay. Figure 21: Front avionics bay University of Minnesota USLI FRR 2012 Page 30

31 Figure 22: Aft avionics bay Figure 23: Magnetic Switches from Aerocon Systems The magnetic switch offered by Aerocon Systems has an On/Off activation range of 1 with a small rare-earth magnet. It has a current capability of 12 amps continuously with an input voltage of 3.5 to 16V. The current draw is under 3 micro-amps and it is comprised of 100% solid-state surface-mount components which will not be affected by accelerations throughout flight as shown by our half scale launch. University of Minnesota USLI FRR 2012 Page 31

32 POWER ANALYSIS For the airbrake system a latching relay which will be able to be switched on and off using the pyro channels from the Raven2 and their standard pulse setting. Testing of the Raven 2 altimeters showed that there was a continuity voltage equivalent to the power supply and a continuity current of 0.5 ma. A standard Duracell 9V battery is rated to 580 mah which means the current supplied when the brakes are not in use will still give a lifetime well within the limits of our 1 hour on the pad requirement. The other altimeters are also well within the recommended operating conditions and as such will also fall within the USLI requirements Flight Reliability Confidence At the moment we are confident with our design however we are still waiting to complete our full scale test flight to verify our calculations Component and Assembly Testing We completed shear pin tests with the nylon key hole, 2-56, and 4-40 shear pins. With no reinforcement, the key hole shear pins failed to shear and instead tore through the phenolic tube. While the 2-56 created less damage, it was still enough that we decided to coat all shear pin phenolic faces with epoxy. The epoxy reinforcement was sufficient for all tests that followed. We hung weight the phenolic tube to test shear pin strength. We found the 2-56 shear strength to be roughly 25 lbs while the nylon 4-40 shear pins were around 35 lbs. Before our sub scale flight test we tested all altimeters in a vacuum chamber which gave positive conclusive results for our barometric sensors. Each altimeters drogue and main lines fired properly and we had visual confirmation using an LED. We performed five half-scale drogue parachute ejection tests. The first three tests were done using 1.7 grams of FFFg pyrodex because we did not need our mentor to be present, as opposed to if we used black powder he would have. The initial test initially appeared successful, but later we discovered that one of the three shear pins tore the phenolic. To confirm this was not a fluke, we performed the testing again and found similar results. At this point we decided to switch from keyhole shear pins to standard 2-56 shear pins and use 1.4 grams of FFFFg black powder because the pyrodex showed to burn differently. The fourth test resulted in slight tear again. At this point we decided to switch to a friction fit and tested an ejection with.6 grams of black powder and it was successful. University of Minnesota USLI FRR 2012 Page 32

33 We performed three main parachute ejection tests. The first two tests we used 2.5 grams of FFFg pyrodex and resulted in a very, very forceful ejection. This test slightly damaged the ejection piston and tore the phenolic tube in two of three shear pin holes. We repaired this and ran a second test, this with 1.8 grams of pyrodex. This test resulted in the charge blowing a hole in the rocket tube. We attributed this to a tight parachute pack and the charge settled directly against rocket tube. To solve these problems we made a second main parachute bay and made canister holders that prevented the charge from facing the rocket tube. The third test used 1.0 grams of FFFFg black powder with the canister and was a success. We performed ten payload ejection tests using a mock payload to guarantee that the payload will eject in any orientation, using the rover controller to remotely trigger the black powder charge. The first two tests used 1.0 grams of black powder with the payload bay orientated horizontally and nose up at a 45 angle with both being successful, but very high energy. The next two tests were performed with.85 grams of black powder, both with nose down at a 45 angle. The first of these tests resulted in the nose cone bulkhead epoxy fillet breaking and the payload getting stuck in the nose cone. To fix this we made a better, bigger epoxy and the second test proved successful. During this test the molex connector on the mock payload was damaged and went unnoticed for the next three tests. The next three tests used 1.1 grams of black powder and this was to show that the mechanical switch on the camera would not accidently switch to the off position on the full-scale. 1.1 grams of black powder allowed the rover to undergo the maximum forces expected during the full-scale launch, just over 20g. The first two tests were unsuccessful because the charge did not fire, once the molex connector was replaced the test was successful. The next two tests were performed to test unlikely, worst-case scenarios for how the rocket can land, nose directly up and nose directly down. Both tests were done with.85 grams of black powder and both were successful. The last test performed was to test if the rover would escape from the payload bay even if the ejection piston severely cocked in the bay. This was guaranteed by cutting the piston to 2.9 in. in a 3.9 in. tube. This test was a success, the piston moved about 3 in. inside the tube, but the rover exited the bay unimpeded. With these tests we know we only need to do two tests on the full-scale and know the amount of black powder required is roughly 85% of theoretical calculations require. The final test on the half-scale was a test to see if our new method to prevent the phenolic from ripping from shear pins was valid. The new method involved putting an epoxy layer on both the inside and outside of the coupler tube that the shear pins would go through. The phenolic is porous and allows the epoxy to soak in, thru-hardening the material. Doing this changed the diameter of the coupler tube very little and does not require sanding to fit inside of the airframe tubing. We did a standard parachute ejection test and the shear pins broke as they are intended to break and did not damage the phenolic. All tests involving black powder were performed under the supervision of our mentor. University of Minnesota USLI FRR 2012 Page 33

34 3.1.4 Workmanship Many of our components were purchased off the shelf, and did not require much workmanship. Some of the components did require modifications for our purposes however, and there was time spent investigating the best techniques to work with phenolic, birch plywood and other materials for cutting, drilling and finishing. Some of our components were made from raw materials as we could not find a suitable off the shelf component to satisfy our requirements. Many of the bulkheads and centering rings are such components, as well as the majority of the airbrake hardware. Care was taken to machine such parts properly to specifications laid out by Solidworks drawings so that all components would fit together properly in final assembly. The team lead assisted and inspected each part that the team members made for the subsystems, maintaining a high quality of workmanship. During machining, our NAR/TRA mentor was also present much of the time, to supervise and assist in proper machining techniques. Prior to final assembly, many of the components were machined and test fitted prior to epoxy or bonding to the other components. This would ensure that we did not end up overlooking minor details such as hole alignment and fit and finish. Since our vehicle was designed to be modular, care was taken to ensure all of the fillets and surfaces of the components were visually appealing, functional and easily repaired, replaced and/or cleaned. For all bonding of components using epoxy either the team lead or the NAR/TRA mentor was present to supervise. The most critical elements of workmanship involve proper bonding of components to optimize structural integrity, and precision of machining of components to allow multiple components to function as a subsystem properly Safety and Failure Analysis The launch vehicle itself has many failure modes since it comprises many components and subsystems. We have considered all failure modes possible and in the sake of safety, have compiled a table of mitigations. The tables are divided into vehicle systems mitigations. Vehicle System: Payload Integration Risk Probability Impact Mitigation Nosecone separation in Low Nosecone free falls from altitude, payload Detailed calculations have been completed to confirm proper University of Minnesota USLI FRR 2012 Page 34

35 flight separates from vehicle amount of shear pins in nosecone Nosecone jams upon deployment Low Payload fails to deploy Use the proper shoulder length, and guarantee that the applied force to separate is evenly distributed Piston jams upon deployment Low Payload fails to deploy Using the proper piston length for the tube involved, and making sure the inner tube is clean and free of contamination Piston bulkhead fails Low Piston, payload and airframe become damaged Use proper epoxy techniques to ensure a rigid bond, and ground testing to find failure load limit Table 5: Safety Analysis Payload Integration Vehicle System: Airbrake Risk Probability Impact Mitigation Electrical failure of actuators or relays Low Airbrakes fail to deploy or deploys at improper altitude Ensure altimeter is functioning and programmed properly, and there is electrical continuity to actuators Mechanical failure of hinge assembly Low Airbrake fails to open or door separates from vehicle Calculations to prevent cable separation, or hinge failure Drag force pulls MMT out of body tube Low Motor assembly and airbrake assembly free fall to ground Use lots of epoxy, and open door slowly One of airbrake doors jam closed Low Vehicle loses stability in flight and tumbles Ensure doors are free to move prior to flight Table 6: Safety Analysis Airbrake University of Minnesota USLI FRR 2012 Page 35

36 Vehicle System: Motor System Risk Probability Impact Mitigation Body centering rings fail to contain motor assembly Low Motor casing rips through rocket in flight Proper epoxy attachment of CR s and calculations of max structural loads on CR s Motor mount centering rings fail to contain motor assembly Low Motor casing rips through rocket in flight Proper epoxy attachment of CR s and calculations of max structural loads on CR s Motor retention system fails Low Motor casing free falls from altitude Ensure motor casing is properly retained with a pre-fabricated motor retainer Propellant fails to ignite Low Mission failure, vehicle fails to leave pad Proper ignition system setup Propellant burns through casing Low Loss of vehicle stability, vehicle damage Inspection and care of motor casing system to prevent damage pre-flight Propellant explodes upon ignition Low Loss of vehicle, personnel hazard Proper assembly of motor, and inspection of casing and assembly Table 7: Safety Analysis Motor Vehicle System: Fin System Risk Probability Impact Mitigation University of Minnesota USLI FRR 2012 Page 36

37 Fin mount fails to stay attached to body tube Low Fin and fin mount free fall from altitude, rocket loses stability Ensure proper attachment of mount to body tube using epoxy techniques and screws Fin fails to stay in fin mount Low Fin free falls from altitude, rocket loses stability Ensure fins are properly and snuggly attached to fin mount Fin not designed strong enough Low Fin flutters causing unstable flight Analysis to determine proper fin thickness Full Scale Launch Test Results Table 8: Safety Analysis Fin System Our full scale launch test results are pending our full scale test flight. It is scheduled for March 31-April 1. We will submit an addendum to our FRR following flight tests. We will have two full scale test flights. The first flight will be without airbrake deployment, and will be a general systems test as well as flight stability and Rocksim verification. Pending the results of the first test flight, the second test flight will be an airbrake systems test flight. Upon updating the Rocksim simulations (if needed) we will employ our airbrakes to attain an apogee altitude of 5280 feet AGL. Both flights will also serve to give the team practice at a full vehicle systems assembly, simulating competition requirements Mass Statement At this stage, our mass statement no longer relies on weight estimates of components from simulations or calculations. Prior to vehicle assembly, after each component was manufactured, the individual weights were recorded. The itemized vehicle component weights can be seen in the following table: Component Qty Length (in) Weight (oz) Margin Max Wt Nosecone Transition University of Minnesota USLI FRR 2012 Page 37

38 Boattail Payload Airframe Tube Booster Airframe Tube Booster Airframe Tube Coupler Band Trailing Airframe Tube Payload Piston Tube Av Sleeve Main Piston Tube Coupler Tube Av Outer Tube Fin Sleeve Av Tube Motor Mount Tube Nosecone Bulkhead Payload Piston Bulkhead Separation Bulkhead Front Av Bulkhead Main Piston Bulkhead Coupler Bulkhead Rear Av Bulkhead Body CR MMT CR Body CR MMT CR MMT CR MMT CR Main Parachute Main Shockcord Drogue Parachute Drogue Shockcord U-bolt (Small) U-bolt (Large) U-bolt (Medium) Quicklink (Small) University of Minnesota USLI FRR 2012 Page 38

39 Quicklink (Large) Quicklink (Medium) Threaded Rods Aluminum Canisters Rivet Nuts Primary Front Alt Secondary Front Alt Primary Rear Alt Secondary Rear Alt Airbrake Alt Volt Battery Battery Holder Magnetic Switches Rf Tracker Front Avionics Sled Surface Fin Mounts Fins Actuators Airbrake Hardware Airbrake Doors Motor Casing Motor Motor Retainer Payload Paint/Epoxy Estimate Total Components 110 Total Length (in) Total Weight (oz) Maximal Weight (oz) Table 9: Mass Balance Statement University of Minnesota USLI FRR 2012 Page 39

40 Prior to assembly, it was difficult to estimate the amount and weight of epoxy required to assembly the vehicle. It was also difficult to estimate the amount and weight of paint also required. However, by weighing the subsystems after assembly, the amount of paint and epoxy could be determined considering we know the weights of each component in the various sub-assemblies. Maintaining the mass statement was crucial during the design, manufacture and assembly of the vehicle since the sum of the component weights affected our flight performance and apogee altitude achieved, as well as affected our static stability margin. With a proper mass statement, our simulations could be accurately maintained and minor modifications could be made to the design as required. In summary, there are three combined weights that we are concerned with: 1. Fully Loaded Weight 46.5 pounds (weight of vehicle, payload and motor) 2. Recovery Weight 41.4 pounds (weight of vehicle, payload after motor burnout) 3. Empty Vehicle Weight 31.4 pounds (weight of vehicle without payload and motor) The fully loaded weight is most critical, since this will determine our apogee altitude. Currently, we are targeting an apogee altitude of 5700 feet AGL without the use of our airbrake system. University of Minnesota USLI FRR 2012 Page 40

41 3.2 Recovery Subsystem Robustness of Recovery System Testing on the half-scale showed two major flaws in the phenolic during black powder ejections, both caused by phenolic being much weaker than G10, our initial rocket tube material. First is that if the charge is directly facing the wall of the rocket tube, it has the capability to tear a hole through the tube. To fix this problem we created canister holders so the charges were orientated parallel to the rocket s axis. On the half-scale we simply epoxied a brass tube to the aviation bay bulkhead. The canisters would sit in here with the wires running through the bulkhead to the altimeters, shown in Figure 24. The holes were made just big enough for the wires so little to no black powder will escape through them. This design prevented all further issues with the charge itself damaging the phenolic tubing. Figure 24: Copper ejection canister holder The design flaw with this design is the epoxy breaks or fractures often during test and one did come apart in the half-scale launch, assuming it broke upon landing. The canisters broke apart from the bulkhead because the only connection it had was the epoxy. Therefore on the full-scale the design has been improved. The new design now has the canister holders sitting inside of the bulkhead to provide support with an epoxy fillet opposite of the flared end of the container, shown in Figure 25.

42 Figure 25: Aluminum ejection canister holder The new canisters are made of a single piece of aluminum with a depth hole lathed out to allow the canister to sit inside it. The closed end of the tube has holes cut into it to allow the wires to feed out and the open end of the container is flared. This flare will prevent the holder from moving through the bulkhead Figure 26 shows a standalone picture of the holder. Figure 26: Aluminum ejection canister holder The second major flaw of the phenolic tubing is that the shear pins have a tendency to tear it during ejections. In previous designs we planned on using keyhole shear pins, but those heavily damaged the tubing. Static testing showed this happened because the pins did not simply break, they first bent and this is when they damaged the tube. The orientation of the pin also effected how it sheared. Therefore these shear pins were considered unreliable and we switched to standard 2-56 and 4-40 shear pins. The shear pins had the same effect the keyhole shear pins, but to a lesser extent. To solve this problem we lined the inside and outside of the coupler tube where the shear pins would be placed with epoxy. The phenolic is porous and allows the epoxy to soak in, thru- University of Minnesota USLI FRR 2012 Page 42

43 hardening the material. Doing this changed the diameter of the coupler tube very little and does not require sanding to fit inside of the airframe tubing. Ejection tests were performed and showed phenolic tearing issues were solved. We have changed a few components of the recovery subsystem; we have selected smaller and lighter parts than those proposed in the CDR. After evaluation, we realized that we do not need to have such high work load limits since a large force will be exerted for a very short period of time when a main parachute opens. When a drogue chute deploys, the force will not be excessive because this event will be happening at apogee. If a chute accidently opens at other than apogee, other parts of the recovery subsystem such as epoxied parts will be destroyed before breaking U-bolts and quick links; we do not need to think about parts that can withstand at emergency situations. We need to think about how we can make a lighter vehicle. The dimensions of the U-bolts on the front AV bay, and on the main piston are following: Figure 27: Main parachute U-bolt CAD drawing The workload limit for this U-bolt is 1090lbs. The dimensions of U-bolts on the coupler bulkhead and on the rear AV bay are following University of Minnesota USLI FRR 2012 Page 43

44 Figure 28: Rear Avionics Bay U-bolt CAD drawing Quicklinks for the front AV bay and the main pistons are 1580lbs of workload limit made of zinc. Quicklinks for the rear AV bay and coupler bulkhead are 660lbs of workload limit and made of stainless steel. Before launching, we would like to test the strength of U bolts and quicklinks for the main piston with a tensile machine. From the test results, if we feel these parts are not strong enough to withstand the force which will be exerted, we will use the parts which we proposed in CDR. Both the drogue parachute bay and main parachute bay shall have three 4-40 shear pins in place to prevent the rocket from breaking apart during break open. Worst case scenario for break open is 4g and with the upper payload section weighing 25 pounds the shear pins must be able to support 100 pounds of force on them. With these shear pins and volume for each bay 1.2 grams of FFFFg black powder are required to break the main parachute bay shear pins and.6 grams are required to break the drogue parachute bay shear pins. RF Transmitter RF transmitters we are using for locating our rocket are same as one used for payload. One RF tracker will be located in the nosecone. The nosecone is not tethered to the rest of the vehicle, so it will have its own transmitter if it separates from the vehicle prematurely. One RF tracker will be located in the boat tail. The boat tail is attached to the rest of the rocket, and all other sections are tethered to one another. University of Minnesota USLI FRR 2012 Page 44

45 Both transmitters were positioned as far away as possible from all of the vehicle onboard recovery electronics as well as the payload electronics. They were also positioned in opposing ends of the vehicle in case of a complete vehicle failure the trackers will cover the most area. The frequency, wattage and range of our RF trackers are summarized below: Tracker Frequency Wattage Range Nosecone MHz 1 milliwatt 2 miles Boat tail MHz 1 milliwatt 2 miles Table 10: RF Tracker Specifications for PT-1B For ease of use we chose sequential tracker frequencies. The range is more than sufficient given competition requirements, and will suffice in worst case scenario if our main parachute deploys at apogee. For protection of av bay electronics we plan on coating the av bay bulkheads with copper shielding tape. Drogue Parachute Selection We are using Rocketman Mach2 high-speed drogue chute 3 ft diameter. Since we are using a big main chute, we need to eliminate the descent rate before the deployment of a main chute, so that we can limit the G-force experienced by the other parts of the recovery system. According to Rocksim, its descent rate is 77.3ft/s. Calculated values from a table published by Rocketman Chute for this chute size was 71.21ft/s. The descent rates for the drogue chute size are shown in the table below. If we choose a bigger drogue chute, the landing site from the launching point will be increased, and there will be a possibility that we will not be able to meet the half mile landing requirement. Chute diameter (ft) descent rate (ft/s) University of Minnesota USLI FRR 2012 Page 45

46 Table 11: Parachute descent rates Main Parachute Selection We will be using Fruity Chute Iris Ultra 144. The specification of the chute is following. Diameter Chute Weight Open Diameter, In Spill Hole, In Flattened Diameter in Open Area, Ft^2 Spill Area Ft^2 Effective Drag Area, Ft^ oz Table 12: Iris Ultra 144 specifications We proposed Fruity Chute Iris Ultra 120 in CDR. However, we knew that the chute is close to breaking the kinetic energy requirement at weight we had for the vehicle. If the weight increased, we will not be able to meet the kinetic energy requirement, so we increased the chute size. As long as we have a big chute, we will be always able to shrink down the size of the chute by reefing. There are different ways of reefing the chute, and we will decide how the chute will be reefed after launching the full scale. Currently, the heaviest part of the vehicle is forward section, which is 21 lbs of mass. According to the result of kinetic energy calculation, at landing, the velocity needs to be ft/s just before breaking the 75ft-lbf kinetic energy requirement Parachutes and Attachment Schemes There shall be two parachute deployments, drogue at apogee and main at 500 feet. Both will be a black powder ejection used to break four 4-40 shear pins shear pins are required instead of 2-56 shear pins because the brakes may cause up to a 5G force on the rocket. Attachment Scheme To protect chutes from burning, we are using a chute protector for a drogue chute and a piston for a main chute. A drogue chute will be connected to a chute protector. A chute protector will be connected to 5/8 tubular nylon shock cord. A shockcord is the three times the length of the rocket. The quicklinks will be used to connect the shockcord to the bulkheads on coupler and the rear AV bay. A main chute will be connected to a shockcord, and this will be connected to the quicklinks, and then U-bolts on a piston and a coupler. The length of shockcord we are using is the three times the length of the rocket. University of Minnesota USLI FRR 2012 Page 46

47 3.2.3 Safety and Failure Analysis For the recovery system, a safety and failure analysis was essential to ensure a safe recovery of the vehicle and minimize the chance of failure. Vehicle System: Avionics System Risk Probability Effects Mitigation Battery failure Low Altimeters cease operating in flight, fail to provide enough power to ejection charges acceleration Altimeter failure Incorrect pressure readings Altimeters programmed incorrectly Ensure correct orientation of battery terminals, fresh batteries every flight, and holder strength prevents movement under Low Parachutes fail Primary and secondary systems are onboard to ensure ejection charges and electronics will function Low Low Incorrect airbrake and parachute deployment altitudes Parachutes may fail to deploy We need to properly size the vent holes for each avionics bay and test it during full scale flight tests Use altimeters from different manufacturers to ensure the altimeter programming failures are not duplicated Table 13: Recovery Systems Failure Analysis - Electronics Vehicle System: Recovery Hardware Risk Probability Effects Mitigation Charges Ignite Pre-launch Med Personnel Hazard We do not put magnet close to the magnetic switch Charges Fail to Ignite Med Parachutes fail to deploy There is redundancy for an altimeter and a black powder charge Charges Ignite Prematurely Low Parachutes deploy on ascent There is redundancy for an altimeter Charges ignite simultaneously Low Possible structural damage to airframe or parachutes Proper wiring and programming University of Minnesota USLI FRR 2012 Page 47

48 BP residue contaminates avionics Insufficient black powder charge Harness linkage fails Bulkhead attachments fail Shock cords tangle in parachute Shroud lines tangle in parachute Improper parachute selection Parachute shreds Improper length of shock cord Improper shock cord selection Low Low Low Low Low Med Low Altimeters cease to operate correctly Parachutes fail to exit the airframe Vehicle tethers come apart from vehicle Vehicle tethers come apart from vehicle Parachute does not open correctly Parachute does not open correctly Descent rate too high, too much kinetic energy on landing The avionics bays must be properly sealed to protect from BP residue, ejection canisters must fit properly and not move or fail Calculations will be essential, and ground tests will be conducted, redundant charges The calculation of maximum opening force of chute has been performed and all components have been chosen based on the results We will test the attachment by applying the maximum opening force applied to the bulk head before putting into the body We are going to hold parachutes using a proper procedure We are going to hold parachutes using a proper procedure Hand calculations have been done, and the results have been verified with Rocksim simulation Low Descent rate too high We bought one of the strongest chutes Low Shock cord zippers The length of the shock cord has airframe upon been chosen by the rule of thumb; deployment three to four times the length of the rocket Low Shock cord snaps The shock cord was chosen based on the result of maximum opening force calculation Table 14: Recovery Systems Failure Analysis - Hardware University of Minnesota USLI FRR 2012 Page 48

49 University of Minnesota USLI FRR 2012 Page 49

50 3.3 Mission Performance Predictions Mission Performance Criteria The rocket must travel to an altitude of 5280 ft. Through Rocksim simulation and Matlab coding, the full scale rocket is predicted to travel to an altitude between 5050 ft. and 5280 ft with airbrake activation. Drogue and main parachutes must be successfully deployed on decent, and the rocket must provide an adequate housing for our payload. The drogue should be deployed at apogee, while the main parachute should be deployed at an altitude of 700ft. The rocket shall return within a 2500 ft. radius of the launch site. After a successful touchdown our team will receive clearance to deploy our rover. Our rover will then survey the land and take scientific measurements. These measurements will be recorded and later analysed by the team. A. MOTOR MOUNT SYSTEM The motor system must be capable of withstanding an impulse of up to 4905 N-sec and a maximum thrust of up to 2700 N. The motor mount system must also be removable so as to allow access to the linear actuators of the airbrake system for maintenance. There must also be enough room for the motor mount tube to be interchangeable between an inner diameter of 75mm and 98mm. The time to switch motor mount tubes must not impede on the USLI requirement of 2 hours for entire rocket assembly. B. FIN SYSTEM Our fin system must be robust enough to withstand the forces from launch. Each fin must also be removable in the chance that one becomes damaged upon landing. The removability of the fins will also give the capability to alter the fin design allowing for the rocket to carry various payloads in future missions. C. AIRBRAKE SYSTEM The airbrake system must be capable of opening without damage to the actuators or the doors at the rocket s maximum velocity. While the airbrakes are not designed to open at maximum velocity, they must still be able to withstand the forces in case of accidental opening. The airbrake doors must close after apogee and before landing. D. AVIONICS SYSTEM The avionics system must be removable and able to withstand the forces from launch and landing. It must also be capable of being activated externally on the launch pad. The door to the avionics bay must be able to withstand the forces from launch and landing. The system must be programmable to supply enough voltage to run the various tasks at user set altitudes or times Flight Profile Simulations

51 A flight profile simulation from Rocksim is shown in Error! Reference source not found.. The altitude with no airbrake deployment, velocity, acceleration, burnout, and apogee time are shown, max predicted values and time data are as follows, respectively: 5839 ft, 706 ft/s, 1437 ft/s 2, 2.1 s, and 18.8 s. Figure 29: Flight simulation plots Predicted altitude resuts from Rocksim with no airbrake deployment are shown below. The results are performed at an altitude of 639 ft, assumed altitude of Huntsville, AL. Simulations 1-5 increase in temperature with 15 degree increments, starting at 30 F. The mean altitude is predicted to be 5839 ft. With air brake deployment, the rocket s new altitude will be between 5050 ft and 5280 ft. ALTITUDE PREDICTIONS Figure 30: Rocksim estimated altitudes University of Minnesota USLI FRR 2012 Page 51

52 MOTOR THRUST CURVE The thurst curve for the Cesaroni L-2375-WT motor is shown below. The rocket motor burnout occurs at 2.1s and max thrust is 2603 N. While we were unable to achieve our full scale test flight by this report, we plan to verify our motor s thrust curve using data collected by our altimeters. Figure 31: Motor thrust curve from Rocksim Validity of Analysis Results from half-scale flight showed the rocket altitude was 250 feet higher than RockSim projected despite the fact Rocksim did not have notable d drag accounted for such as the excessively robust fin mounts, non-tapered fins, and uneven paint job. This may have been due to RockSim over accounting for drag caused by the transition. Therefore, on the full scale, our initial altitude estimation will be high. We can validate this on our first flight, then adjust on our second flight when airbrakes will be used Stability Margins The CP and CG are shown in Figure 32. The static margin of the full scale rocket is The CG and CP are located inches and inches from the front of the nose cone, respectively. The total rocket length is 126 in. After the Cesaroni L2375-Classic motor is depleted, the static margin remains at until apogee. University of Minnesota USLI FRR 2012 Page 52

53 Figure 32: Rocksim 2-D image Kinetic Energy Management We will be able to meet the 75 ft-lbf kinetic energy requirement with Iris Ultra 144. According to Rocksim, Fruity Chute Iris Ultra 144 will decent at around ft/s. The results of kinetic energy calculation are shown below. Weight Kinetic Energy Forward section lbs ft-lbf Mid section lbs ft-lbf Rear section lbs ft-lbf Table 15: Kinetic energy calculations For 20 ft/s of wind, we need to reef the chute to 130 inch outer diameter. In this case, the rocket will descend at ft/s according to Rocksim. For this descent rate, the values of kinetic energy will be following Weight Kinetic Energy Forward section lbs ft-lbf Mid section lbs ft-lbf Rear section lbs ft-lbf Table 16: Kinetic energy calculations with maximum windspeed Altitude and Range with Variable Wind Conditions According to Rocksim, in all the wind conditions, the vehicle will land at less than 2500 feet from launching site. Therefore, we will meet the 2500 feet of landing requirement. For 20ft/s of wind, we need to reef the chute. Drift distance 0ft/s of wind 0 University of Minnesota USLI FRR 2012 Page 53

54 5ft/s of wind 567 ft 10ft/s of wind ft 15ft/s of wind 1728ft 20 ft/s of wind* 1878ft Table 17: Drift range with various wind speed *reefing; Outer Diameter: 130 inch The setting of Rocksim Drogue chute will be deployed at apogee, and the main chute will be deployed at 700ft. The engine chosen was Cesaroni- L2375. University of Minnesota USLI FRR 2012 Page 54

55 3.4 Vehicle Verification Requirement Verification 1. The launch vehicle shall carry a science or engineering payload following one of two options: a. Option 1: The engineering or science payload may be of the team s discretion. b. Option 2: The Science Mission Directorate (SMD) at NASA HQ will provide a $3,000 sponsorship Status Verification Verified Our team has chosen option 1 and the payload design can be seen in section The launch vehicle shall deliver the science or engineering payload to, but not exceeding, an altitude of 5,280 feet above ground level (AGL). One point will be deducted for each foot achieved below the target altitude. Two points will be deducted for each foot achieved above the target altitude. Status Verification Unverified Our team has not had a full scale test flight. Currently our rocksim simulations combined with our airbrake design calculations in matlab give us a projected altitude of 5,280 ft. 3. The vehicle shall carry one Perfect Flight MAWD or ALT15 altimeter for recording of the official altitude used in the competition scoring. Teams may have additional altimeters to control vehicle electronics and payload experiments. At the flight hardware and safety check, a NASA official will mark the altimeter which will be used for the official scoring. At the launch field, a NASA official will also obtain the altitude by listening to the audible beeps reported by the altimeter. Status Verification Verified Our rocket possesses a total of five flight altimeters which are each capable of providing an official score. We will report to a NASA official before choosing the official flight altimeter. 4. The recovery system electronics shall have the following characteristics: a. The recovery system shall be designed to be armed on the pad. b. The recovery system electronics shall be completely independent of the payload electronics. c. The recovery system shall contain redundant altimeters. The term altimeters includes both simple altimeters and more sophisticated flight computers. d. Each altimeter shall be armed by a dedicated arming switch. e. Each altimeter shall have a dedicated battery. f. Each arming switch shall be accessible from the exterior of the rocket airframe.

56 g. Each arming switch shall be capable of being locked in the ON position for launch. h. Each arming switch shall be a maximum of six (6) feet above the base of the launch vehicle. Status Verification Verified a. Each altimeter is connected to a magnetic switch which we can turn on while the rocket is on the pad. b. The four recovery altimeters are completely independent from all other electronics. c. We have redundant altimeters for both drogue and main parachute deployments. d. Each altimeter has its own separate magnetic arming switch. e. Each altimeter has its own battery. f. All switches can be turned on from the exterior of the rocket airframe using a rare earth magnet. g. Our half scale flight test showed the switches stayed locked ON through all flight conditions. h. No switches are located above six feet above the base of the vehicle. 5. The recovery system electronics shall be shielded from all onboard transmitting devices, to avoid inadvertent excitation of the recovery system by the transmitting device(s). Status Verification Verified The avionics bay closest to the payload electronics and the radio tracker is lined with copper tape. 6. The launch vehicle and science or engineering payload shall remain subsonic from launch until landing. Status Verification Verified The maximum Mach number achieved by our Rocksim simulations is The launch vehicle and science or engineering payload shall be designed to be recoverable and reusable. Reusable is defined as being able to be launched again on the same day without repairs or modifications. Status Verification Verified The vehicle is designed with drogue and main parachutes to be recoverable and reusable without repairs. University of Minnesota USLI FRR 2012 Page 56

57 8. The launch vehicle shall stage the deployment of its recovery devices, where a drogue parachute is deployed at apogee and a main parachute is deployed at a much lower altitude. Tumble recovery from apogee to main parachute deployment is permissible, provided that the kinetic energy is reasonable. Status Verification Verified The vehicle is designed with drogue and main parachutes for staged deployment. 9. Removable shear pins shall be used for both the main parachute compartment and the drogue parachute compartment. Status Verification Verified Both drogue and main parachute compartments have removable shear pins. 10. The launch vehicle shall have a maximum of four (4) independent or tethered sections. a. At landing, each independent or tethered sections of the launch vehicle shall have a maximum kinetic energy of 75 ft-lbf. b. All independent or tethered sections of the launch vehicle shall be designed to recover within 2,500 feet of the launch pad, assuming a 15 mph wind. Status Verification Pending Verification We will not be able to confirm our kinetic energy upon landing until we complete our full scale flight tests. 11. The launch vehicle shall be capable of being prepared for flight at the launch site within 2 hours, from the time the waiver opens. Status Verification Pending Verification The vehicle is designed with drogue and main parachutes to be recoverable and reusable without repairs. 12. The launch vehicle shall be capable of remaining in launch-ready configuration at the pad for a minimum of 1 hour without losing the functionality of any onboard component. Status Verification Verified Each altimeter is capable of being launch ready for at least 1 hour. The airbrake also has low energy usage in launch ready position and meets the 1 hour requirement. 13. The launch vehicle shall be launched from a standard firing system (provided by the Range) using a standard 10 - second countdown University of Minnesota USLI FRR 2012 Page 57

58 Status Verification Verified We are using a standard electronic firing system. 14. The launch vehicle shall require no external circuitry or special ground support equipment to initiate the launch (other than what is provided by the Range). Status Verification Verified The vehicle is designed so that it requires no external circuitry to initiate launch. 15. Data from the science or engineering payload shall be collected, analyzed, and reported by the team following the scientific method. Status Verification Pending Verification The payload is designed so that data can be collected but we have not yet reported any scientific data. 16. An electronic tracking device shall be installed in each independent section of the launch vehicle and shall transmit the position of that independent section to a ground receiver. Audible beepers may be used in conjunction with an electronic, transmitting device, but shall not replace the transmitting tracking device. Status Verification Verified We have an RF tracker in the launch vehicle as well as the payload. 17. The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR), Tripoli Rocketry Association (TRA) and/or the Canadian Association of Rocketry (CAR). Status Verification Verified We are using a commercially available L2375 motor from Cessaroni. 18. The total impulse provided by the launch vehicle shall not exceed 5,120 Newtonseconds (L-class). This total impulse constraint is applicable to any combination of one or more motors. Status Verification Verified The L2375 has a total impulse rating of 4905 Newton-seconds. 19. All teams shall successfully launch and recover their full scale rocket prior to FRR in its final flight configuration. University of Minnesota USLI FRR 2012 Page 58

59 a. The purpose of the full scale demonstration flight is to demonstrate the launch vehicle s stability, structural integrity, recovery systems, and the team s ability to prepare the launch vehicle for flight. b. The vehicle and recovery system shall have functioned as designed. c. The payload does not have to be flown during the full-scale test flight. If the payload is not flown, mass simulators shall be used to simulate the payload mass. If the payload changes the external surfaces of the launch vehicle (such as with camera housings and/or external probes), those devices must be flown during the full scale demonstration flight. d. The full scale motor does not have to be flown during the full scale test flight. However, it is recommended that the full scale motor be used to demonstrate full flight readiness and altitude verification. e. The success of the full scale demonstration flight shall be documented on the flight certification form, by a Level 2 NAR/TRA observer. f. After successfully completing the full-scale demonstration flight, the launch vehicle or any of its components shall not be modified without the concurrence of the NASA Range Safety Officer. Status Verification Failed We did not launch our vehicle prior to this document. We have scheduled a launch before competition and plan to append the analysis before competition. 20. The following items are prohibited from use in the launch vehicle: a. Flashbulbs. The recovery system must use commercially available low-current electric matches. b. Forward canards. c. Forward firing motors. d. Rear ejection parachute designs. e. Motors which expel titanium sponges (Sparky, Skidmark, MetalStorm, etc.). f. Hybrid motors. Status Verification Verified We do not use any of the prohibited items. 21. Each team shall use a launch and safety checklist. The final checklist shall be included in the FRR report and used during the flight hardware and safety inspection and launch day. Status Verification Verified The launch and safety checklists are included within this report and will be used on launch day. University of Minnesota USLI FRR 2012 Page 59

60 22. Students on the team shall do 100% of the work on the project, including design, construction, written reports, presentations, and flight preparation with the exception of assembling the motors and handling black powder charges. Status Verification Verified All design, construction, written reports, presentations, and flight preparation has been performed by students. Our team official and our team mentor can verify this requirement. 23. The rocketry mentor supporting the team shall have been certified by NAR or TRA for the motor impulse of the launch vehicle, and the rocketeer shall have flown and successfully recovered (using electronic, staged recovery) a minimum of 15 flights in this or a higher impulse class, prior to PDR. Status Verification Verified Gary Stroick is NAR certified and has a minimum of 15 flights in L or a higher impulse class prior to PDR. 24. The maximum amount teams may spend on the rocket and payload is $5000 total. The cost is for the competition rocket as it sits on the pad, including all purchased components and materials and the fair market value of all donated components and materials. The following items may be omitted from the total cost of the vehicle: a. Shipping costs. b. Ground Support Equipment. c. Team labor. Status Verification Verified The cost of all components on the pad does not exceed $5,000. The budget is documented within this report. University of Minnesota USLI FRR 2012 Page 60

61 3.5 Safety and Environment (Vehicle) Safety and Mission Assurance Analysis The safety officers for our rocket vehicle will be Gary Stroick and Mark Abotossaway. Gary has many years of experience with high powered rocketry, and is familiar with most NAR and FAA safety codes. He has also served as RSO at the Tripoli Minnesota launches. Mark Abotossaway is the Team Lead, and has his NAR Level One certification. His responsibilities will include shop safety and hazardous material handling as well as briefing the team on all safety related issues weekly, at the team meetings. He is also responsible to maintain a safe work environment for the team, ensuring that they have all the proper equipment and training needed to safely and successfully fabricate a high powered rocket Personnel Hazards Leading up to the operational phase of the Life Cycle, addressing all personnel hazards is essential. Aside from the hazards of constructing a high powered rocket, which has been addressed in the Critical Design Report, there remain many different hazards involved with the testing and launching of high powered rockets. Testing of systems prior to test flight has inherent risks. One of the key risks involves black powder ejection tests of both our parachutes and our payload. Launch preparation has many hazards involved and team awareness of all hazards is the best prevention against incurring any incidents. Launch preparation begins in the workshop, with initial systems prep. It is followed by final systems preparation at the launch site, and this is perhaps the most dangerous situation, as there are many different systems preparations happening simultaneously, and nerves and anticipation will affect judgment. To this regard, the team has practiced full launch preparation initially with our scale vehicle. Prior to our full scale flight, we will also practice launch preparation in the workshop numerous times Environmental Concerns All team members have become familiar with the materials and compounds used in the fabrication process through the Material Safety and Data Sheets. The team members follow the proper disposal procedures of all materials. Absolutely under no circumstances are any chemicals poured down the drain. In addition to preventing damage to the rocket, a safelanding also prevents damage to the ground and restricts the environmental impact where the rocket lands. Team members with the most gas-efficient cars have driven to test

62 launches and the team currently plans to rent a passenger van to transport the team, rocket and rail. 3.6 Payload Integration Overview The rover will be contained in the payload bay located at the top of the rover, just below the nose cone. The rover will be pushed outwards using a black powder charge with an ejection piston. The black powder ejection will be ignited manually by the rover pilot once the RSO gives the all clear. This is done using an electronic switch located on the rover that is connected to the black powder canister via a molex connector. During ejection the piston will continue to push out the rover until a shock cord connecting the ejection piston to the payload bay bulkhead stop the piston s movement, allowing the rover to continue moving out of the piston. When the rover is ejected the molex connector allows the rover to separate from the ejection canister and piston. Holding in the rover during flight will be seven 4-40 shear pins. The large amount of shear pins are required to prevent the rover from falling out of the rocket during main parachute open, which has a force on it of 20 Gs. This results in the rover, nose cone, and piston have a 200 pound force on the shear pins and the shear pins are rated to hold 245 pounds of force total. An additional 40 pounds of force to push out the rover during ejection and a 10% contingency force results in a 314 pound black powder charge, or 2.5 grams given the volume of the payload bay. Ten payload ejection tests were performed on the half-scale to optimize the amount of black powder and ensure integrity of the design. For full-scale two tests shall be done, one in which the nose cone is pointed vertically upwards and the other with the nose cone pointed vertically downwards, which were found to be worse case scenarios for half-scale Compatibility The rover shall be contained by the nose cone bulkhead, Figure 33, and payload ejection piston bulkhead, Figure 33. University of Minnesota USLI FRR 2012 Page 62

63 Figure 33 Nose Cone Bulkhead Figure Figure 34 Ejection Piston Bulkhead Both bulkheads were designed to hold the rover s claw bracket in place during flight, to prevent shifting, transverse forces, and vibrations. The bracket is 3 in. in diameter and both bulkheads have a 3 in. diameter,.25 in. depth notch machined out to hold the bracket. The fits are shown in Figure 35 and Figure 35. Figure 35 Ejection Piston Bulkhead Fit Figure 36 Nose Cone Bulkhead Fit After the all clear from the RSO, the electronic switch will be remotely triggered, completely a circuit containing a 9V battery source and the black powder ejection charge. The electronic switch and the 9V battery will be housed on the rover, see Section 4 for details, and will interface with the black powder charge via a molex connector, see below figure. University of Minnesota USLI FRR 2012 Page 63

64 Figure 37 Connector used for rover ejection The molex connector will be used to allow the payload to separate from the rocket during ejection. The female end of the connector will be attached to the rover, while the male end will attach to the piston. A small hole will be drilled in the payload piston to accommodate the ignition wires. The hole will be covered with epoxy, both to protect the payload and to secure the wires attached to the male end of the molex connector in place. As the payload is ejected from the rocket, the ends of the molex connectors will separate, preventing damage to the wiring during ejection and freeing the payload of any trailing wiring Housing Integrity i. Component Integrity The payload bay is made of phenolic airframe tubing pre-glassed with fiberglass, unlike the rest of the rocket which is not glassed. This prevents the major issue that arose during ejection testing of the charge breaking the phenolic if not orientated correctly. During half-scale payload ejection tests the payload bay, also pre-glassed on the halfscale, never showed signs of damage and had a hoop stress of 215 psi acting upon it. The full-scale payload bay shall undergo a hoop stress 220 psi, therefore no damage is expected. The purpose of the G10 in the nose cone bulkhead is to prevent bending that would occur during main parachute open if it were made entirely of wood. This bulkhead is the only support for the rover during that time. The fiberglass is orientated as it is because the material is structurally stronger this way. The payload ejection piston bulkhead has an epoxy fillet on both sides to assure it does not break or come loose during any part of flight or ejection. University of Minnesota USLI FRR 2012 Page 64

65 The ejection piston is 7 in. in length, which is.5 in. short of the recommended length for it to be equal to the diameter of the tube. This was done to shave weight. Testing on the half-scale showed that even if the piston severely cocks the payload will still be able to move with more than enough energy to break the shear pins and exit the rocket. Both bulkheads on the half-scale were made of.25 in. thick wood and had a 10.5 psi acting upon them. On the full-scale both bulkheads are thicker, 1 in. on the nose cone bulkhead and ejection piston bulkhead ranges from.25 in. to.75 in., with 7 psi acting upon it, therefore neither are expected to fail. U-bolts are used for both the payload bay bulkhead and the ejection piston bulkhead and both are rated for 430 pounds, well below the 314 pounds of expected force created by the black powder charge. The shock cord is sown onto the payload bay bulkhead and the other end of the shock cord that connects to the ejection piston will be connected via a quicklink rated to 600 pounds. Although the switch is triggered electronically, there is a mechanical mechanism which closes to complete the circuit it is connected to. One concern for potential failure is the premature closing of this switch due to the forces experienced during the rocket launch. Such a closure would be devastating, causing the payload to eject during launch. To ensure this wouldn't happen, the switch was tested under the maximum forces expected during the half scale launch, just over 20g. The test was conducted by placing the switch on the half scale rover and triggering the ejection with the appropriate amount of black powder. After ejection, the switch was examined to verify the switch position had not changed due to the forces of ejection. By placing the switch in the same orientation on the full scale rocket, we can ensure the switch will now prematurely close during launch. ii. Assembly Integrity Figure 38 Integration Diagram Before assembly is to start we must make check all u-bolts are securely in place, shock cord is tied sowed or bowline knotted, bulkhead epoxy fillets show no signs or fracture or damage, and batteries are out of rover controller. University of Minnesota USLI FRR 2012 Page 65

66 Weigh black powder charge and fill canister, filling in excess canister volume with dog barf Connect payload (3) to ejection piston (2) via molex connector Connect black powder canister to ejection canister molex connection Fit payload into ejection piston bulkhead, inside ejection piston piston bulkhead Connect ejection piston to payload bay via quicklink to u-bolt Fit rover (3) into nose cone bulkhead (4) Fit ejection piston (2) and rover (3) into payload bay by pushing on nose cone until ejection piston is flush against payload bay bulkhead (1), concurrently putting nose cone into place Screw seven 4-40 shear pins into place The batteries are not placed into rover controller until rocket is on launch pad. University of Minnesota USLI FRR 2012 Page 66

67 4. Payload Criteria 4.1 Experiment Concept As stated before, the purpose of the rover is to simulate the potential value in sending inexpensive probes to landing zones on extraterrestrial bodies in order to gain more detailed scouting for future missions. The data collected by the scout would include a detailed mosaic of the landing zone terrain, and initial atmospheric readings. The data collected by the scouting rover will aid in greatly reduce the risk faced by future missions. This project is intended to be a simplified simulation of an actual mission and is intended simply to explore the potential of such a device. It is our belief that this project represents an ambitious but reasonable challenge for undergraduate students and is therefore of suitable challenge for this competition. 4.2 Science Value Payload Objectives The rover will be capable of being deployed from the airframe after touchdown without damage and in a drivable configuration. Furthermore, the rover must be capable of responding to control inputs from the ground station and returning video feed back to the pilot. Through this, the rover is also expected to be able to complete predetermined courses, including a straight line back-and-forth loop, a square pattern and a stationary rotation. During its entire time on the ground, the rover must be capable of taking and recording air temperature and humidity measurements in addition to recording solar radiation levels. After deployment, the rover must be operational for no less than five minutes. In completing these objectives the payload will simulate a small probe sent to an extraterrestrial planet to scout landing zones for future missions Payload Success Criteria 1. Payload must not sustain any damage during mission. 2. Payload must be successfully deployed on the ground after rocket has safely landed. 3. Rover must receive and execute control inputs within a range of 2,640 ft from launch site. 4. Rover must receive and execute control inputs for a minimum of five minutes after deployment.

68 5. Rover must collect temperature, relative humidity and solar radiation readings every minute for a minimum of five minutes after deployment. 6. Rover must complete the following courses: a. Drive forward 20 feet, turn 180 degrees, drive forward 20 feet and return to start. b. Drive in a square measuring 10 ft. by 10 ft. c. Survey terrain by spinning 360 degrees in one location. 7. Rover must maintain proper orientation during mission in order to receive control inputs and take valid data. 8. The collected data must be analyzed and compared to expected results to verify rover collected valid data. The payload will be considered successful if all objectives above are completed on the launch day. This means that the rover must be integrated into the airframe prior to launch, then safely and fully deployed after touchdown with all subsystems fully operational. After deployment, the rover must be undamaged and in a position that is capable of receiving and responding to all input commands. Through the commands it receives from the pilot, the vehicle must make a full rotation while remaining in one place. Next it must move forwards 20 feet, turn around and return to its starting position. Finally, the rover is expected to complete a course consisting of a 10 foot by 10 foot square, or similar course if there are major obstacles in the original course. After these requirements have been fulfilled, the rover must continue to respond to the pilot s control inputs, send video feed and take measurements from all sensors for the remaining time for a total of at least five minutes after deployment. Once the rover has been recovered, the data stored within the data logger must be readable and valid. This data will be compared to readings made by the same sensors before launch on the same day, as well as to official weather reports during operation. If the recovered data is very similar to the readings from earlier in the day and both sources are close to the official values, the mission will be considered successful. The acceptable range of values has not yet been determined and will depend on the instrumental uncertainty observed during testing Experiment Approach Because this project is intended to be an engineering challenge and not a scientific project, the data collected during the mission is not particularly valuable in an experimental sense, but will instead serve as validation that the flight as deployment left University of Minnesota USLI FRR 2012 Page 68

69 the sensors operational. To this end, the data collected during the mission must be reasonably close to the data collected by the same sensors before flight, and both data sets must be acceptably close to published air temperature and humidity data on the date of flight in the launch area. The quality of the data and of the video feed will serve as a major discussion point in determining the feasibility of using a similar device as a pre-mission scouting device Experimental Test and Measurements, Variable and Controls As stated above, the measurements taken by the rover will include the local air temperature, humidity and radiation intensity. The proposed hardware used to accomplish this goal is outlined in section 4.3, which includes the uncertainty error associated with each instrument. The variables in the experiment will be the air temperature, humidity and solar radiation level. The pre-launch readings and published air condition data will serve as the control data, but it is worth noting that the data collected during the mission is expected to conform to these results and not differ significantly. It is important to note that the readings made during the mission are not expected to be exactly the same as the prelaunch readings, as they will be taken later in the day when air conditions are not identical, nor should they be exactly the same as the published air condition data which will be taken at a different location. The launch data, however, is still expected to be reasonably close to the aforementioned control groups, which will indicate that the data collected is valid and therefore the sensors were not damaged or otherwise made inoperable during the mission. If the collected data is vastly different from the control data, the sensors will be inspected for damage, which would result in a failed mission objective Relevance of Expected Data and Accuracy/Uncertainty As stated above, the collected data will serve as confirmation that the sensors have remained operational during the mission. Therefore, the relevance of the expected data is that it will serve as indication of the completion of one of our mission objectives. More broadly, this would serve as design validation in promoting this style of probe for use in actual extraterrestrial exploration missions buy showing that the sensors and airframe are sufficient to remain operational during and after deployment. The uncertainties in the measurements made by each sensor are listed in section 4.3, and the expected bias uncertainty of the measurements will be determined during testing. The total uncertainty of all measurements is expected to be in the range of five to eight percent of any given reading. This figure will be refined and used to determine University of Minnesota USLI FRR 2012 Page 69

70 the validity of collected mission data when after testing has been completed Preliminary Experiment Process Procedures As stated in other sections, the focus of this project is to serve as an engineering challenge rather than as a scientific experiment, and as such, the experiment will be primarily focus on confirming the rover s performance. Therefore the experimental procedure process at this point starts with investigating the uncertainty of the sensors, and determining what range of values can be considered valid for a given reading. Next, the pre-launch readings will be recorded by the rover on the launch day in order to give baseline values to compare our mission data to. After that, the mission will process with launch, deployment and recovery, at which point the stored data can begin to be analyzed. The official values for temperature and relative humidity will also be considered to determine if the collected data can be considered valid and therefore if this mission objective can be considered completed. This decision will be a very important consideration in our post-launch discussion of what practical purpose this style of rover could have in larger-scale operations and what potential design improvements would have to be considered to make this possible. 4.3 Payload Design Payload Design A. Control System The control system consists of an RC transmitter and receiver, a wirelessly transmitting camera, one electronic switch, one battery pack and two high torque servos. The control system s primary functions are to give the pilot visual reference of the rover s position and to relay and execute command inputs. In essence, the system transmits a video feed to the pilot, who uses this information to make navigational decisions. The pilot inputs commands using the RC transmitter and the rover receives and executes these commands. Power Supply: Voltage Requirements: Battery Capacity: Operational Range: Torque: Transmitter Frequency: SPMB4500NM 6V 4500mAh 2500ft 21 kg-cm total 2.4 GHz Table 18: Control System Specifications University of Minnesota USLI FRR 2012 Page 70

71 The rover will be controlled from the ground station. A pilot will use a DX5e RC transmitter, along with the LRS Pro System, to send control inputs. Control inputs include, left wheel forwards and backwards, right wheel forwards and backwards, as well as toggling the electronic switch on and off. Another team member at the ground station will use a 14db patch antenna to receive transmissions from a wireless camera. The patch antenna will connect to a computer at the ground station providing a live video feed to the pilot, allowing for real time control of the rover. A conceptual drawing of the interaction between the ground station and rover is shown in the following figure. LRS Pro Receiver S6020 Servo DX5e RC Transmitter D Figure 39: Conceptual Drawing of Control System and Ground Station The SPBM4500NM battery pack will be mounted on the top of the rover and will provide power to the drive servos, the electronic switch and the RC receiver. University of Minnesota USLI FRR 2012 Page 71

72 The RC receiver acts as a relay between the RC Transmitter at the ground station and the electronic switch and servos mounted on the rover. The electronic switch provides a means for the pilot to manually control the ignition of black powder ejection charges used to deploy the payload, as well as to turn on the wireless camera. The servos will independently drive the wheels. Independent control allows the rover to turn on a dime without the addition of a complex steering system. The servos are very durable, containing metal gear boxes, allowing the system to survive the forces of take off, parachute deployment and ground ejection. The servos also deliver high torque, allowing the rover to climb over difficult terrain and to flip the rover over should it lose proper orientation, as seen in the following figure. This system allows the rover to automatically correct its orientation, provided the rover is in range of the RC transmitter. Figure 40: Automatic Orientation Correction System University of Minnesota USLI FRR 2012 Page 72

73 The wirelessly transmitting camera, the BVGM-1, will be powered by a 9V battery and will be mounted on the top shelf of the rover. This position gives the camera maximum ground clearance, increasing its transmission range, and also provides the greatest line of sight. The control components and their mounting locations can be seen in the following figure. Electronic Switch RC Receiver Camera Data Logger Servo Figure 41: Rover Drawing With Control Component Locations B. Data Collection System The data collection system consists of a data logger and two atmospheric probes, used to collect and store temperature, relative humidity and light intensity readings (specifications for electronic components given below. The system will operate from payload ejection to recovery, with a minimum of five minutes of data collection. Upon payload recovery, the collected data will be uploaded to a computer for analysis. Temperature Range: -40ºC to 75ºC Light Intensity Range: 0 to 1280 W/m 2 for wavelengths of light ranging from 300 to 1100 nm RH Range: 0-100% Memory: 512K non-volatile flash Power Supply: 4AA batteries ( housed in Data Logger) Battery Life: 1 year Sampling Rate: 1 sample per second Table 19: Data Collection system Specifications University of Minnesota USLI FRR 2012 Page 73

74 The data logger will be stored inside the chassis to offer maximum protection, while the sensors will be mounted on the top shelf of the rover to ensure maximum atmospheric exposure, as shown in the following figure. In this configuration the system is capable of recording and storing atmospheric readings once every second for up to one year. Light Sensor Temp/RH Probe Data Logger Figure 42: Rover Drawing Showing Data Collection System All components of the data collection system are designed to be used in outdoor environments and contain protective shielding. The Data Logger has a hard plastic casing that surrounds the electronics, providing structural support as well as waterproofing the system. The system has yet to be built or tested, but should be capable of meeting all data collection requirements. C. Structural/Mechanical Elements The rover chassis is made out of a rectangular frame made of G10 fiberglass with a thickness of 3/16 in. This material will be sufficiently strong to withstand all loads experienced during flight, deployment and ground operation, and will not interfere with wireless signals necessary for control inputs or output video transmission. Additionally, the rover frame is large enough to house all of the control, data collection and other electrical components described in sections A, B and D. The data logger will be placed in the lower rover shelf, and the main battery, secondary battery, wireless camera, temperature/humidity probe and light intensity sensor will all be University of Minnesota USLI FRR 2012 Page 74

75 located on the upper shelf. Furthermore, the RC receiver, RF tracker and drive servos will be located in the shoulder bulkheads on either side of the shelves. A model of the rover with its electronic component layout can be seen below: Figure 43: Solidworks model of assembled rover The figure above also demonstrates the claw mechanism that will be used to propel the rover forward. The claws will act as the drive wheels of the rover and are mounted on the servo axel, which is fixed to the outer bracket, which is attached to the arms of the claw. With this system, the torque generated by the servo turns the entire claw assembly, which effectively acts as the wheels of the rover. Also mounted on this axel is a matching bracket that holds a linkage to each arm that will force the arms outward as the bracket is pushed away from the side of the rover. The claw assembly will be mounted to the servo by bolting the axel to a mount on the servo horn so the claw can be removed if necessary. The purpose for this system is that the effective radius of the assembly can be changed during the mission by moving the inner bracket in or out. Thus, when the rover is loaded into the airframe, the bracket will be forced into the inward-most position to close the legs down to a radius of 3.7in, but when the rover is deployed, a compressed spring will automatically push the bracket out to its outward- University of Minnesota USLI FRR 2012 Page 75

76 most position to give the claw mechanism an effective radius of roughly 6 in. In this way, the rover can gain significant ground clearance without expending the payload section of the airframe. This will make the rover much more capable of handling rough terrain than the original design of simply having dive wheels. Additionally, if the legs are forced inwards the spring will compress, effectively acting as a suspension system. Note that in the above picture, the left claw is shown in the fully deployed drive configuration, while the right claw is shown in the compact flight configuration even though both claws will both be in the same mode during the mission. A photograph of the manufactured claw assembly is shown in below: Figure 44: As-built rover claw assembly. Note that the drive axle in this picture has not been trimmed and is longer than it will be in the final rover. Another important mechanical component of the rover is the outrigger arm, which will be attached to a tensioned cable for automatic deployment during rover ejection. The outrigger will provide the counter-torque needed for forward driving as well as a stop for when the rover is spinning to orient itself after deployment. Unlike previous wheeled designs, the new outrigger is simply a long peg in order to make the component as compact as possible. A figure of the deployed outrigger is shown below: University of Minnesota USLI FRR 2012 Page 76

77 Figure 45: Deployed outrigger arm D. Electrical Elements i. Wireless Camera A BoosterVision BVGM-1 Camera, as specified in the CDR, will be mounted on the top front of the rover and will transmit a live video feed to the ground station to aid in navigation. The camera is powered by a 9V battery and transmits at 2.4GHz. With the aid of a 14db patch antenna, this signal can be received at over 1 mile on the ground. Initial testing of the camera verifies its range capabilities. The camera was taken to a corn field in North Branch, Minnesota where the range was tested using the 14db patch antenna. Signals were received and processed by the receiver until about 3000ft, well over the required 2500ft. Once the rover is constructed this test will be repeated to verify that this range still holds once the camera is mounted on the rover itself. University of Minnesota USLI FRR 2012 Page 77

78 Figure 46: Wireless Camera Manufacturer: BoosterVision Model: BVGM-1 Power Requirement: 9V Battery Resolution: CMOS 380 TV lines Frequency: 2.4GHz Weight: lbs (with battery) Dimensions: 0.75x0.75x0.75 in. Cost: $73.75 Table 20: BVGM-1 Specifications ii. Electronic Switch The ejection of the payload must be triggered after the all clear is given by the RSO. To accommodate this, the rover will house an electronic switch, the RC 100X. The switch will be mounted on the top of the rover and will connect directly to the LRS Pro receiver, see Section vii. The other end of the switch will connect to a quick release mechanism, which will connect wires to the black powder ejection charges. Upon receiving the all clear from the RSO, the pilot will send a signal from the DX5e transmitter to the electronic switch, which will allow electricity to flow through the wires and trigger the ejection charges. After ignition, the wires connected to the charges will separate and the payload will be ejected. The switch will be operated at 6V, receiving power from the receiver battery pack. Upon triggering, the switch will close and complete a circuit including the wireless camera, a 9V battery and the black powder ejection charges. This set up is shown in greater detail in the schematics following component discussion. University of Minnesota USLI FRR 2012 Page 78

79 The electronic switch was tested by performing static ground deployment tests on the half scale model. The switch was successfully remotely triggered, causing the payload to be ejecting while simultaneously turning on the BoosterVision camera. Figure 47: Electronic Switch Manufacturer: RCATS Model: RC-100X Power Requirements: 6V (Supplied by Receiver Battery Pack) Cost: $29.95 Table 21: RC-100X Specifications iii. Servos Two high torque servos will be mounted on the rover to direct drive the wheels. Each servo will be operated independently, allowing the rover to turn on a dime without any additional steering system. High torque servos were chosen to give the rover enough power to drive across the rugged terrain experienced in a field. These servos were also chosen for their durability, their metal gear boxes providing more strength then the plastic gears of competing servos. University of Minnesota USLI FRR 2012 Page 79

80 Figure 48: S6020 Servo Manufacturer: Spektrum Model: S6020 Power Requirements: 6V (see G) Torque: 10.5 kg-cm Gear Type/Material: Metal Motor Type: Brushed Speed: 0.19sec/60 degrees Weight: 0.106lbs Dimensions: 1.5x1.6x0.8 in. Cost: $49.99 Table 22: S6020 Specifications iv. Battery Pack The battery pack will be mounted on the top of the rover and will act as a power supply for the RC receiver, the electronic switch and the drive servos. Given the configuration will need to be powered for at least one hour; the largest possible battery life is desired. This battery was chosen because it has about twice the milliamp hours of the average RC battery pack. For conservative estimation assume both the servos and the electronic switch will draw 200mA. This yields a total operation time of 7.5 hours, far greater than the required limit. Part of the excess charge is to ensure the battery remains at 6V throughout the mission. The torque of the servos, and thus the performance of our rover, significantly drops with a decrease in voltage. The large University of Minnesota USLI FRR 2012 Page 80

81 capacity of this battery eliminates the risk of operating at low voltages. Figure 49: SPMB4500NM Batter Pack Manufacturer: Spektrum Model: SPMB4500NM Battery Type: NiMH Number of Cells: 5 Cell Size: Sanyo 4/3A Voltage: 6.0 V Capacity: 4000 to 4999 mah Connector Type: EC3 Weight: lbs Dimensions: 0.7x3.5x2.8 in. Cost: $72.99 Figure 50: SPMB4500NM Batter Pack v. RF Transmitter One Radio Frequency Transmitter will be placed on the rover to aid in recovery. The rover will separate from the rocket, so it s important that each vehicle has its own recovery system. University of Minnesota USLI FRR 2012 Page 81

82 Figure 51: PT-1B Transmiter Manufacturer: Communications Specialists, Inc. Name: PT-1B Transmitter Band: MHz Number of Channels: 128 Output: 1mW Range: 2 miles Power Requirement: CR2032 Battery, 30 days of operation Key Features: Waterproof Shockproof Weight: lbs Dimensions: 1.1 in. diameter, 0.51 in high Cost: $49.95 Table 23 PT-1B Specifications vi. RC Transmitter Control inputs will be sent to the rover from the ground station using the DX5e RC transmitter, with the aid of the LRS Pro transmitter. The LRS attaches through the trainer port of the DX5e and overrides the normal transmission from the DX5e. By placing the DX5e in trainer mode as a student, the LRS Pro system will retain the settings of the DX5e and allow controls to be sent with the DX5e controller. The LRS Pro will emit signals at 433MHz and 0.5W, to be received by the LRS Pro receiver. This transmitter was added to boost the range of the rover to well over one mile on the ground. Additionally, because it operates at a different frequency than the camera, we will experience less interference in the camera image. Range tests are planned in the following weeks to verify the performance of the new transmitter. The LRS Pro is powered by a 9.6V battery pack. University of Minnesota USLI FRR 2012 Page 82

83 Figure 52: Picture of the LRS Pro (Silver) mounted to RC Controller vii. RC Receiver The receiver will be mounted on the upper shelf of the rover to get as much ground clearance as possible. The receiver has two flexible antennae, which will be mounted perpendicular to each other to maximize the range of the rover. Two S6020 servos, used to drive the wheels will connect directly to the receiver through the open ports shown in the following figure. Each port is programmed to a different control channel of the RC Transmitter and the AR600 will relay signals on the various control channels to both the servos and the electronic switch. University of Minnesota USLI FRR 2012 Page 83

84 Figure 53: LRS Pro RC Receiver (Green) with Antennae Connected The LRS Pro Receiver will be powered by a 5 cell NiMH battery pack, the Spektrum SPMB4500NM battery pack discussed in the CDR. The battery pack will connect directly to an open port on the AR600 receiver and supply 6V to power the receiver, the electronic switch and both servos. viii. Electronic Schematics S6020 Servo S6020 Servo LRS Pro RC - Receiver SPMB4500NM Battery Pack RC-100X Electronic Switch Ejection Charge Quick Release Figure 54: Electrical Schematic of Payload Control System University of Minnesota USLI FRR 2012 Page 84

85 The above schematic shows the electrical connections between the various components of the rovers control system. All components receive power from the SPMB4500NM battery pack and all components are directly or indirectly connected through the AR600 receiver, which acts as a relay between the components and the RC transmitter. It should be noted that the positions and sizes of the components in the schematic are not accurate. The schematic simply shows the wired connections between the components. HOBO H Data Logger S-LIB-M003 Light Sensor Temp/RH Probe Figure 55: Electrical Schematic of Data Collection System The above schematic shows the electrical connections between the atmospheric sensors and the Data Logger. The Data Logger is powered by its own collection of AA batteries, which will supply power via serial connection to both the light sensor and the temp/rh probe. The Data Logger will collect and store this data during the mission. After recovery of the rover, the data logger will be connected to a computer through a USB connection and the data will be uploaded for analysis. Again, the schematic only shows the wired connections and doesn t reflect the relative sizes or locations of the various components. University of Minnesota USLI FRR 2012 Page 85

86 Camera Electronic Switch Black Powder 9V Battery Figure 56: Electrical Schematic of Ejection/Camera Circuit The above schematic shows the circuit connecting the electronic switch, the wireless camera, the black powder ejection charge and the 9V battery. The camera and the black powder charge are wired in parallel, each receiving the required 9V to operate. Additionally, immediately after ejection, the black powder connection is broken the circuit reduces to just the closed switch, the 9V and the camera Precision of Instrumentation A. Data Logger The data logger will be housed within the payload chassis to protect the system from the surrounding environment. The logger has four serial ports which will be used to connect with the temperature/relative humidity sensor as well as the solar radiation sensor. This data logger was chosen over the Watchdog model presented in PDR because of its ability to take measurements every second, as opposed to every minute. Additionally it has a delayed start feature; allowing data collection to start after the vehicle has been launched, therefore collecting only the desired data. The data logger will collect and store data during the mission to be uploaded and analyzed after the rover has been recovered. University of Minnesota USLI FRR 2012 Page 86

87 Upon construction of the rover, the data collection system will be tested in the field. The rover will be ejected from the payload tube and then will operate under simulated competition conditions, i.e. in a corn field. We will ensure the data logger is capable of recording valid data by comparing the results to other published/expected values for the measured parameters. The test will also confirm the electronics are capable of withstanding the forces and accelerations experienced during the competition. Figure 57: HOBO Data Logger Manufacturer: Onset Model: HOBO H Power Requirements: 4 AA Batteries Battery Life: 1 year Memory: 512K non-volatile flash Logging Interval 1 second to 18 hours Weight: 0.8 lbs Size: 3.5x4.5x2.125 in. Cost: $368 Figure 58: Data Logger Specifications B. Temperature/Relative Humidity Sensor The temperature/relative humidity probe will transmit data to and receive power from the data logger via a serial connection. The probe will be positioned on the upper shelf of the rover to fully expose the sensor to the surrounding environment. University of Minnesota USLI FRR 2012 Page 87

88 Figure 59: HOBO Temp/RH Probe Manufacturer: Onset Model: S-THB-M002 Power Requirements: Powered by Data Logger Temperature Range: -40ºC to 75ºC Temperature Resolution: 0.02ºC at 25ºC Temperature Accuracy: ±0.21ºC from 0º to 50º C Relative Humidity Range: 0-100% Relative Humidity Resolution: 0.1% at 25ºC Relative Humidity Accuracy: ±2.5% from 10% to 90%, max. of ±3.5% Dimension: 0.39 in diameter, 1.39 in length Weight: 0.24 lbs Cost: $189 Table 24: Temp/RH Probe Specifications C. Solar Radiation Sensor The solar radiation sensor, a Silicon Pyranometer, will transmit data to and receive power from the data logger via a serial connection. In order to accurately measure light intensity, the sensor will be positioned on the upper shelf of the rover, fully exposing the sensor to the sun. University of Minnesota USLI FRR 2012 Page 88

89 Figure 60: Solar Radiation Sensor Manufacturer: Onset Model: S-LIB-M003 Power Requirements: Powered by Data Logger Wavelength Range: 300 to 1100 nm Measurement Range: 0 to 1280 W/m 2 Accuracy: ±10 W/m 2 or ±5% (whichever is greater) Resolution: 1.25 W/m 2 Azimuth Error: ±2% at 45º from vertical, 360º rotation Dimensions: in high, 1.25 in diameter Weight: 0.25 lbs Cost: $210 Table 25: Silicon Pyanometer Specifications D. Precision of Instrumentation The accuracy of the instrumentation is given in the components summaries above, that is the bias uncertainties in the measurements of temperature, relative humidity and solar radiation by the array of sensors connected to the HOBO Data Logger. Statistical analysis of the expected 300 to 600 readings taken after deployment will yield precision uncertainty. We will assume that during the very short time after deployment during which readings are taken there is no significant change in the levels of temperature, relative humidity and solar radiation. Thus any variance between readings is indicative of precision uncertainty. Given the bias uncertainties reported by the manufacturers, it is expected that total uncertainty will fall in the range of 5-8%. E. Repeatability of Measurement University of Minnesota USLI FRR 2012 Page 89

90 Due to the independence of our measurements and the actual rocket launch, the experiment is highly repeatable. In order for the experiment to be a success essentially 4 factors must be met. The rover may not sustain any damage during the mission, the rover must collect atmospheric data, the rover must transmit a live video feed, and the rover must receive and execute command inputs from the pilot. The largest risk to the rover, in terms of sustaining damage, is take off and ground deployment. While full scale rocket launches are costly and time consuming, simulating the ejection is fairly simple. By simply packing the rover into a cylindrical tube measuring roughly 7 to 7.5 inches in diameter, one could recreate the conditions of payload ejection. Additionally the remaining three requirements pertaining to data collection and command execution may be completed with simple ground tests and do not require a rocket launch. The simplicity of testing this experiment makes our measurements very easy to repeat and verify. This repeatability opens up the possibility for extensive ground testing prior to launch, a crucial factor to the success a complex payload such as this rover Workmanship The rover chassis will be composed of G10 fiberglass panels, and thus the workmanship required to successfully fabricate the component is no more advanced than for the airframe. All fiberglass components can be made with basic tools available in the student machine shop. Similarly, the epoxy system that will be used to bond the fiberglass parts together is the same as that used on the airframe and thus the team is familiar with its use. The electronics will all be off the shelf components that require little to no modification before use and are thus not a fabrication concern. The brackets that hold the rover claw components was considered to be too difficult to manufacture with the machines available in the student shop, however they were able to be 3D printed out of ABS plastic and reinforced with fiberglass. The rover components requiring the highest level of workmanship includes the entire claw assembly, which has already been manufactured and fit together as intended Test and Verification Program A. Range Testing As the rover must be remotely operated, it is important to ensure that both the camera and the RC transmitter/receiver operate within the full required range of 2500ft. To simulate the conditions of competition, the transmitting components will be range tested in a corn field in North Branch Minnesota. At this point the camera has been range tested, and signals have been received for distances up to 3000ft, satisfying the range requirement. The LRS Pro transmitter has University of Minnesota USLI FRR 2012 Page 90

91 been implemented, but we are waiting to range test until after the rover has been completely constructed, at which point we can more closely simulate the competition conditions. B. Electronics Testing A series of ground ejection tests on the half scale model have been conducted. In addition to testing the ejection system, the tests also confirmed the integrity of the electronics. The black powder charges were calculated to match the maximum forces expected on the full scale flight in order to verify the electronics could withstand the loads. The camera, electronic switch, battery pack, 9V battery, and RC receiver were all tested and shown to survive the associated forces. Additionally, the tests concluded that the associated g forces would not prematurely close the electronic switch, ensuring no catastrophic failures. The light intensity, temperature and relative humidity sensors, along with the HOBO data logger have all been tested for functionality. We were able to record readings every second for as long as we desired. The data for an initial test is shown below. The test was conducted in doors, with the only source of light being a lamp. The light intensity was varied by moving the sensor closer and farther away from the light source, while the temp. and relative humidity readings were varied by breathing on the prove. The test was meant to show that the sensor took accurate readings, and responded appropriately to external changes. The light intensity sensor showed a decrease in light intensity as the sensor moved away from the lamp, as expected. When breathing on the temp./rh probe, the temperature increased along with the relative humidity, as expected. The values for temp, light intensity, and relative humidity, recorded without disturbances, were of the level we were expecting. No official values of these parameters were available to compare too, however the data appears to be reasonable. University of Minnesota USLI FRR 2012 Page 91

92 Figure 61: A test of the HOBO data logging system C. Drive Testing Upon construction of the rover, and prior to competition, the rover will be tested to confirm the torque supplied by the servos meets the established requirements. The primary test will ensure the rover is capable of flipping its orientation, and the secondary test will determine what slopes the rover is capable of climbing up. Additionally the operational life of the battery will be determine to verify a long enough life for competition requirements, i.e. a minimum of 1 hour in the on position plus the approximate length of the mission. Finally, the rover must be tested on a variety of surface conditions, primarily a corn field in North Branch, Minnesota, to determine how rugged of terrain it will be capable of crossing. D. Ejection Testing As previously detailed, the ejection of the payload was successfully tested a total of 7 out of 10 times. Two of the failures were due to a broken wire that went unnoticed. The wire connected the 9V battery to the black powder, and supplied the ignition current. Once this wire was noticed, all following tests were a success. The third failure was due to a failure in the nose cone bulk head, against which payload sits. The epoxy job used to hold this bulkhead in place was done poorly and was the cause for failure. After fixing this bond, all following ejection tests were a success. University of Minnesota USLI FRR 2012 Page 92

93 The ejection was tested in 4 different configurations, nose cone pointing up at about 45º, nose cone parallel to ground, nose cone pointing down at 45º and nose cone pointing directly into ground, 90º. In all configurations the payload was ejected from the payload ejection piston, clearing both the nose cone and the payload bay, landing in a drivable configuration. The same ejection system will be implemented on the full scale rocket, and will be tested in several configurations to ensure full scale success. Additionally, because the payload half scale model was more of a mass simulator, than a structural simulator, the full scale ejection tests will verify the strength of the rover structure. 4.4 Payload Verification A. Terrain Handling The rover will be operating in difficult terrain, navigating the tilled rows and downed stalks of a corn field. This terrain presents a significant challenge in both obstacle avoidance and terrain handling. The use of spoked wheels greatly improves the terrain handling of the rover. Such a wheel allows the rover to skip over obstacles, rather than continuously drive over them. Additionally the new wheel design expands the effective wheel radius, further enhancing the terrain handling capabilities of the rover. Through the use of a BVGM-1 wirelessly transmitting camera, the pilot will have a 60 degree field of view from the point of view of the rover. Using this live video feed the pilot will be able to make judgements about the terrain and chose course that avoid particularly difficult locations. The use of the camera will aid the pilot in successfully completely the three drive courses set out in Section 4.3 as part of the success criteria. The aid of a camera will help the pilot to avoid the most difficult obstacles, but as the rover will be operating on a cornfield, there will be no clear path at any time. The rover must be rugged enough to deal with the difficulties presented by such a field, both in structural design and drive power. The use of high torque servos ensures the rover will be capable of crawling over difficult terrain. Each servo is capable of delivering 10.5 kg/cm or torque, for a maximum of 21 kg/cm. This level of torque allows the rover to climb up slopes as steep as 30º, assuming a continuous wheel. However, the addition of the spoked wheel improves this value. B. Orientation Maintenance University of Minnesota USLI FRR 2012 Page 93

94 As discussed in Section 4.3.1, the rover is capable of correcting its orientation. Under normal circumstances the rover s deployable outrigger counteracts the reactive torque generated by the drive servos. However, when the rover is in any other orientation, this reactive torque is unbalanced and causes the rover to automatically flip into the proper orientation. This system will work as long as the moments due to gravity, the only moments opposing this correction, are less than the moment due to the reactive torque. The S6020 servos provide sufficient torque to overcome the moments due to gravity and flip the rover into the correct orientation. The maximum moment due to gravity is given by, where W is the weight of the rover, d is the maximum distance to the center of gravity and M is the moment due to gravity. Solving this equation for d, and knowing the maximum weight of the rover and the maximum moment the servos can counteract, we can determine the maximum distance to the center of gravity. d M W 21kgcm * g 2.73kg * g 7.69cm 3.03in. Given that the radius of the rover is 6 in. and given that most of the weight is evenly distributed about the axis of rotation, the center of gravity should be well within 3.03 inches, guaranteeing the rover will be capable of correcting its orientation. C. Command Execution The rover s control system depends on two wireless connections, the live video feed from the wireless camera to the ground station, and the RC transmissions sent to the rover by the DX5e RC transmitter. As long as the rover is capable of receiving these transmissions the rover will be capable of executing command inputs. Testing of the wireless camera system has verified the range of the camera is greater than the required 2500ft, ensuring this connection should hold no matter where the rover is in the field. The range of the RC transmitter/receiver system has yet to be tested, but ground vehicles using the LRS Pro system have been shown to receive and execute command inputs at ranges over one mile on the ground. The constructing of the rover out of G10-fiberglass will prevent the rover from blocking command signals. Future ground testing will verify the rover is capable of executing command inputs for a multitude of ranges and orientations. D. Data Collection The data logger and sensors used for the payload are designed to be used in outdoor environments and have protective casing. Along with the protection offered by the chassis, the electronics should be shielded from any potential damage from rough terrain. The data collection system collects data once every second and has a one year University of Minnesota USLI FRR 2012 Page 94

95 battery life, far exceeding the requirements of data collection once every minute for a minimum of five minutes. Initial testing has shown the HOBO data logger and associated sensors are capable of collecting valid data for user specified intervals. E. Ejection As mentioned before, a series of payload ejection tests were conducted on the half scale rocket. The tests show that the ejection system works, and that the rover electronics are capable of surviving the associated forces. While the structure of the half scale payload did not match the full scale model, the design itself verifies the full scale rover will be capable of surviving the same forces as the half scale. The highest load that the rover will experience over the entire mission will be during ejection, at 288lbf. This necessitates that the frame be strong enough to handle this high load without failure. The narrowest section of the frame in terms of cross-sectional area is at the bracers, which are half of an inch long and 3/16in thick for each of the six braces, which totals in^2 of contact area. When a full 288lb load is applied to this area, the maximum stress totals 0.512ksi, which is well below G10 fiberglass specked bonding strength of 2.0ksi and it s expected compressive strength of 60.0ksi. Keep in mind that this is still a very conservative estimate because the bracers will be notched into place and thus will have more strength than the simple epoxy bond would provide. All other loading conditions are lower than this example and all other cross-sectional areas are greater than this example, thus the frame can be considered sufficiently strong for the entire mission. The more likely failure during ejection would be of an electrical component, but since this data is not made available by the manufacturer, this limit will have to be determined during testing. 4.5 Safety and Environment (Payload) Safety Officer As stated in the Safety and Environment Vehicle section, the team safety officers will be Gary Stroick and Mark Abotossaway Failure Modes Payload Failure Modes Risk Likelihood Consequence Mitigation Status Rover out of range of RC Controller Highly Unlikely Rover fails to receiver control inputs and is uncontrollable. Implementation of LRS Pro greatly improves the Complete University of Minnesota USLI FRR 2012 Page 95

96 Rover lands in unfavorable orientation. Rover is damaged during take off, flight, and or landing. Highly Likely Unlikely Rover has difficulty maneuvering. Rover may be incapable of driving. range, guaranteeing control authority within the required bounds. The reactive torque generated by the servos will flip the rover into the proper orientation. Configure the wheels to move inwards during flight and meet sturdy bracers connected to the frame Complete Complete Rover electronics are damaged during take off, flight or landing. The mechanical component of the electronic switch closes due to high g forces. Unable to receive camera signal. Unlikely Highly Unlikely Unlikely Rover may be incapable of receiving control inputs or taking data. Payload ejection prematurely triggered, causing a complete failure of the entire mission. Also poses significant safety threat. Impossible to navigate, driving blind. A series of 10 ejections tests on the half scale rocket have shown the electronics are capable of surviving all expected forces. Payload ejection tests simulating the maximum forces experienced by the switch have shown the switch will not close under expected forces. Ground testing has shown the camera to have an operable range of approximately Complete Complete Complete University of Minnesota USLI FRR 2012 Page 96

97 Rover is unable to navigate over difficult terrain. Rover sustains damage during black powder ejection Rover ejection takes place before bystanders are clear Unlikely Unlikely Highly Unlikely Rover cannot move, and therefore cannot meet all mission success criteria Rover incapable of meeting all success criteria Potentially poses a risk of bodily harm to bystanders. 3000ft. Use of high torque servos to guarantee rover has sufficient power. The development of expanding spoked wheels to give the rover greater ground clearance. Full scale ejection test will verify structural integrity of rover. An electronic switch has been implemented to prevent triggering of charges until all clear signal is given. Complete Proposed Complete Table 26: Summary of Payload Failure Modes The complexity of this payload presents several safety concerns. The rover is ejected from the payload via black powder charges, presenting a potential danger to anyone who may be near the rocket at the time of ejection. To mitigate this danger, the ejection of the payload will not take place until the RSO verifies that the rocket is clear of people. However, if for whatever reason the charges went off before the RSO gave the go ahead, the ejection could become a safety hazard. To deal with this, we took steps to minimize the amount of charge loaded in the payload bay, using just enough to accomplish payload ejection. Additionally the charges are now wired to an electronic switch controlled by the rover pilot, preventing ignition of the black powder before safety is verified. Another concern is premature aerial payload ejection. If the payload bay separates from the nose cone in the air, perhaps due to weak shear pins failing during parachute deployment, then the payload could fall out of the rocket. As the payload has no parachute of its own, a failure of this kind would cause the rover to free fall, thereby posing a safety concern to those on the ground. This risk has been mitigated by careful calculation of the forces experienced throughout the rocket launch, and appropriate shear pins have been selected to prevent such a failure. University of Minnesota USLI FRR 2012 Page 97

98 In addition to the safety concerns to observers, there are several failure methods that would prevent the rover from completing its mission. The failure modes fall into three general categories, structural damage, terrain obstacles, and communication failure. These failure modes and their associated risk mitigation techniques are summarized in the table above Personal Hazards and Mitigation A. Battery Handling The Nickel Metal Hydride battery must be handled with care to prevent damage to the battery and injury to those handling it. The battery will be stored in a dry location at room temperature. Additionally, the battery will be disconnected from rover for the purpose of storage. When charging the battery, one team member will always be present to ensure no circuit malfunctions and to prevent overcharging. Taking these precautions will prevent damage to the battery and will prevent sudden failure of the battery, such as an explosion, that could cause severe bodily harm. B. Chassis Construction and Assembly All team members involved in the manufacturing process of the rover will follow the same rules and safety procedures as laid out for the airframe in section. Most importantly, this included the safety review that all team members have already received in order to receive authorization to work in the Mechanical Engineering Student Shop which goes over safety procedures when operating large pieces of machinery necessary for the construction of the aluminum chassis, as well as general shop safety tips and procedures. Additionally, the MSDS forms for all potentially hazardous materials such as the epoxy we are planning to use will be available in the storage area of Akerman 103B and the team will be briefed on their content by a team safety officer. The personal protective equipment recommended in these forms are readily available in both the Mechanical Engineering shop as well as in the Akerman workspace, including safety goggles, gloves and personal respirators, as well as first aid kits and fire extinguishers. The team will recommend that team members manufacture components in pairs, to keep one another alert and focused while in the shop and alert each other of potential hazards and safety violations. C. Rover Deployment Because the deployment plan dictates that the airframe will land with active black powder charges, it is extremely important that the rover not be deployed in a way such that anyone could potentially be hit by the moving airframe components or the rover during deployment. This scenario is very unlikely since the landing zone will be kept clear of all persons, however if the airframe should somehow land near people, the rover cannot be deployed until they have moved sufficiently far away. The rover is equipped with an electronic switch that requires a control input to trigger the black powder charges, thus preventing an uncontrolled ejection. After touchdown, the team will standby and wait for the range safety officer to signal that the landing zone is clear University of Minnesota USLI FRR 2012 Page 98

99 and the rover has clearance to deploy safely Environmental Concerns As per the given requirements, all separable sections of the airframe (including the rover) will be equipped with an RF tracker to ensure their recovery. Thus no part of the rocket can be left in the field assuming no structural failures. University of Minnesota USLI FRR 2012 Page 99

100 5. Launch Operations Procedures 5.1 Checklists A. Recovery Preparation Procedural Checklist Front Avionics Assembly 1. Assemble Avionics Bay o Confirm fresh batteries (with a multi meter) o Confirm correct programming of primary and secondary altimeters (main parachute deployment) o Confirm magnetic switches functioning properly o Switch electronics off o Disconnect leads to switches 2. Assemble ejection charges o Weigh black powder for main ejection charges o Strip ejection canister leads o Check ejection canister resistance with a multi meter o Pour black powder into ejection canisters and fill remaining space with tissue paper o Run leads through aluminum holders in bulkhead o Connect charges to respective altimeters 3. Place entire avionics assembly and bulkhead in position in avionics sleeve 4. Run harness leads through access holes on outside of avionics sleeve 5. Tighten nuts to secure bulkhead in position Hardware 1. Assemble upper booster tube o Attach main piston shock cord to avionics bulkhead u-bolt with proper quicklink o Slide outer airframe tube over main piston and avionics sleeve o Pass avionics harnesses through respective holes o Secure airframe tube to avionics sleeve with rivets o Connect avionics harnesses to magnetic switches located on outside of airframe tube (ensure switches are off) o Ensure piston is flush against the avionics bulkhead 2. Attach main parachute o Properly fold main parachute ensuring shroud lines are not twisted or tangled o Attach main parachute to proper position on shock chord o Daisy chain both sides of shock cord out from main parachute o Attach proper end of shock cord to main piston u-bolt with proper quicklink o Coil shock cord into tube followed by folded main parachute o Ensure parachute is not packed in tube too tight and there is 6 inches remaining in tube to attach to coupler

101 o Attach remaining end of shock cord to u-bolt on coupler using proper quicklink o Slide coupler tube into upper booster section o Insert shear pins Rear Avionics Assembly 1. Assemble Avionics Bay o Confirm fresh batteries (with a multi meter) o Confirm correct programming of primary and secondary altimeters (drogue parachute deployment) o Confirm magnetic switches functioning properly o Switch electronics off o Disconnect leads to switches 2. Assemble ejection charges o Weigh black powder for drogue ejection charges o Strip ejection canister leads o Check ejection canister resistance with a multi meter o Pour black powder into ejection canisters and fill remaining space with tissue paper o Run leads through aluminum holders in bulkhead o Connect charges to respective altimeters 3. Reconnect leads to magnetic switches (ensuring switches are off) 4. Place entire avionics assembly and bulkhead in position in avionics sleeve 5. Tighten nuts to secure bulkhead in position Hardware 1. Prepare drogue parachute o Properly fold drogue parachute ensuring shroud lines are not tangled o Attach parachute swivel to parachute protector with proper quicklink o Wrap parachute protector around parachute o Attach opposite side of parachute protector to shock cord at respective position o Daisy chain shock cord from parachute to ends o Attach shock cord to inside coupler u-bolt with proper quicklink o Coil shock cord into coupler tube followed by wrapped parachute (ensure parachute pack is not too tight with parachute protector facing bulkhead) o Attach remaining end of shock cord to u-bolt on rear avionics bulkhead o Slide coupler (with parachute inside) in to lower booster section o Insert shear pins B. Motor Preparation Procedural Checklist (Note: Motor Preparation will most likely be handled by the NAR/TRA mentor or the Team Lead) Following the complete preparation of the vehicle and payload, the motor reloads will then be inserted in to the casing, as the final assembly step University of Minnesota USLI FRR 2012 Page 101

102 For reference, see the Cessaroni Motor Manual in the Appendix C. Igniter Installation Pre-launch: Check continuity of igniter at the bench with a multi meter Insert igniter. Be sure it is positioned correctly. Secure igniter in position. Assure that launcher is not hot. Visually confirm that key IS NOT in remote device and that arming switch is off. Be sure all connectors are clean. Attach leads to ignition device. Be sure leads don't touch each other and that circuit is not grounded by contact with metal parts D. Setup On Launcher The setup on launcher will be performed by three team members with either the NAR/TRA mentor or the team lead supervising 1. Ensure launch rail assembly is fully locked and stable 2. Tilt launch rail over to horizontal 3. Slide launch vehicle onto rail ensuring the rail buttons slide smoothly in to rail. (Support vehicle so as not to put undue force on rail buttons) 4. With both buttons in rail, and vehicle lowered to bottom of rail, slowly raise rail until vertical, supporting both the vehicle and the rail 5. Lock the rail into the vertical position 6. Ensure blast deflection plate is in place and secure 7. Turn on all electronics sequentially using magnetic switches, ensuring electronics are on by audible beeps a. Begin with primary and secondary main altimeters b. Follow with primary and secondary drogue altimeters c. Turn on airbrake system altimeter 8. Visually inspect vehicle to ensure there is no damage during rail loading, and all shear pins/rivets are in place 9. Proceed to Igniter Installation E. Launch Procedures The Launch Procedure will consist of checklists for all systems preparations, beginning with vehicle assembly at the assembly station. The checklists will be maintained by the NAR/TRA mentor and the Team Lead, while the remaining team members will implement assembly procedures sequentially. 1. Prepare the recovery system (see Recovery System checklist) 2. Prepare payload system (see Payload System checklist) 3. Assemble the motor (see Motor Preparation checklist) University of Minnesota USLI FRR 2012 Page 102

103 4. Setup launcher on pad (see Setup on Launcher checklist) 5. Install igniter (see Igniter Installation checklist) 6. Return to launch control and prepare ground station for observation of flight and control of payload 7. Verify that the range is clear and the sky is clear. 8. All team members will proceed to pre-designated positions and prepare for recovery tasks a. One team member will man payload controller and monitor video feeds b. One team member will man video patch antenna c. One team member will man RF receiver antenna d. Remaining 5 team members will make up recovery team i. Two members will remain stationary at different positions as recovery spotters ii. Three members will proceed to recover vehicle and payload when instructed and safe to do so 9. Announce the launch to observers and give a ten second countdown before motor ignition 10. Ensure launch pads are clear and RSO opens range before proceeding to recover vehicle 11. After vehicle is safely recovered, proceed to postflight inspection (see Postflight Inspection checklist) F. Troubleshooting If vehicle sections will not slide together, check that there is nothing on the coupler or inside of the airframe. If they are clean, additional sanding may be required. If there are continuity problems with the ejection canisters, check the canister connections and the altimeter connections. If those areas are okay, disconnect the ejection canister and check its continuity. If the ejection canister is faulty, replace with a new canister. If there is still a problem, examine the wiring for any possible breaks or loose ends. If the rocket does not fire when the launch button is pressed, wait at least one minute before approaching the rocket. Remove the igniter from the motor and launch leads and inspect the igniter head. If the head is burnt, replace the igniter with a fresh igniter and make sure that the head is at the forward end of the motor. If the head still looks intact, check the continuity of the igniter. If there is no continuity, replace the igniter. Check the lead clips to ensure a good connection for current flow. G. Postflight Inspection Checklist University of Minnesota USLI FRR 2012 Page 103

104 Post flight inspection will take place at the field preparation area. All team members should be present to inspect all systems for damage. 1. Upon first locating the vehicle and payload in the field, ensure that all ejection charges have fired a. If not, switch off the electronics and proceed to remove ejection charge from vehicle by cutting leads b. Twist leads together and transport unused ejection charges back to assembly area 2. Record the altitudes from the altimeters prior to shutting off electronics 3. Ensure that the vehicle is intact a. If components are damaged or missing, search the immediate area to locate 4. Recover and power off payload 5. Transport all sections and components back to the assembly area 6. Disconnect the shock cords and parachutes from the vehicle bulkheads and inspect for any singe marks or holes 7. Inspect the fin system for any damage or cracked fillets 8. After allowing time to cool, remove the motor casing from the motor mount tube and inspect for damage a. Clean motor casing immediately if needed 9. Remove the motor mount/airbrake system from the lower booster section and inspect for damage 10. Remove the rear avionics bulkhead (ensuring the charge canisters are empty) 11. Inspect the avionics bay for damage a. Proceed to download the altimeter data to computer 12. Remove the upper booster tube and transition, inspecting for damage 13. Remove the main piston shock cord from the front avionics bay u-bolt 14. Remove the front avionics bay bulkhead (ensuring charge canisters are empty) 15. Inspect the avionics bay for damage a. Proceed to download the altimeter data to computer 16. Clean black powder charge residue from booster tube interiors 17. Clean black powder charge residue from avionics bulkheads 18. Inspect payload for damage 19. Inspect payload bay and payload ejection piston for damage 20. Inspect nosecone shoulder for damage University of Minnesota USLI FRR 2012 Page 104

105 6. Activity Plan 6.1 Budget The project costs have not changed considerably since the Critical Design Review. For detailed project costs see Appendix. Our project costs are summarized in the following tables: Funding Source Amount Department of Aerospace Engineering University of Minnesota $2500 Senior Design Class Funds $1500 Minnesota Space Grant Consortium $6000 TOTAL ALLOCATED FUNDS $10000 Table 27: Funding Sources and Amounts. Expenditures Project System Amount Half Scale Subtotal $910 Full Scale Subtotal $3070 Payload Subtotal $1492 Testing and Supplies Subtotal $1992 Travel Subtotal $2386 TOTAL EXPENDITURES $9850 Table 28: Total Expenditures The budget has not changed substantially since the Critical Design Report. For the Flight Readiness Review, we calculated our on the pad costs. The following table summarizes our on the pad costs, with separate tables for the vehicle and the payload. University of Minnesota USLI FRR 2012 Page 105

106 The table is listed in order of the components as they appear on the vehicle. Component Unit Cost Qty Total Cost ProRated Status Nosecone $ $ $ Purchased Bulkhead $ $1.00 $0.50 Custom RF Tracker $ $49.95 $49.95 Purchased Airframe Tube $ $ $ Purchased Payload Piston Tube $ $21.73 $12.68 Purchased Payload Piston Bulkhead $ $0.50 $0.50 Custom U-Bolt (small) $ $7.54 $7.54 Purchased Quicklink (small) $ $1.99 $1.99 Purchsed Separator Bulkhead $ $1.00 $1.00 Custom Rivet Nuts $ $7.84 $7.84 Purchased Threaded Rods $ $4.58 $4.58 Purchased Avionics Sleeve $ $13.36 $5.01 Purchased Avionics Sled $ $2.00 $2.00 Custom Primary Altimeter $ $ $ Purchased Secondary Altimeter $ $ $ Purchased Batteries $ $10.00 $10.00 Purchased Nylon Battery Holders $ $8.00 $8.00 Purchased Aluminum E Canister $ $1.00 $1.00 Custom Outer Av Bulkhead $ $8.55 $8.55 Purchased U-bolt (large) $ $1.26 $1.26 Purchased Quicklink (large) $ $3.47 $3.47 Purchased Airframe Tube $ $39.50 $24.69 Purchased Body Transition $ $ $86.96 Purchased Magnetic Switches $ $50.00 $50.00 Purchased Main Piston Tube $ $13.36 $7.79 Purchased Main Piston Bulkhead $ $8.55 $8.55 Purchased U-bolts $ $2.51 $2.51 Purchased Quicklinks $ $10.41 $10.41 Purchased Main Parachute $ $ $ Purchased Main Shockcord $ $14.80 $14.80 Purchased University of Minnesota USLI FRR 2012 Page 106

107 Component Unit Cost Qty Total Cost ProRated Status Outer Band $ $39.50 $0.82 Purchased Coupler Tube $ $13.36 $13.36 Purchased U-bolts $ $2.51 $2.51 Purchased Quicklinks $ $10.41 $10.41 Purchased Bulkhead $ $0.40 $0.40 Custom Drogue Parachute $ $80.00 $80.00 Purchased Drogue Shockcord $ $16.40 $16.40 Purchased System: Lower Booster Airframe Tube $ $39.50 $16.46 Purchased Front Bulkhead $ $0.40 $0.40 Custom Aluminum E Canisters $ $1.00 $1.00 Custom U-bolt $ $1.26 $1.26 Purchased Quicklinks $ $6.94 $6.94 Purchased Avionics Inner Tube $ $13.36 $4.45 Purchased Primary Altimeter $ $ $ Purchased Secondary Altimeter $ $85.00 $85.00 Purchased Batteries $ $10.00 $10.00 Purchased Nylon Battery Holders $ $8.00 $8.00 Purchased Airbrake Alt $ $ $ Purchased Battery $ $5.00 $5.00 Purchased Threaded Rods $ $2.29 $2.29 Purchased Rivet Nuts $ $1.96 $1.96 Purchased Switches $ $75.00 $75.00 Purchased Av Outer Tube $ $13.36 $5.01 Purchased Fins Support Tube $ $13.36 $8.91 Purchased Surface Fin Mounts $ $24.00 $24.00 Custom Fins $ $21.00 $21.00 Custom Body Centering Ring 1 $ $0.40 $0.40 Custom MMT Centering Ring 1 $ $0.20 $0.20 Custom Motor Mount Tube $ $16.50 $14.21 Purchased Body Centering Ring 2 $ $0.75 $0.75 Custom MMT Centering Ring 2 $ $0.75 $0.75 Custom Motor Casing $ $ $ Purchased Motor Reload 4-G $ $ $ Purchased Motor Retainer $ $48.00 $48.00 Purchased University of Minnesota USLI FRR 2012 Page 107

108 Component Unit Cost Qty Total Cost ProRated Status Airbrake Hinge Assembly $ $12.00 $12.00 Custom Airbrake Doors $ $6.00 $6.00 Custom MMT Centering Ring 3 $ $0.75 $0.75 Custom Airbrake Actuators $ $ $ Purchased Airbrake Relays $ $5.12 $5.12 Purchased Trailing Airframe $ $39.50 $1.65 Purchased Boat Tail $ $ $ Purchased RF Tracker $ $49.95 $49.95 Purchased Boat Tail Centering Ring $ $0.40 $0.40 Custom Full Scale Vehicle Total $2, Table 29: On the pad vehicle costs. The majority of the components were purchased new. The table accounts for prorating parts that were modified from the purchased parts. It also prorates completely fabricated parts from the cost of the raw materials. 6.2 Timeline Parachute Ejection Test Friday, March 30 th, 2012 Payload Ejection Test Saturday, March 31 st, 2012 Full Scale Flight Sunday, April 1 st, 2012 Analysis of Full Scale Flight Data Tuesday, April 3 rd, 2012 FRR Addendum Written and Given to USLI Representatives Sunday, April 8 th, 2012 Leave for Huntsville, AL Wednesday, April 18 th, 2012 University of Minnesota USLI FRR 2012 Page 108

109 6.3 Educational Engagement Community Outreach is not only important to our success in this competition, but for the future of the aerospace industry. The goals of all of our community events were to inspire those younger than us to not only become interested in aerospace, but math, science and engineering as a whole. At this time, we have performed 5 events. The first event was an Urban 4H kickoff on October 29 th, This event was smaller with 15 children and 6 parents present. In the first half of this event we had the children make CD hovercraft and straw rockets to emulate fluid power and fin design of rockets. In the second half of the event we went outside to shoot off 3 in. water propelled rockets as well as larger foot-long rubber rocket. With these rockets we were able to show how the difference in the propellant (water or air) affects how much power is generated, as well as how compressed air can be used. This allowed the kids to apply the lessons learned from the indoor activities on rockets. The second event was the Math & Science Family Fun Fair on November 19 th, This event was six hours and had a total of 2,200 attendees including the children and adults. The event is for children from K-12, with majority of them being in 4 th -8 th. We had our own room and used the CD hovercraft and straw rockets again, along with the air pneumatic circuit kit, and used a LCD screen attached to a laptop to show videos of previous launches. A rocket used last year by a student organization was on for display and our Team Lead answered questions about it. An estimated 250 kids, with an equal amount of parents visited our room. The number is an estimate because, although we kept a tally as people walked through the door, the event was fairly busy the entire six hours. University of Minnesota USLI FRR 2012 Page 109

110 The third and fifth events were science nights at Galtier Elementary s, a K-6 school which had 47 attendees and Lincoln Center Elementary s, another K-6 which had 200 attendees. For both had a booth or table and had our half-scale rocket on for display. There was also a pasteboard demonstrating how rocket s fins and propulsion works. We had kids make straw rockets, as they are perfect for these types of events because they are easy to make and the kids can take them home with them after they make it. With these events we are projected to meet with 883 people total, with approximately 550 of them being middle school students. Error! Reference source not found. is a schedule of our past and future events. University of Minnesota USLI FRR 2012 Page 110

111 Event Date Attendees Status Urban 4-H Kickoff Octover 17th, Finished Family Fun Fair November 19th, [2200 total] Finished Galtier Elementary Science Fair February 16th, Finished AIAA Event February 23rd, Finished Lincoln Center Elementary March 22nd, Finished Science Night Marcy Open February 17th, 2012 (40 estimated) Itinerary Marcy Open February 24th, 2012 (40) Itinerary Table 30 Educational Outreach Events University of Minnesota USLI FRR 2012 Page 111

112 7. Conclusion In summary, Team Artemis is prepared to fly our vehicle and payload at the USLI competition in Hunstville, Alabama. We feel confident in our design.

113 APPENDICES

114 APPENDIX I Flight Readiness Review Flysheet University of Minnesota USLI FRR 2012 Page 114