USLI Critical Design Report

Transcription

1 UNIVERSITY OF MINNESOTA TWIN CITIES USLI Critical Design Report University Of Minnesota Team Artemis 1/23/2012 Critical Design Report 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 CDR 2012 Page 2

3 Table of Contents 1. CRITICAL DESIGN SUMMARY TEAM SUMMARY LAUNCH VEHICLE SUMMARY PAYLOAD SUMMARY CHANGES MADE SINCE PRELIMINARY DESIGN VEHICLE CRITERIA PAYLOAD CRITERIA Changes to Payload Electronics Changes to Payload Structure ACTIVITY PLAN VEHICLE CRITERIA DESIGN AND VERIFICATION OF LAUNCH VEHICLE Mission Statement Major Milestones Design Review Functional Requirements Workmanship Testing Manufacturing and Assembly Integrity of Design Mass Statement SUBSCALE FLIGHT RESULTS RECOVERY SUBSYSTEM Parachute Selections Parachute Ejection Parachute Components Recovery Hardware University of Minnesota USLI CDR 2012 Page 3

4 3.4 MISSION PERFORMANCE PREDICTIONS PAYLOAD INTEGRATION Payload Integration Assembly Shear Pins and Black Powder Required LAUNCH CONCERNS AND OPERATIONS PROCEDURES SAFETY AND ENVIRONMENT (VEHICLE) Safety Officer Failure Modes and Mitigation PERSONNEL HAZARDS AND ENVIRONMENTAL CONCERNS PAYLOAD CRITERIA TESTING AND DESIGN OF PAYLOAD EXPERIMENT Payload Systems Design Verification Workmanship Testing Status Payload Integration Instrumentation and Measurement Payload Electronics Safety Analysis PAYLOAD CONCEPT FEATURES AND DEFINITION SCIENCE VALUE SAFETY AND ENVIRONMENT (PAYLOAD) ACTIVITY PLAN BUDGET TIMELINE EDUCATIONAL ENGAGEMENT CONCLUSION APPENDIX I ACTIVITY PLAN BUDGET SHEETS University of Minnesota USLI CDR 2012 Page 4

5 Appedix Component Checklist Components Recovery Components Components Pre-Flight Procedural Checklist Electronics Preparation Forward Avionics Bay Aft Avionics Bay Install Avionics Tracking Electronics Main Ejection System Drogue Ejection System Final Launch Preparations This checklist is to be used for final launch preparations at the launch pad. At least one of the team safety officers is to be present for final launch preparations Load Rocket on Pad Prepare Igniter Final Launch Sequence Misfire Procedures Post-Flight Checklist Normal Post Flight Recovery Flight Failure Checklist APPENDIX 3 ACTIVITY PLAN EDUCATIONAL OUTREACH University of Minnesota USLI CDR 2012 Page 5

6 List of Tables: Table 1: Major Milestone Summary Table 2: Table of Functional Requirements Table 3: Mass Balance Statement Table 4: Iris Ultra Chute Data Table 5: Maximum Opening Force for Various Chutes Table 6: Electrical Components for the Avionics Bay Table 7: Featherweight Raven2 Specifications Table 8: Featherweight Raven2 Specifications Table 9: Enticore Electronics AIM USB Altimeter Specifications Table 10: Rocksim Simulation Results Table 11: Failure Modes and Risk Mitigation for Recovery System Table 12: Failure Modes and Risk Mitigation for Avionics System Table 13: Failure Modes and Risk Mitigation for Payload Integration Table 14: Failure Modes and Risk Mitigation for Airbrake System Table 15: Failure Modes and Risk Mitigation for Motor System Table 16: Failure Modes and Risk Mitigation for Fin System Table 17: Control System Specifications Table 18: Data Collection system Specifications Table 19: Temp/RH Probe Specifications Table 20: Silicon Pyanometer Specifications Table 21: BVGM-1 Specifications Table 22: RC-100X Specifications Table 23: S6020 Specifications Table 24 PT-1B Specifications Table 25: Spektrum SPMB4500NM Specifications University of Minnesota USLI CDR 2012 Page 6

7 Table 26: Summary of Payload Failure Modes Table 27: Funding Sources and Ammounts Table 28: Total Expenditures Table 29: Half Scale Expenditures Table 30: Full Scale Expenditures Table 31: Payload Expenditures Table 32: Testing and Supplies Cost Table 33: Travel Expenditures Table 34: Reports and Design Timeline Summary Table 35: Table Launch Schedule Table 36: Manufacturing and Assembly Timelines Table 37: Half Scale Timeline Summary Table 38: Educational Engagement Timeline Summary Table 39: Educational Outreach Overview List of Figures: Figure 1: Complete Vehicle Assembly Figure 2: Nosecone and Payload Assembly Figure 3: Front Avionics Bay Assembly Figure 4: Main Parachute Tube and Transition Assembly Figure 5: Coupler Between Main Parachute Main Parachute and Fin Section Figure 6: Rear Avionics Bay University of Minnesota USLI CDR 2012 Page 7

8 Figure 7: Motor Mount System Assembly Figure 8: Motor Mount Cutaway Figure 9: Fin Dimensions Figure 10: Fin Mount Dimensions Figure 11: Schematic of Airbrake Assembly Figure 12: Firgelli L16 Linear Actuator Figure 13: Airbrake Wiring Schematic Figure 14: Load and Current Curves for the L16 Actuator (Provided by Firgelli) Figure 15: Latching Relay Specifications Figure 16: Distance profile found for the rocket without airbrakes (blue), with variable airbrake opening velocity (green), and instantaneous airbrake opening velocity (red).. 38 Figure 17: This shows the force on each airbrake door for instantaneous brake deployment (red), and variable speed brake deployment (green) Figure 18: This shows the rocket deceleration without brakes (blue), with variable opening speed brakes (green), and instantaneously opening brakes (red) Figure 19: PT-1B Transmitter Figure 20: Featherweight Raven2 Altimiter Figure 21: PerfectFlite Stratologger Altimeter Figure 22: Entecore Electronics AIM USB Altimeter Figure 23 Front avionics bay wiring schematic Figure 24: Aft avionics bay wiring schematic for recovery Figure 25: Aft avionics bay wiring schematic for airbrake system Figure 26: Magnetic switch from Aerocon Systems Figure 27: Flight Profile Rocksim Simulation Figure 28: Rocksim Altitude Predictions Figure 29: Thurst Curve for the Cesaroni L1115-Classic motor Figure 30: Drag Assessment Rocksim Results Figure 31: Location of CG (Blue) and CP (Red) University of Minnesota USLI CDR 2012 Page 8

9 Figure 32: Payload Integration Assembly Figure 33: Ejection Piston Assembly Figure 34: Black Powder Canister Figure 35: Keyhole Shear Pin Figure 36: Three-View rover schematic featuring all major dimensions Figure 37: Conceptual Drawing of Control System and Ground Station Figure 38: Automatic Orientation Correction System Figure 39: Rover Drawing With Control Component Locations Figure 40: Rover Drawing Showing Data Collection System Figure 41: Labeled CAD Model of Rover Assembly Figure 42: Molex Connector Figure 43: HOBO Data Logger Figure 44: Data Logger Specifications Figure 45: HOBO Temp/RH Probe Figure 46: Solar Radiation Sensor Figure 47: Wireless Camera Figure 48: Electronic Switch Figure 49: S6020 Servo Figure 50: SPMB4500NM Batter Pack Figure 51: SPMB4500NM Batter Pack Figure 52: PT-1B Transmiter Figure 53: AR600 RC Receiver Figure 54: Electrical Schematic of Payload Control System Figure 55: Electrical Schematic of Data Collection System University of Minnesota USLI CDR 2012 Page 9

10 1. Critical Design 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 Projected Vehicle Dimensions Length: Diameter: Empty Vehicle Weight: Loaded Weight: Motor: Recovery: Rail Size: 114 inches inches payload section inches booster section 24.6 pounds pounds (with motor and payload) CTI Pro75 L1115 Dual Deployment, redundant black powder ejection One inch button See Appendix for complete CDR 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. University of Minnesota USLI CDR 2012 Page 10

11 2. Changes Made Since Preliminary Design 2.1 Vehicle Criteria A. Airframe Changes The full scale vehicle airframe material has been changed from G10 fiberglass to preglassed and phenolic tubing. The reason for the change was to save overall weight on the fully loaded vehicle and enable the vehicle to attain a higher apogee altitude. The nosecone was also changed from 3-to-1 ogive to a 4-to-1 ogive to fit the new airframe. A fiberglass boat tail was also added to reduce drag and make the airflow over the tail of the vehicle more streamline. B. Motor Changes The motor was also changed from the preliminary design to coincide with the lighter payload and vehicle. Previously we had selected a Cessaroni Technologies L1115 but have decided to use a Cessaroni Technologies L1115 instead. The old motor selection was found to overpower our vehicle substantially, and L1115 is better suited to our new vehicle weight and performance goals. Coincidentally, the two motors have the same dimensions and only differ in burn time and total impulse. 2.2 Payload Criteria Changes to Payload Electronics A. Data Collection System There were several drawbacks of the data logger chosen for the PDR, chiefly the logging interval of one minute. The rover may collect data for a period as short as five minutes, thereby limiting us to only five data points. Additionally, the light intensity sensor gives an average reading over the entire logging interval, in this case one minute. As the rover will be changing orientation, possibly disturbing the azimuth angle as much as 90 degrees, the average will not accurately reflect the true value of light intensity. For these reasons a data logger with a one second logging interval was chosen, the HOBO H Micro Station, discussed in greater detail in Sections and To accommodate this change, compatible temperature, relative humidity and light intensity sensors were chosen to meet atmospheric collection requirements. B. Payload Ejection System At PDR we were informed that the ejection of the payload on the ground must be triggered only after safety confirmation by the RSO. To accommodate this change, the black powder charges will be triggered via an electronic switch. The switch will plug into the AR600 receiver used to control the rover and will use an independent control channel on the DX5e RC controller. Upon receiving the all clear from the RSO, the pilot will trigger the switch, igniting the black powder. A quick release mechanism will be used to separate the ignition wires from the rover after the electronic signal is sent, University of Minnesota USLI CDR 2012 Page 11

12 thereby preventing damage to the rover and its electrical components. This system is discussed in greater detail in Section Changes to Payload Structure Several minor improvements have been implemented into the payload structure since the Preliminary Design Report. Perhaps most noticeably, the lower bay has been expanded considerably to house the new, larger data logger, which necessitated that the upper shelf be made much shallower to remain within the wheel radius. Additionally, the layout of the electronics had to be revised with this change, putting the main battery, camera, receiver, temperature sensor, IR sensor and 9V battery on the upper shelf, as well as moving the tracker down to the lower shelf with the data logger. Another important change was the removal of the drive suspension, which was determined to be overly complicated and not entirely necessary since the non-axial loads on the rover are expected to be small. The servos will instead be supported by two bulkheads on either side of the rover which will have slots cut to house the drive assembly. Furthermore, the axial suspension concept which separated each shelf outward with RC shocks has been removed and the shelves will instead be made of a continuous plate. To compensate for this, the wheels will move inwards in launch configuration and mate with bracers that are attached to the frame, thus transferring all axial loads onto the chassis. As the rover is ejected, the wheels are spring-loaded to move outwards into drive configuration, allowing for free rotation. These changes are detailed further in Section Activity Plan A. Budget Changes The project budget has changed since the PDR. Initial estimates of costs from the preliminary design report, especially travel costs, were realized to be low. Our expenditures have gone up but our allocated funding has gone up as well. B. Timeline Changes The project timeline has changed since the PDR. We feel we are currently behind schedule but since we are a rookie team, we have no way to gauge if our progress is on schedule. C. Educational Engagement Changes Since PDR, we have added two new events, the Galtier Middle School Science Night and a presentation at an AIAA event. As for our other events, only scheduling has changed, which is discussed in section 5.3. University of Minnesota USLI CDR 2012 Page 12

13 3. Vehicle Criteria 3.1 Design and Verification of Launch Vehicle Mission Statement It is the mission of the University of Minnesota USLI team to design, fabricate and fly a high powered rocket with a scientific and engineering payload successfully and safely and to satisfy all requirements put forth by the NASA University Student Launch Initiative Student Launch Projects. Our primary focus will be to satisfy all of the requirements with a minimal amount of cost and materials, optimizing each system and component on our rocket to its maximum capabilities without sacrificing performance. With regards to the NASA USLI competition, it is the University of Minnesota Team Artemis mission to design, test and fabricate a high powered rocket to carry a two-wheeled cylindrical rover to an apogee altitude of 5280 feet. It is also our mission to return the rover and vehicle safely to the ground within 2500 feet of the launch point, and then deploy the rover from the vehicle. We intend to meet and exceed the 24 technical requirements put forth by the NASA USLI guide. To achieve our mission successfully, we are designing all systems in tandem with each other to guarantee the maximal amount of overlap and project efficiency. With our payload estimated weight between 6 to 7 pounds, our rocket vehicle must be designed to encompass the payload and as such will be a rather large and heavy vehicle. We plan on using the largest motor possible under USLI requirements to exceed the competition apogee altitude. We have designed an airbrake system which will be used in conjunction with ballast (following test flights) to apogee our vehicle and payload at competition altitude of 5280 feet above ground level (AGL). With regards to recovering our rover and vehicle safely and effectively, we will use specific parachute calculations, to minimize drift, which will be essential for correct operation of our payload upon landing. Launch vehicle trajectory calculations will also ensure the straightest trajectory possible to apogee, with as little down range motion as possible. Mission success criteria for our team are various. With regards to the overall project, completion of all reports on time will be a success. A successful construction of a half scale and full scale vehicle will be deemed a success. Our mission will be successful upon completion of our launch vehicle and payload and a safe and successful launch and recovery at USLI with all systems working as planned. Our mission success during competition flight will depend mainly how far away we land from the launch pad, as well as what orientation we land in. University of Minnesota USLI CDR 2012 Page 13

14 3.1.2 Major Milestones A summary of our major milestones can be seen in the following table. It is not an exhaustive list however, and full timelines for design, testing, manufacturing, reports and outreach can be found in Section 5.2. NASA USLI important dates are highlighted in red. Milestone Date Dates Team Formation September 7th, 2011 Initial Concepts Developed September 7th 21st, 2011 Proposal Submitted September 28th, 2011 Website Established November 4th, 2011 Preliminary Design Phase September 28th November 14th, 2011 Preliminary Design Report Submitted November 28th, 2011 Half Scale Design Phase November 28th December 3rd, 2011 Half Scale Systems Tests December 5th 22nd, 2011 Half Scale Full Assembly December 26th January 17th, 2012 Half Scale Flight Test January 18th, 2012 Critical Design Phase November 28th January 14th, 2012 Critical Design Report Submitted January 23rd, 2012 Rebuilding of Scale Vehicle January 30 th February 3 rd, 2012 Tooling and Manufacturing Prep February 6 th 10 th, 2012 Systems Assembly and Testing February 6 th 17 th, 2012 Ground Testing and Final Assembly February 20 th 24 th, 2012 Final Paint and Finishing February 27 th March 2 nd, 2012 Full Scale Flight Test March 3 rd, 2012 Last Day for Full Scale Flight March 10th, 2012 Flight Readiness Review Submitted March 26th, 2012 Travel to Huntsville, Alabama April 18th, 2012 Launch Day April 21st, 2012 PLAR Due May 7th, 2012 Table 1: Major Milestone Summary University of Minnesota USLI CDR 2012 Page 14

15 3.1.3 Design Review The complete launch vehicle consists of many individual systems and sections which must work in unison for a successful flight. The design has been challenging to optimize all systems while still meeting all launch and competition criteria. Figure 1: Complete Vehicle Assembly A general listing and discussion of all vehicle components, assemblies and systems is given here. Further detail will be provided in Section A. Nosecone Section The nosecone we have chosen is a 4:1 ogive cone constructed of fibreglass. It is 29 inches in length and will contain a radio transmitter which will sit on a wood bulkhead located in the shoulder of the nosecone. This nosecone was chosen due to its decrease in vehicle drag as well as cost and ease of attainment. A more detailed drawing of the nosecone and bulkhead is given in Section B. Payload Section The payload section is comprised of an outer pre-glassed phenolic tube of length inches. Within this outer airframe tube lays our payload (14 inches), piston (7.5 inches), and front avionics bay. The nosecone is held on with shear pins which are broken when the payload is ejected using a black powder charge. A more detailed description of this system is given in the payload integration section 3.5. C. Forward Avionics Bay University of Minnesota USLI CDR 2012 Page 15

16 The front avionics bay is located on the aft side of the bulkhead epoxied to the payload section outer airframe. The external shell of the bay is comprised of a coupler with outer diameter equal to the inner diameter of the main parachute tube. It is inches in length. This bay coupler will be epoxied into a groove which is cut half the depth into the bulkhead attached to the payload airframe. Two aluminum threaded rods will also be used connected to countersunk rivets on the bulkhead. The tray will then slide onto these rods followed by a G10 fiberglass bulkhead which will keep the avionics bay safe from black powder charges for main parachute deployment. A detailed view of the front avionics bay is located in section D. Main Parachute and Transition Section The main parachute tube is comprised of pre-glassed phenolic tubing with an outer diameter of 6.17 inches and length 30 inches. The transition section is composed of a 4:1 ogive fiberglass nosecone with the tip cut off to allow the diameter of the main parachute tube to fit snugly. The length of the transition section is estimated to be 14 inches. A drawing of this section is given in Within the main parachute tube there is also a 6 inch piston to deploy the main parachute. E. Coupler Section The aft end of the main parachute tube is connected with shear pins to the drogue coupler. This coupler is 11 inches in length and composed of phenolic tubing with a 1 inch band in the middle to ensure that the drogue parachute deployment does not put stresses on the main parachute shear pins. The bulkhead separating the main and drogue parachutes is made of 1/8 inch thick G10 fiberglass. A detailed view of this section is shown in The drogue parachute will be housed partly in this coupler and partly in the lower airframe tube. F. Rear Avionics Bay The rear avionics bay is similar to the front avionics bay in that aluminum threaded rods connected to an aft plywood bulkhead provide stability for the sled. The sled will slide onto this and will be capped in the forward direction by an 1/8 inch G10 fiberglass bulkhead which will shield the avionics from the drogue black powder charges. G. Fins and Fin Mount Section The fin and fin mount section is composed of a 6.17 inch outer diameter pre-glassed phenolic tube with a length of 22 inches. The fin mounts will be placed in slots cut into the airframe which will provide stability and strength. The fins and the mounts are composed of 3/16 inch G10 fiberglass sheet and are detailed in University of Minnesota USLI CDR 2012 Page 16

17 H. Airbrake Section The airbrake doors are 6 inches in length and are comprised of G10 fiberglass to ensure strength. The hinges epoxied to the airbrake doors are machined with 6061 aluminum and are 4 inches in length. The actuators are connected to the doors by nylon cord. A more detailed explanation is given in Functional Requirements The system level functional requirements are found in the USLI handbook. Since the beginning of our design, we have focused on designing our systems to satisfy all of the requirements. Verification of our design features will commence after the critical design report. The following tables list major requirements, the design feature that satisfies the requirement and requirement verification. Functional Requirement The launch vehicle shall carry a science of engineering payload. The launch vehicle shall deliver the science or engineering payload to, but not exceeding an altitude of 5,280 feet above ground level. Design Feature Satisfying Requirement The vehicle has a payload bay defined in Section 3 of this report which is capable of housing our designed rover payload as defined in Section 4. The entire rocket airframe shall be matched with a motor to ensure the projected height does not exceed 5,600 feet. An airbrake system will slow the rocket down to a projected 5280 feet. How the Requirement will be Verified The ability of the launch vehicle to fly with the weight of our payload will be tested at our full scale test launch. The dimensions outlined in this document show that the rocket is capable of housing the payload. The airbrake system shall be simulated in ANSYS software prior to launch and will also be tested at the full scale test launch. Note: The design being used for the airbrakes has been previously built (using 4 inch diameter tubing versus our team s airframe diameter of 6 inches) and flown by our team lead, Mark Abotossaway. University of Minnesota USLI CDR 2012 Page 17

18 The vehicle shall carry one altimeter to be determined for recording the official altitude used in the competition. The recovery system shall be designed to be armed on the pad. Requirement The recovery system electronics shall be completely independent of the payload electronics. The recovery system shall contain redundant altimeters. Each altimeter shall be armed by a dedicated arming switch. Each altimeter shall have a dedicated battery. Each altimeter arming switch shall be accessible from the exterior of the rocket airframe. We have space designed for the official altimeter in the front avionics bay. There will be holes in the airframe just large enough to fit the head of a screw driver to turn on the avionics bays screw switches. Design feature satisfying requirement There is no electrical connection between the avionics bay and the payload. There will be two sets of redundant altimeters for the drogue and main parachute events. Avionics systems are further defined in Section Four magnetic switches will be supplied for the four separate altimeters. Four 9V batteries will be supplied for the four separate altimeters. Each magnetic switch shall be turned on with a small magnet from the outside of the airframe. Our full scale test flight as well as material compression tests will be used to verify the strength of the avionics bay. Ground tests will be run to verify proper connection of all components as well as proper open and closed circuits when turning the switch on and off. How the requirement will be verified N/A The redundancy of the altimeters and the electrical circuitry will be tested on the ground using a vacuum chamber to simulate changes in barometric pressure for the altimeters. Ground testing of the avionics bay circuitry will be done to ensure each switch is connected to only one altimeter. Ground testing of the avionics bay circuitry will be done to ensure each battery is supplying power to only one altimeter. We will do multiple tests on the strength of the magnet and minimum distance for the switches to be activated. University of Minnesota USLI CDR 2012 Page 18

19 Each altimeter arming switch shall be capable of being locked in the ON position for launch. Each altimeter arming switch shall be a maximum of six feet above the base of the launch vehicle. The recovery system electronics shall be shielded from all onboard transmitting devices. Screw switches are designed to lock into place and are not affected by the accelerative forces on launch. While the aft avionics bay is within six feet of the base of the vehicle, the distance of the front avionics bay from the base is still pending due to the variable lengths of the main parachute bay and the airbrake doors. If the final design pushes the front avionics bay above six feet, wiring and switches will run underneath the airframe s transition section to fulfill this requirement. The RF tracker is placed in the nose cone, while the closest altimeter is separated from the tracker by over 2 ft. The screw switches will be tested on our half scale and full scale test launches to ensure a closed circuit from launch to landing. A tape measure will be used to verify the distance of the switches from the base of the rocket. We will have multiple tests at a distance of 2 ft. to check if any interference is observed by the two electrical systems. The launch vehicle and science payload shall remain subsonic from launch until landing. We will be using an L-motor, which produces a max impulse of 4864 N-s. Rocksim analysis will show us our max speed throughout the flight. This speed must stay below 340 m/s. The launch vehicle and science or engineering payload shall be designed to be recoverable and reusable. All payload materials are made of metal or strong composites and rocket is made from G12 fiberglass. An RF tracker is placed in the nose cone for tracking purposes. We will do tensile and compression test on the G12 fiberglass to ensure structural strength to absorb impact. We will also be using the same RF tracker in the ½ scale testing to make sure the tracker is working properly. University of Minnesota USLI CDR 2012 Page 19

20 The launch vehicle shall stage the deployment of its recovery devices. The rocket will have 4 altimeters. Two, for redundancy, will be activated at apogee to release the drogue, and two will be activated at 800ft. altitude to release the main parachute. We will test that the altimeters are working properly in a vacuum chamber and at our full scale launch. Removable shear pins shall be used for both the main parachute compartment and the drogue parachute compartment. The launch vehicle shall have a maximum of four independent or tethered sections. At landing, each independent or tethered section of the launch vehicle shall have a maximum kinetic energy of 75 ft-lbf. The current design of deployment for the main and drogue parachutes is by using black powder charges large enough to break the shear pins used to keep the airframe together during flight. The airframe is comprised of two separate sections. 1. Nosecone, 2. Payload/main tube, drogue coupler, fin can section (these three components are tethered together) Using the estimated weights and descent rates from the Rocksim analysis, we will select properly sized parachutes to allow each section to land with less than the max kinetic energy Shear pin calculations will be tested on the ground to ensure the charges are large enough to deploy each parachute. The shear pin strength will also be designed to withstand the forces from the deceleration of the airframe due to the parachutes and will be tested at the full scale test launch. N/A We will verify our design feature during a test flight. University of Minnesota USLI CDR 2012 Page 20

21 All independent or tethered sections of the launch vehicle shall be designed to be recovered within 2,500 feet of the launch pad assuming a 15 mph wind. The launch vehicle shall be capable of being prepared for flight at the launch site within 2 hours. The launch vehicle shall be capable of remaining in launchready configuration at the pad for a minimum of 1 hour without losing the functionality of any onboard component. The launch vehicle shall be launched from a standard firing system using a standard 10 second countdown. The simulated launch angle is design to be as small as possible by changing fin design and using a motor with a high impulse to ensure stability after leaving the rail. The drogue parachute will be sized to ensure minimal drift and the main parachute deployment will be set for an altitude of 800ft which will ensure a fully open parachute and minimal drift time. Rocket design, payload integration, and avionics bay integration is currently being designed to ensure minimal preparation time. A flight checklist will also be designed to decrease setup time at launch. The altimeters used are designed to run off the 9V batteries which we are implementing into our rocket and the limit switches on the airbrake actuators will ensure minimal power drain while sitting on the launch pad. The payload electronics are also designed to be capable of being in the launch-ready configuration for at least one hour prior to launch. More payload electronics details are covered in section 4. We will be using a Cessaroni L1115 which is designed to be launched using a standard launch system. This will be verified with the help of Rocksim as well as tested at our full scale test launch. Each member of our team will be assigned a task to help with during the preparation of our rocket for flight. Each member will also practice there portion of preparation during the full scale test launch. We plan to run a ground test to make sure all components are still in working order with sufficient power for at least one hour after being in launch-ready configuration. Our motor and firing system will be tested at our full scale launch test. University of Minnesota USLI CDR 2012 Page 21

22 The launch vehicle shall require no external circuitry or special ground support equipment to initiate the launch. Data from the science or engineering payload shall be collected, analyzed, and reported by the team following the scientific method. 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. The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant which is approved and certified by the National Association of Rocketry, Tripoli Rocketry Association, and/or the Canadian Association of Rocketry. Our rocket only requires a standard firing system and requires no external circuitry or special ground support equipment to initiate launch. The payload will transmit scientific data outlined in section 4 which will be analyzed and reported by our team following the scientific method. An RF tracker will be installed in the nosecone of the airframe as well as housed within the payload. We will be using a commercially available Cessaroni L1115 solid motor which is certified by the National Association of Rocketry, Tripoli Rocketry Association, and/or the Canadian Association of Rocketry. N/A The payload data transmission and collection will be tested on the ground at distances up to 2,500 feet prior to launch. Each transmitter/receiver will be tested prior to launch as well as at the half and full scale test launches to verify their capabilities. N/A University of Minnesota USLI CDR 2012 Page 22

23 The total impulse provided by the launch vehicle shall not exceed 5,120 Newton-seconds (Lclass). All teams shall successfully launch and recover their full scale rocket prior to FRR in its final flight configuration. Flashbulbs, forward canards, forward firing motors, rear ejection parachute designs, motors which expel titanium sponges, and hybrid motors are prohibited from use in the launch vehicle. Each team shall use a launch and safety checklist. 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. The Cessaroni L1115 solid motor has an impulse of 4864 N-s. We will test our design near the end of February with a mass matching that of the payload in the payload bay. Flashbulbs, forward canards, forward firing motors, rear ejection parachute designs, motors which expel titanium sponges, and hybrid motors are not used in our design. Our team will design a launch and safety checklist to include in our flight readiness report. All design, construction, written reports, presentations, and flight preparation has and will continue to be completed strictly by student members on the team with exception of assembling the motors and handling of black powder charges. N/A We have designated a strict schedule to ensure completion of the construction of our design by our preliminary test launch date. N/A We will have our team mentor, Gary Stroick, look over our launch and safety checklist to ensure we are following necessary safety precautions. Our team official, William Garrard, and our team mentor, Gary Stroick, have been informed of this requirement and are able to give a written statement attesting to the verification of the requirement. University of Minnesota USLI CDR 2012 Page 23

24 The rocketry mentor supporting the team shall have been certified by NAR or TRA for the motor impulse of the launch vehicle and shall have flown and successfully recovered a minimum of 15 flights in this or a higher impulse class, prior to PDR The maximum amount teams may spend on the rocket and payload is $5000 total. This includes component and material prices while excluding shipping costs, ground support equipment, and team labor. Our team mentor, Gary Stroick, has been involved with high power rocketry since 1998 and is certified level 3 with the Tripoli Association of Rocketry. He also holds a low explosives import and user permit. Currently a board member of Tripoli Minnesota and was the vice president from 2007 to Our budget is displayed in Section 5 and it can be verified that we are within the $5000 on the pad. N/A N/A Table 2: Table of Functional Requirements Workmanship The degree of skill that is required to fabricate and assembly our launch vehicle and payload is essential to our mission success. We realize that every single component on the vehicle is equally important. Given that we are a rookie team and most of our team members lack experience with rocketry and machining we have limited the number of parts and systems that will be custom manufactured by us. Although we have relied mainly on pre-purchased and prefabricated components, the parts that we machine will have to be done so carefully. The most experienced team members will do most of the machining. University of Minnesota USLI CDR 2012 Page 24

25 With regards to assembly, we plan on carrying out trial runs and test fits whenever possible. We will also experiment with various epoxy application techniques and painting and finishing techniques as well, before we apply them to our full scale vehicle Testing Our test schedule following the CDR will be intense. There will be full component testing followed by functional testing and static testing of most of our vehicle parts. Component testing will begin when parts start to arrive. Each system lead is responsible for their own systems tests and the team will decide priority and necessity of all tests, whether structural or functional. Structural tests will consist of calculating or verifying the integrity of our fiberglass and plywood bulkheads and centering rings Our altimeters will be tested in vacuum chambers multiple times with various programming to ensure correct operation in all configurations. We will need to build the vacuum chamber. Our airbrake system will be tested on a half scale flight prior to full scale to ensure correct operation and no mechanical issues. The test will also help to confirm our drag calculations. Our payload ejection system will be tested to confirm the proper amount of black powder charge and to ensure structural integrity of the airframe and payload following the event. Payload ejection tests will also be carried out to test our remote detonation system. The team still needs to fabricate a mock payload for static tests. Ground tests will be carried out on the parachute ejections. We will confirm the amount of charge required and make sure the tube connections release smoothly, and the number and size of the shear pins is correct. Our magnetic switches will be tested for durability and sensitivity. Our RF trackers will be tested for range and signal strength. These tests will most likely be carried out at the TRA launch range at North Branch, Minnesota. Payload tests are discussed in Section 4 (Payload). A detailed timeline of test schedule can be found in Section Manufacturing and Assembly All manufacturing and assembly of the launch vehicle remains to be completed. This will require effective communication and planning among team members as well as effective utilization of workspace, workshop and all other facilities. University of Minnesota USLI CDR 2012 Page 25

26 Airframe tubes will need to be cut to length in the machine shop or by hand with a miter box to ensure squareness. Multiple bulkheads and centering rings will need to be cut to specification in the machine shop on either the lathe or a milling machine. The airbrake hinge and door attachments need to be completed. The avionics bays need to be assembled. The pistons need to be assembled. The fin mounts and fins need to be cut. For payload manufacturing and assembly see Section 4 (Payload). For a detailed timeline of manufacturing and assembly see Section 5. Our primary shop will be the Mechanical Engineering Student Machine shop which has is open various hours weekly for a total of 12 hours per week. We also have access to the Physics Student Machine shop if needed. These will be used for detailed custom machining or precision machining. Our workspace in Akerman Hall is equipped with various hand tools for small systems assembly and finishing as well as testing. We have access 24 hours a day. A few team members also have access to other student project shops and workspace. University of Minnesota USLI CDR 2012 Page 26

27 3.1.8 Integrity of Design NOSECONE AND PAYLOAD SECTION Figure 2: Nosecone and Payload Assembly Our team will be mounting an RF tracker with Velcro inside the 4:1 Ogive fiberglass nosecone with the base of the tracker placed against the nosecone bulkhead. The opposing side of this bulkhead will be fit up against the payload. A more detailed description of the payload bay design is given in the Payload Integration Section 3.5. Aft of the payload piston is the bulkhead which forms part of the enclosure for the front avionics bay. The outer airframe tube of the payload section is pre-glassed phenolic tubing which will be sufficient to handle the loads during flight as well as forces exerted during payload deployment. University of Minnesota USLI CDR 2012 Page 27

28 FRONT AVIONICS BAY Figure 3: Front Avionics Bay Assembly The front avionics bay has an outer shell composed of phenolic whose outer diameter is equivalent to the inner diameter of the main parachute airframe tube. Black powder charges will be wired through the main parachute side bulkhead which is composed of G10 fiberglass. The load bearing rods are threaded aluminum and will keep the avionics bay closed throughout flight as well as keeping loads from compressing or straining the avionics board. To assemble for flight, the avionics sled will be loaded in the bay followed by securing the main parachute side bulkhead onto the threaded rods. Then the main parachute tube will slide over the avionics bay while the transition connects with the outer payload tube. Washers will be used in congruence with nuts on all threaded items to properly distribute the loads. University of Minnesota USLI CDR 2012 Page 28

29 MAIN PARACHUTE TUBE AND TRANSITION Figure 4: Main Parachute Tube and Transition Assembly The main parachute tube is composed of pre-glassed phenolic which will be capable of withstanding all loads during flight and landing. The transition is composed of fibreglass but will not be a load-bearing section. All of the distributed load will be applied to the main parachute tube held within the transition. University of Minnesota USLI CDR 2012 Page 29

30 DROGUE COUPLER Figure 5: Coupler Between Main Parachute Main Parachute and Fin Section The coupler separating the main parachute section from the motor and fin section is composed of phenolic with a bulkhead of G10 Fiberglass. Washers will be used with the threaded U-bolts to distribute loads from the parachutes. The band around the coupler is used to ensure charges for the drogue parachute do not affect the shear pins of the main parachute tube which could cause failure to separate. University of Minnesota USLI CDR 2012 Page 30

31 Figure 6: Rear Avionics Bay AFT AVIONICS BAY The aft avionics bay is similar to the front avionics bay in that it threaded aluminum load bearing rods provide strength to distribute the load of the parachute deployment away from the avionics bay. The sled is 3 inches in length and is made of G10 fiberglass. The rods are secured with nuts and washers to distribute the load to the 1/8 inch G10 fiberglass bulkhead which drogue charges will be wired through. The rear bulkhead is comprised of plywood. University of Minnesota USLI CDR 2012 Page 31

32 MOTOR MOUNT Figure 7: Motor Mount System Assembly Figure 8: Motor Mount Cutaway The design for the motor mount system has not changed since PDR. The current design for our motor mount system is removable to allow access to the actuators for our airbrake system. The foremost centering ring connected to the airframe has an inner University of Minnesota USLI CDR 2012 Page 32

33 diameter equivalent to the outer diameter of our motor casing to assure proper centering of our motor while the foremost centering ring attached to the inner motor mount tube will press against the former ring to distribute the load of the motor. The two centering rings aft of the front rings will be bolted together using three countersunk rivets in congruence with three machine screws to assure a secure fit of the motor mount to the airframe. Aft of those centering rings is the airbrake section followed by the boat tail added to reduce drag. All centering rings in the above diagram are fabricated out of plywood. FIN AND FIN MOUNT SYSTEM The full scale fins have a trapezoidal shape shown in Figure9 with chamfered trailing and leading edges made of G10 fiberglass with a thickness of in. Iterative processes done in Rocksim allowed our team to find an optimal sweep angle for maximum altitude. Flutter speeds are obtained at 71% of the speed of sound for our fin design, but the rocket will only obtain max mach number of.57, thus flutter will not be an issue. Our fin mount system allows for different fins to be easily interchanged. This allows our rocket to have a modular effect; however, due to safety concerns affecting the change in the rocket s static margin, the full scale rocket will only have one set of fins that were previously discussed that will be flown for this competition. Figure 9: Fin Dimensions University of Minnesota USLI CDR 2012 Page 33

34 Figure 10: Fin Mount Dimensions The fin mounts for our rocket will be created using three pieces of 3/16 inch G10 fiberglass sheet that have been epoxied together as shown above. A 1/4 inch tab will be placed into slots in the airframe for added strength and stability while the entire fin mount will be epoxied to the airframe. The leading and trailing edge of each fin mount will be chamfered to reduce drag. A similar scaled down design was flown on our subscale rocket and we had visual confirmation that the design did not fail. 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 in Figure 11. University of Minnesota USLI CDR 2012 Page 34

35 Figure 11: Schematic of Airbrake Assembly Construction and Sizing The airbrake doors are composed of reinforced G12 fiberglass 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. Door sizing is set to 6 inches but may change after dynamic load testing is completed with the linear actuators. Nylon cord will be used to connect each door with their respective linear actuator. A second redundant cord will connect the door with the forward bulkhead to relieve stresses to the open actuators. The actuators will be mounted to the motor mount tube, and the actuator arm length is set to 140mm using Firgelli L16 linear actuators. A picture of the Firgelli L16 can be seen in Figure 12. A gear ratio of 63:1 will be used giving a maximum lifting force of 100 N. University of Minnesota USLI CDR 2012 Page 35

36 Figure 12: Firgelli L16 Linear Actuator Wiring and Component Specification 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 dual 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 in Figure 13. The load and current curves for the actuator and the specifications of the latching relay can be seen in Figure 14. University of Minnesota USLI CDR 2012 Page 36

37 Figure 13: Airbrake Wiring Schematic Figure 14: Load and Current Curves for the L16 Actuator (Provided by Firgelli) University of Minnesota USLI CDR 2012 Page 37

38 Relay Part Number Relay Type Coil Current Coil Voltage Contact Form Contact Rating Operate Time Release Time TX2-LT-9V-TH Latching, Dual Coil 15.5 ma 9V DPDT 2A 4 ms 4 ms Figure 15: Latching Relay Specifications Airbrake Analysis 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 16, 17 and 18 show the distance, force per airbrake door, and rocket deceleration Altitude Profile Altitude (ft) Time (sec) Figure 16: Distance profile found for the rocket without airbrakes (blue), with variable airbrake opening velocity (green), and instantaneous airbrake opening velocity (red) University of Minnesota USLI CDR 2012 Page 38

39 Figure 16 shows three profiles for distance of our rocket. The highest altitude, shown in blue, shows our final altitude of 5850 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 4300 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 5050 ft and 5280 ft. 60 Force per Brake Force (lbf) Time (sec) Figure 17: This shows the force on each airbrake door for instantaneous brake deployment (red), and variable speed brake deployment (green) Figure 17 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 mitigation for damage is to make the redundant nylon cord capable of withstanding a minimum of 54 lbf. The forces applied to the airbrake doors when the speed of the actuators is applied gives a University of Minnesota USLI CDR 2012 Page 39

40 maximum force of 19 lbf while the linear actuators are capable of withstanding 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 load before damage. The actuators are capable of loads greater than this without damage when the load is applied in the direction of motion and they will resist motion with a back drive force of 10.3 lbf. 5 Deceleration Deceleration (g) Time (sec) Figure 18: This shows the rocket deceleration without brakes (blue), with variable opening speed brakes (green), and instantaneously opening brakes (red) Figure 18 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. POWER ANALYSIS While PDR showed concerns that the relay connecting the avionics bay with the airbrake system could cause unwanted battery drainage, the design has been changed to include a latching relay which will be able to be switched on and off using the pyro channels 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 University of Minnesota USLI CDR 2012 Page 40

41 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 Mass Statement A mass balance statement is shown is tabular form below. SubAssembly Component Qty Length (in) Weight (oz) Margin Max Wt Cone Nosecone Rf Tracker Bulkhead Payload Airframe Tube Payload Payload Piston Payload Bulkhead U-bolts Bulkhead Av Bay Av Sleeve Av Sled Altimeter Altimeter Battery Battery Threaded Rods Bulkhead Inner Bulkhead Outer U-bolt Upper Booster Airframe Tube Coupling CR Transition Switches Recovery Main Piston Main Bulkhead U-bolts Main Parachute Main Shockcord Coupler Band Coupler U-bolt Bulkhead Recovery Drogue Parachute University of Minnesota USLI CDR 2012 Page 41

42 Drogue Shockcord Deployment Bag Lower Airframe Tube Av Bay Front Bulkhead U-bolt Av Sled Altimeter Altimeter Battery Battery Airbrake Altimeter Battery Threaded Rods Switches Rear Bulkhead Tracker Motor Body CR MMT CR Motor Mount Tube Actuators Body CR MMT CR Surface Fin Mounts Fins Airbrake CR Airbrake Assembly Airbrake Doors Motor Casing Motor Motor Retainer Boattail Total Components 79 Total Length Total Weight Maximal Weight Table 3: Mass Balance Statement University of Minnesota USLI CDR 2012 Page 42

43 Using Rocksim calculated weights as an initial starting point we began a mass balance analysis. Although our Rocksim model was accurate, we did not include every small item such as bolts and washers and screws. Initially, we wanted to confirm by independent calculations, that the Rocksim weight projections were correct. As we finalized our design, we switched the Rocksim weights with a mass override to the estimated or stated weights of the products we ordered. Some custom parts that are still to be made could not be accounted for in Rocksim correctly though. There are higher margins on these parts since we do not know if Rocksim accurately describes our part. The total estimated mass can be seen in the table as ounces. This does not include every single bolt or smaller component on the vehicle. As such, we added 32 ounces to the total from the table to end up with ounces. We believe this will account for the external finishing and the internal epoxy as well. The total estimated weight in the table also coincides with the Rocksim vehicle weight. We believe our mass estimate is fairly accurate and will perhaps increase slightly after final assembly. Using the margins of growth for each component, we also calculated an estimated maximum vehicle weight of ounces. Aside from any design changes we have selected a motor which will allow us to put the maximum weight vehicle to the expected altitude. Our vehicle is not particularly sensitive to mass growth. Given that we plan to overshoot the expected altitude and use our airbrake system to slow the vehicle to the exact altitude, we have overpowered our vehicle. Without the use of the airbrake system, our vehicle could carry an extra 4 pounds (around 10% of our total weight) before becoming too heavy and not reaching our intended altitude. Our rail exit velocity will also decrease at this weight, but still be at a safe level (60 feet per second on a 120 inch rail). The mass balance statement table also gives us our section weights upon landing to help calculate our kinetic energy. The upper section weight (forward of the coupler) is oz while the rear weight is oz. For safety and failure analysis, see Section 3.7. University of Minnesota USLI CDR 2012 Page 43

44 3.2 Subscale Flight Results Following the preliminary design report in November 2011, we began to design a half scale launch vehicle to test flight performance and a few other systems. Starting with the full scale Rocksim design from the PDR, we reduced the geometric dimensions of the vehicle by half. Due to the non-linear scaling of the avionics, parachutes and motor, the half scale design was similar in shape to the full scale, but not an exact half scale replica. We decided that the geometric scaling was most important for flight performance tests and flight stability tests. Given that we are purchasing stock fiberglass tubes for our airframe, we needed to round our half scale airframe diameters to an existing airframe diameter. This lead our booster tube to be 3.5 inches diameter, with 4.0 inches diameter payload tube. The size of the booster tube dictated our motor selection. We found that there was no motor that would duplicate our predicted full scale flight performance by half, with the booster tube size limitation. It must also be noted that the half scale weight was less than half as well, so a 54mm 3 grain motor was selected for half scale flight which would give us an apogee altitude of approximately 2500 feet. The parts for our scale rocket were ordered in mid-december, as well as a few payload parts for initial testing as well. Construction began on all custom parts at the end of December and continued in the beginning of January. However, due to misinformation on lead times, airframe tubes had not arrived at the end of the first week of January. For the sake of staying on schedule, it was decided to build custom tubes for our scale rocket, and complete a flight test prior to CDR. At the same time, we had contemplated switching airframe materials on the full scale design to save weight and ensure the target altitude was reached. Pre-glassed phenolic tubes from Public Missiles Ltd were selected for the full scale to coincide with the custom airframe tubes that we would have to fabricate for the scale vehicle. Since phenolic tubes do not come in a stock size of 3.5 inches in diameter, we decided we would cut a tube and reduce the area to our dimensions. With such short notice, we could not find PML phenolic tubes and chose to use LOC Precision tubes. Final vehicle fabrication and assembly was completed January 14 th and the flight test was completed January 18 th. Due to multiple systems failures however, the team failed to recovery the vehicle. Although there is no flight data, we can confirm that the scale launch vehicle was structurally and aerodynamically sound. The flight failed when both parachutes failed to eject from the tubes, despite confirming proper amounts of FFFFg black powder were used to deploy the parachutes in ground testing. Recovery failed due to a failed (or University of Minnesota USLI CDR 2012 Page 44

45 damaged) RF tracker. Due to a heavy wind-cock the vehicle did pitch a few degrees off vertical when it cleared the launch rail, and due to the ballistic trajectory, was last seen headed south. An hour and a half search was done, but had to be ended due to limited sunlight. We are currently making plans to recover the vehicle and build an all new airframe. The team is still waiting to receive the initial fiberglass airframe tubes purchased in mid- December. Pending recovery of the initial scale launch vehicle and analysis of the failures, modifications will be made to the second scale launch vehicle. We will also design the second scale launch vehicle to accommodate the airbrake system, which we will test in our next flight. University of Minnesota USLI CDR 2012 Page 45

46 3.3 Recovery Subsystem Parachute Selections Theoretical Calculations The parachutes are selected based on the kinetic energy requirement at landing, which is 75 ft - lbf for each section, keeping in mind that the drift in windy conditions should also be minimal as we need to land no further than 2500 feet from the launch tower. The weight of the rocket at upper section is 22lbs and lower section is 14lbs. If we use the equation (3.2.10) in PDR, we will be able to determine a parachute we can use as a main. Assumptions made to determine the parachute size are that the air density is at sea level, and the descending rate of the parachute is steady. At sea level, density is , and the kinetic energy requirement is 75. where is called nominal fully inflated drag area. Since is proportional to, as long as the kinetic energy requirement of upper section is met, lower section is automatically satisfied. The for upper section of the rocket is Therefore, we need to choose a main chute whose a nominal fully drag area is more than 84.58ft^2. According to table 4, Iris Ultra chute diameter 96 works for the rocket without breaking the kinetic energy requirement. Rocksim Analysis From the theoretical analysis above, we can choose Uris Ultra chute diameter 96. However, from the analysis of RockSim, we are not able to satisfy the kinetic energy requirement at landing. According to the analysis, a vertical velocity at landing is 16.95ft/s and the maximum weight of each section for this velocity is 17.2lbs. A chute larger than this size is one whose diameter is 120. University of Minnesota USLI CDR 2012 Page 46

47 From the analysis of Rocksim, a vertical velocity at landing is 14.3ft/s, and the maximum weight of each section for this velocity is lbs. Therefore, this chute satisfies the kinetic energy requirement. If you would like to know more about kinetic energy at landing, please go to section Assumptions Made for RockSim Analysis The drogue parachute deploys at apogee, and the main parachute deploys at 700feet altitude. The size of drogue chute chosen was 2 foot diameter Rocketman chute. A steady state descending rate for this drogue chute is ft/s, and an assumption was made that it has since specifications cannot be found. The simulation was run in both light wind conditions (3-7 MPH), and slightly breezy wind conditions (8-14 MPH). Diameter Effective drag area ( ) Nominal fully inflated drag area, ( Table 4: Iris Ultra Chute Data Parachute Ejection Both main and drogue parachutes will be deployed from the rocket using black powder charges. The main parachute will also have the assist of an ejection piston. The canisters will be assembled in the same way as payload ejection, shown in section 3.5, except they will be set inside the rocket bulkheads instead of a piston bulkhead. The amount of black powder required will be found using University of Minnesota USLI CDR 2012 Page 47

48 where V is the volume of the payload bay, P is the pressure, R is the gas constant for air, and T is the burn temperature of black powder. The value for R is 266 in-lb f /lb m and T is 3307 Rankine, according to the USLI sponsored booklet. The purpose of the black powder charge is to break the three keyhole shear pins, rated at 65 pounds, which are inserted into the rocket to hold the bays together during flight. The drogue parachute bay has a diameter of and a length of 6 which results in.7 grams of black powder. The main parachute bay has a diameter of and a length of 14 which results in 1.7 grams of black powder., Parachute Components U-Bolts From PDR, the maximum opening force can be calculated using equation (3.2.11). The calculations results are shown in table 5. According to the table, If we use Iris Ultra chute 120, the load of lbs will be exerted at maximum. We need to choose components whose breaking loads are ideally greater than the maximum opening force. However, if the components are too large to fit, we could use the components whose breaking loads are close or little less than the maximum opening force. The maximum opening force for drogue chute cannot be calculated using equation (3.2.11) due to the lack of availability of information. However, one thing we can say is that the maximum opening force for a drogue chute is not as big as main chute. Selected U-bolt is shown in the table below A Pipe size B C Workload limit 3 2 ½ 4 ½ 1 5/8 2020lbs Shock Cord Shock cord was chosen based on the value of maximum opening force. Fruity chute sells a shock cord, which can withstand the force up to 1800lbs. From the rule of the thumb, the cord length is three to four times the length of the rocket; the length of the rocket is 117 inch, and four times of the value is about 444 inch that is yards. University of Minnesota USLI CDR 2012 Page 48

49 Name of the cord Nylon webbing shock cord Size of the cord 5/8 Workload limit 1800lbs Quick Links The quick link was selected using the value of maximum opening force and the size of the shock cord. Because Fruity Chute does not tell us the thickness of the cord, we made an assumption that the thickness is 0.5 inch at maximum. The opening, A in the figure needs to be equal or larger than 0.5 inch, and the value of B cannot be very big. As I mentioned this above, the workload limit needs to be more than 1378lbs of force. Dia A B C Workload 3/8 1/2 2 7/16 3/ lbs S(ft^2) Cd v (ft/s) density (slug/ft^3) load (lbs) Iris Ultra Iris Ultra Iris Ultra , Iris Ultra , Iris Ultra , Table 5: Maximum Opening Force for Various Chutes University of Minnesota USLI CDR 2012 Page 49

50 3.3.4 Recovery Hardware RF TRANSMITTER We selected the PT-1B RF transmitter for our tracker. This transmitter has a stated range of 2 miles. Since the transmitter is only 0.45 oz, one will be mounted in the nosecone with Velcro and another in the motor mount section. If the deployments of the chute were to fail, the nosecone would impact the ground at landing, and the section of tailcone would be the highest point of the rocket above the ground. As long as one of the transmitters is above the ground, we would be able to locate and recover the rocket. Figure 19: PT-1B Transmitter 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. Components Quantity Featherweight Raven2 Altimeter 2 PerfectFlite StratoLogger Altimeter 2 Entecore Electronics AIM USB Altimeter 1 9 Volt Battery 4 Magnetic Switch 4 Table 6: Electrical Components for the Avionics Bay University of Minnesota USLI CDR 2012 Page 50

51 Figure 20: Featherweight Raven2 Altimiter Manufacturer Featherweight Altimeters LLC Outputs 4 programmable outputs 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 Dimensions 0.8 x 1.8 x 0.55 Table 7: Featherweight Raven2 Specifications Figure 21: PerfectFlite Stratologger 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 8: Featherweight Raven2 Specifications University of Minnesota USLI CDR 2012 Page 51

52 Figure 22: Entecore Electronics AIM 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 9: Enticore Electronics AIM USB Altimeter Specifications 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 from Aerocon Systems will be implemented to be turned on with a magnet from the outside of the airframe at launch. The wiring schematics are shown in Figures 23, 24 and 25 for each avionics bay. University of Minnesota USLI CDR 2012 Page 52

53 Figure 23 Front avionics bay wiring schematic Figure 24: Aft avionics bay wiring schematic for recovery University of Minnesota USLI CDR 2012 Page 53

54 Figure 25: Aft avionics bay wiring schematic for airbrake system Figure 26: Magnetic switch 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. University of Minnesota USLI CDR 2012 Page 54

55 3.3.5 Kinetic Energy at landing Upper section of the rocket is 22 lbs, and lower section of the rocket is 14 lbs. The scientific requirement of the kinetic energy at landing for each section is 75ft-lb. The equation of the kinetic energy, is given by where m is the mass of each section, and is the velocity at landing. At landing, the velocity would be 14.31ft/s. Kinetic energy for the upper section of the rocket is Kinetic energy for the lower section of the rocket is Since the values of kinetic energy of upper and lower sections are lower than 75ft-lb, we are meeting the kinetic energy requirements. The distance of landing from launching site for each wind speed is shown below. Wind speed(mph) Drift (ft) Table 10: Rocksim Simulation Results When the simulations were run, Cesaroni L1115-Classic engine was chosen. Launch guide was set to 120 inch. The Drogue chute diameter was 24 inch and was set as 1.2, and this deployed at apogee. The main chute diameter was 120 inch and its spill hole diameter was 24 inch, and was set as 2.2. This deployed at 700ft altitude. Since the drift of each wind speed is less than 2640ft (0.5mile), we are meeting the drift requirement. University of Minnesota USLI CDR 2012 Page 55

56 3.4 Mission Performance Predictions PERFORMANCE CRITERIA The rocket must travel as close to an altitude of 5280 ft. Final performance standards predict the full scale rocket to travel to an altitude between 5050 ft. and 5280 ft. with the airbrake activation. Drogue and main parachutes must be successfully deployed on decent, while providing 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 then later analysed by the team. MOTOR SYSTEM The motor system must be capable of withstanding an impulse of up to 4900 N-sec and a maximum thrust of up to 3700 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. 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. AIRBRAKE SYSTEM The airbrake system must be capable of opening without damage to 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. The airbrake system must also be capable of slowing the rocket down to achieve a projected altitude of 5280 ft. 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 University of Minnesota USLI CDR 2012 Page 56

57 landing. The system must be programmable to supply enough voltage to run the various tasks at user set altitudes or times. FLIGHT PROFILE SIMULATION A flight profile simulation from Rocksim is shown in Figure 27. The altitude with no airbrake deployment, velocity, acceleration, burnout, and apogee time are shown. Figure 27: Flight Profile Rocksim Simulation Preditced altitude resuts from Rocksim with no airbrake deployment are shown in Figure 28. The mean altitude is predicted to be After paint and finish mass is added the rocket and air brake deployment is factored, the rockets new altitude height should be between 5050 ft and 5280 ft. ALTITUDE PREDICTIONS Figure 28: Rocksim Altitude Predictions University of Minnesota USLI CDR 2012 Page 57

58 MOTOR THRUST CURVE Figure 29: Thurst Curve for the Cesaroni L1115-Classic motor. DRAG ASSEMENT All drags were calculated in Rocksim. Figure 30 shows the percentage of each subcomponent (nose + body, base, fins, and lugs) effect on Coefficient of drag C d and also predicts the full scale s C d. = University of Minnesota USLI CDR 2012 Page 58

59 Figure 30: Drag Assessment Rocksim Results CENTER OF PRESSURE (CP) AND CENTER OF GRAVITY (CG) LOCATIONS The CP and CG are shown in Figure 31. The static margin of the full scale rocket is The CG and CP are located 71.9 inches and 78.1 inches from the front of the nose cone, respectively. After the Cesaroni L1115-Classic motor is depleted, the static margin remains at Figure 31: Location of CG (Blue) and CP (Red) University of Minnesota USLI CDR 2012 Page 59

60 3.5 Payload Integration Payload Integration Assembly The payload bay will be located between the front aviation bay and the nose cone. It shall be separated from the av bay by a payload bulkhead and from the nose cone by a.5 wooden centering ring. All bulkheads and centering rings will be.5 thick and made of wood. The payload ejection is a similar system to the main parachute ejection, which is why this method was preferred over others. The rover, sitting inside of an ejection piston, will be propelled out of the rocket using a black powder charge, once the rocket has fallen to the ground. The rover ejection will be manually triggered once the RSO has confirmed it is okay to do so. This is done by having the e-match which triggers the black powder charge to be connected to an electronic switch on the rover. The switch will be turned to the on position by the rover controller and will ignite the charge. A more in depth explanation of how this works is shown in section Because there is the chance, however unlikely, that this ejection happens mid-flight, the nose cone has an additional RF tracker in it. Figure 32 shows the assembly of this system. The rover would fit snugly between the piston centering ring and the nose cone centering ring. The rover is sitting inside of the ejection piston to be more efficient with space. The piston is 7.5 in length, the same as the diameter, which is a general rule of thumb to prevent at against cocking, which would cause unwanted friction. Figure 32: Payload Integration Assembly To allow the rover to exit the ejection piston, the piston bulkhead will be connected to the payload bulkhead via u-bolts and a shock cord and will stop the piston from exiting the rocket. This is a tricky assembly because we cannot put our hand inside of the tube to assemble this because the piston will be in the way. To do it, we will have to first University of Minnesota USLI CDR 2012 Page 60

61 connect the two u-bolts together with the shock cord and fully assemble the piston. Then, when the piston is ready to be put into the rocket, we will lightly tape the payload bulkhead u-bolt to the piston s u-bolt. We can then slide the piston into place and feed the payload bulkhead u-bolt through its holes in the payload bulkhead. Then we just have to screw the washers and nuts onto the threads and the system is in place. If the system somehow becomes damaged or we wish to change something we can unscrew the nut and pull out the piston. The main concern with this is that the black powder energy now gets trapped inside the rocket, but the u-bolts are rated to 430 pounds so they will be able to support the 278 pounds black powder charge, which will be discussed shortly. Also, the u-bolts will be made of stainless steel because of the direct exposure to black powder. This ejection piston is slightly more complicated than the parachute ejection piston due to the rover having to sit square inside of the rover to prevent cocking during ejection or tangential forces acting upon it during flight. These worries are not needed with shape shifting parachutes.. Figure 33 shows the piston assembly. Figure 33: Ejection Piston Assembly University of Minnesota USLI CDR 2012 Page 61

62 The addition of the centering ring was designed specifically for the rover to sit against to avoid conflict with the black powder canisters or u-bolts. The opposite wheel, in the nose cone, will butt up against a centering ring for this purpose. When assembling the rocket, the nose cone centering ring will be custom fit to the rover to make sure the rover sits snug inside of the rocket and there is no shifting of weight during flight The black powder canisters are an improvement upon previous year s canisters, which were simply plastic tubes that were taped to the bulkheads, with an e-match epoxied into the rear. Figure 34 shows how the black powder canisters are put together. Figure 34: Black Powder Canister University of Minnesota USLI CDR 2012 Page 62

63 The canister itself is made of aluminum and will be made out of 1 diameter aluminum rods. All but the largest step of the canister will be lathed down to a diameter of.8. The outside of the lathed section will be threaded. The lathed down section will be put through a.85 hole in the piston s bulkhead and a modified nut will be used as a cover. The inside of the cover will also be threaded and will mate with the canister. When they are mated together the step and cover will keep the canister in place during the black powder explosion. The cover will have a 1/8 hole one the rear side which will allow the e-match to be fed through. This e-match will be epoxied into the hole. Because e-matches are one time use only wire, the cover and e-match will have to be replaced every flight or test. These will be easy to switch out and we will make plenty of covers to allow this. The canister are placed near the middle of the bulkhead so when the charges fire, they do not cock one of the bulkheads, either the one it is sitting in or the payload bulkhead. The benefit of this system is that we will not have to replace the canister every flight like we had to in previous designs and because of our piston design, this is crucial Shear Pins and Black Powder Required Three keyhole shear pins are required to hold the system in position during the flight of the rocket. These keyhole shear pins, as shown in Figure 35, are used instead of typical 2-56 or 4-40 shear pins because they are more likely to stay in place during flight. Given the large G forces the rocket will be under as well as landing, these were preferred. The shear pins are rated to a force of 65 pounds each. Figure 35: Keyhole Shear Pin The primary reason we are using three shear pins is to prevent cocking of the nose cone when the charges are fired. We did calculations to show we do not need a fourth shear pin to hold the nose cone, payload, and piston in place during flight. According to RockSim, the largest acceleration spike, up to 19 G, is when the main parachute opens. At this time the nose cone will be facing downwards, therefore this is the worst case scenario for the shear pins. Using this acceleration and the mass of the nose cone, University of Minnesota USLI CDR 2012 Page 63

64 rover, and piston, the force applied to the shear pins can be calculated. The weights of the nose cone, rover, and piston are 1.5 lb, 7 lb, and.8 lb, respectively. An additional.25 lb was included to account for epoxy and possible additional weights not initially accounted for. This calculation is where m is the mass of the components in pound-mass and is the acceleration of gravity in ft/sec. Dividing this force by the 65 pounds shear force required to break the shear pin shows 3 pins are needed (2.79 exactly). A force will also be required to push the rover out of the rocket tube. Two possible worst case scenarios were present. The first, being if the rocket landed with the nose cone facing an obstruction such as a rock or high inclined hill; The second being if the rocket landed with the nose cone facing directly upward. The latter was chosen because if there were to be an obstruction in the way, the black powder charge would force the av bay to move backwards instead of the payload to move forwards. Because the av bay is lighter, this should not be a problem. This would still release the rover from its casing and allow it to roam freely. Therefore, in the worst case scenario, the black powder charge would have to provide enough force to break the shear pins and move the rover from inside the rocket to directly above the rocket and then have it fall to the ground. To find the force required to push the rover upwards conservation of energy is used. Initially the rover will have kinetic energy due to force from the black powder charge, kinetic energy is where v is velocity. Once the rover has reached its maximum height, it will have only potential energy, where h is the maximum height of the rover. Using these two equations it can be seen that Important to note is how mass is not a factor in determining required initial velocity. Height will be 2.4 feet (28.8 inches), which will result in a velocity of 12.5 ft/sec. Height is determined by the length of the payload bay subtracted by the length of the bay the rover does not have to travel, which is the length of the piston s bulkhead to the av bay s bulkhead. The length is about This length will get the rocket just above the rocket, but an additional 14.6 will be desired to confirm the rover will be able to fully eject from the payload bay and fall to the ground. With a velocity known, total momentum of the ejection system can be known. Momentum is mass multiplied by the velocity. Because the nose cone and ejection University of Minnesota USLI CDR 2012 Page 64,,.,

65 piston will be exiting the payload bay, and in the worst case scenario, the nose cone and piston will also travel 28.8 inches upwards, their masses will have to be included along with the rover s. The impulse force equation, thus, is used to calculate the required force, where Δt is how long the force is applied for. This time is currently an unknown because it cannot be calculated, because it is how long the black powder force is acting upon the piston. This force only acts upon it when the piston is inside of the tube. From simply observing previous black powder ejections a Δt of.1 seconds seems more than reasonable. This time would seem to be low, but will count as a worst case scenario. With this Δt a force of 37 pounds is required. Total force required to push out the payload system is 232 pounds, with an additional 20% for contingency the force become 278 pounds. The contingency is to account for pressure escaping around the piston, friction, and manufacturing error in the shear pins. Using the formula the number of grams of black powder can be found. With an inner diameter of the payload bay being 7.5, this results in an area of 44.2 in 2 and a pressure of 6.3 psi. With a length of in the payload bay, the total volume of the payload bay will be in 3. These values result in 2.1 grams of FFFFg black powder being required. Given how many calculations were made to come to this result, this number is more of an estimation than trusted number. Ground testing will have to be done to find a more correct amount.,,, University of Minnesota USLI CDR 2012 Page 65

66 3.6 Launch Concerns and Operations Procedures On a launch day (test flight or competition) all team members will be responsible for their independent systems. The team lead and the team mentor will supervise. Launch Readiness Component Checklists Prior to leaving the University work shop, component checklists will be completed by each team member. Each system will have a component and tool checklist, and a toolbox. The various checklists can be seen in the Appendix. A second team member will be present to assist with checklists, and the system lead will sign off on the completed checklist. Preflight Procedures and Checklists Upon arrival at the launch site, all system boxes and complete airframe will be brought to prep table. If the site does not have available launchers, the team member in charge of launch system will proceed to assemble and position our launch rail in the designated area. One of the safety officers will assist. Our vehicle and payload has been designed so that most systems can be prepared simultaneously. Our vehicle airframe separates into three main sections, which can be prepared at the same time, and then brought together in a final sequential assembly. Each team member will have a designated system to work on. The team lead and the team mentor (safety officers) will monitor the assemblies and maintain checklists. Although certain team members are assigned to system assemblies, there will always be a secondary team member trained to assist and be completely familiar with each system, in the event that team members are unavailable for launches. The team lead will also be familiar with all system assemblies to assist. For detailed system assembly procedures and checklists, see Appendix. Launch Procedures and Checklists Upon completion of vehicle and payload assembly and integration, the rocket will be presented to the RSO to complete the flight card. The team will be divided at this point, with a minimal number of team members escorting the vehicle to the awaiting launch pad. The team mentor will lead the group and supervise procedures at the pad. The avionics systems team member, the recovery systems team member and the launch systems team member will carry the rocket to the pad and load it on the rail. University of Minnesota USLI CDR 2012 Page 66

67 For details of these procedures, see Appendix. The team mentor will carry a checklist and confirm that all procedures have been followed. All remaining team members will prepare the ground station for flight observation and post flight recovery. The payload systems team will prep the controllers and computers for the RC rover payload experiment following vehicle landing. Following a safe vehicle landing, the payload team will await instruction from the RSO to engage the black powder ejection system and release the rover from the vehicle. The tracking systems team member will test the RF transmitters and receivers to confirm proper function. For detailed procedures and checklists, see Appendix. Launcher Setup The rail will be lowered from vertical to horizontal position. Next the rail lugs on rocket will be lined up with the rail and the rocket will be slid down the rail to the rail stop. Then rocket will be returned to the upright, vertical locked position and altimeters will be activated. Lastly, the launch officer will wire the igniters. Post Flight Procedures Upon successful completion of our flight, with a safe landing of our vehicle and a successful RC rover deployment and experiment, the assigned team members will set off to recover the vehicle and payload from the field, when the RSO deems the field open to recovery. The team lead will supervise the group in recovery. He will be accompanied by the recovery systems team member, and one payload systems team member. There will be a multitude of RF recovery trackers on board the rocket and payload, and the remote video feed will assist in locating the payload. The RF receiver will be handled by the recovery systems team member. Provided there is not systems failures, the vehicle, nose cone and payload should be situated within a short proximity of each other, with all three RF transmitters broadcasting. After locating the vehicle, the first step will be to verify that the black powder charges have ignited and the canisters are empty. If so, the final altitude can be discerned from any of the onboard altimeters audibly prior to powering off all electronics. The vehicle will also be inspected for damaged of missing parts, which must be located in the vicinity prior to returning to staging area. For detailed procedures and checklists, see Appendix. Troubleshooting University of Minnesota USLI CDR 2012 Page 67

68 The team is planning multiple test flights of both half scale and full scale to address all technical difficulties and issues prior to competition flight, as well as to familiarize all team members with all procedures and checklists. In the event there are technical issues, the primary systems team member and the secondary systems team member will address the issue first. If issues persist, the team mentor will assist while the team lead maintains the rest of the team. University of Minnesota USLI CDR 2012 Page 68

69 3.7 Safety and Environment (Vehicle) Safety Officer 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 Failure Modes and Mitigation During design of our rocket vehicle, safety issues were taken into account as they pertained to the rocket vehicle. We are aware that there is inherent risk in designing, building and flying a high powered rocket. We have addressed all failure modes of the rocket while it is in flight, and have developed contingency plans. Due to the failure of our first half scale launch vehicle, we have a new awareness of the failure modes of our vehicle. We have upgraded our failure modes and risk analysis. The following tables summarize all foreseeable failure modes of the vehicle, and include mitigation. Vehicle System: Recovery System Risk Consequence Mitigation Charges fail to ignite Parachutes fail to deploy Proper wiring to avionics is important, will check power to leads prior to ejection canister being loaded, ematches will be centered in the Charges ignite prematurely Charges ignite simultaneously BP residue contaminates avionics Parachutes eject on ascent, possibly causing structural damage Possible structural damage to airframe or parachutes Altimeters cease to operate correctly canister Proper wiring to avionics, and programming of avionics is important Proper wiring and programming The avionics bays must be properly sealed to protect from BP residue, ejection canisters must fit properly and not move or fail University of Minnesota USLI CDR 2012 Page 69

70 Insufficient black powder charge Harness linkage fails Bulkhead attachments fail Shock cords tangle in parachute Shroud lines tangle in parachute Improper parachute selection 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 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 Descent rate too high, too much kinetic energy on landing Hand calculations have been done, and the results have been verified with Rocksim simulation Parachute shreds Descent rate too high We bought one of the strongest Improper length of shock cord Improper shock cord selection Shock cord zippers airframe upon deployment Shock cord snaps chutes The length of the shock cord has been chosen by the rule of thumb; three to four times the length of the rocket The shock cord was chosen based on the result of maximum opening force calculation Table 11: Failure Modes and Risk Mitigation for Recovery System Vehicle System: Avionics System Risk Consequence Mitigation Battery failure Altimeters cease operating in flight, fail to provide enough power to ejection charges Ensure correct orientation of battery terminals, fresh batteries every flight, and holder strength prevents movement under acceleration Altimeter failure Parachutes fail Primary and secondary systems are onboard to ensure ejection charges and electronics will function Incorrect pressure readings Incorrect airbrake and parachute deployment We need to properly size the vent holes for each avionics bay and test it University of Minnesota USLI CDR 2012 Page 70

71 Altimeters programmed incorrectly altitudes Parachutes may fail to deploy during full scale flight tests Use altimeters from different manufacturers to ensure the altimeter programming failures are not duplicated Table 12: Failure Modes and Risk Mitigation for Avionics System Vehicle System: Payload Integration Risk Consequence Mitigation Nosecone separation in flight Low Detailed calculations will be done to confirm proper amount of shear pins Nosecone jams upon deployment Piston jams upon deployment Piston Bulkhead fails Low Low Low in nosecone Using the proper shoulder length, and guarantee that the load to separate is evenly distributed Using the proper piston length for the tube involved, and making sure the inner tube is clean and free of contamination Use proper epoxy techniques to ensure a rigid bond, ground testing to find failure load limit Table 13: Failure Modes and Risk Mitigation for Payload Integration Vehicle System: Airbrake Risk Consequence Mitigation Electrical failure of actuators or relays Airbrakes fail to deploy or deploy at improper altitude As long as altimeter is functioning and programmed properly, and there Mechanical failure of hinge assembly Drag force pulls MMT out of body tube One of airbrake doors jam closed Airbrake fails to open or door separates from vehicle Motor assembly and airbrake assembly free fall to ground Vehicle loses stability in flight and tumbles is electrical continuity to actuators Calculations to prevent cable separation, or hinge failure Use lots of epoxy, and open door slowly Table 14: Failure Modes and Risk Mitigation for Airbrake System University of Minnesota USLI CDR 2012 Page 71

72 Vehicle System: Motor System Risk Consequence Mitigation Body centering rings fail to contain motor assembly Motor CATO s on CR s Motor mount centering rings fail to contain motor assembly Motor retention system fails Motor CATO s Motor casing free falls from altitude Proper epoxy attachment of CR s and calculations of max structural loads Proper epoxy attachment of CR s and calculations of max structural loads on CR s Ensure motor casing is properly retained with a pre-fabricated motor retainer Table 15: Failure Modes and Risk Mitigation for Motor System Vehicle System: Fin System Risk Consequence Mitigation Fin mount fails to stay attached to body tube Fin and fin mount free fall from altitude, rocket loses stability Fin fails to stay in fin mount Fin not designed strong enough Fin free falls from altitude, rocket loses stability Fin flutters causing unstable flight Ensure proper attachment of mount to body tube Ensure fins are properly and snuggly attached to fin mount Table 16: Failure Modes and Risk Mitigation for Fin System Personnel Hazards and Environmental Concerns Personnel hazards will exist during the course of this project, and all steps will be taken to prevent any accidents from occurring. Construction, testing and assembly of our high powered rocket (consisting of various materials including fiberglass, aluminum and wood) will require the use of various specialty tools. Many of the tools required are contained in the Mechanical Engineering machine shop. All team members who will be working on constructing the rocket have completed a shop safety course. University of Minnesota USLI CDR 2012 Page 72

73 Team Artemis has put forth a general rule that will require any member working on any component to do so in pairs. This rule will be held in strict adherence especially when working in the machine shop. The purpose of the rule is to not only prevent accidents by providing assistance in proper shop techniques, but also so that each component that is fabricated will have more than one person who understands the fabrication process. With regard to the various hazardous materials to be used during the testing and construction of our high powered rocket, we will keep all materials in a locked storage cabinet in our Aerospace workspace in Akerman Hall 130B. Upon the purchase of any hazardous material, the team lead and the safety officer will present details of how to handle the material properly during the weekly meetings. All MSDS sheets will be kept in a binder located at the storage cabinet. Again, it will be a mandatory rule that all team members must work in pairs when handling any hazardous material. The shop in Akerman Hall 130B will also contain all safety equipment that will be required for the safe construction of our rocket. It is the responsibility of the team lead and the safety officer to ensure that the first aid kit and the fire extinguishers located in the workspace are functioning properly at all times. Other safety equipment to be purchased will include respirator masks to be used when cutting or sanding fiberglass, applying epoxy and applying paint or primer. Safety goggles will also be purchased to be used as needed. Ear plugs and latex gloves will also be purchased and placed by the storage cabinet to be used as needed. University of Minnesota USLI CDR 2012 Page 73

74 4. Payload Criteria 4.1 Testing and Design of Payload Experiment Payload Systems Figure 36: Three-View rover schematic featuring all major dimensions 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. University of Minnesota USLI CDR 2012 Page 74

75 Power Supply: Voltage Requirements: Battery Capacity: Operational Range: Torque: Transmitter Frequency: SPMB4500NM 6V 4500mAh 2500ft 21 kg-cm total 2.4 GHz Table 17: Control System Specifications The rover will be controlled from the ground station. A pilot will use a DX5e RC transmitter 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. University of Minnesota USLI CDR 2012 Page 75

76 AR600 RC Receiver S6020 Servo DX5e RC Transmitter Figure 37: 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, see Section 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, see Sections and 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 University of Minnesota USLI CDR 2012 Page 76

77 the following figure. This system allows the rover to automatically correct its orientation, provided the rover is in range of the RC transmitter. Figure 38: Automatic Orientation Correction System 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. University of Minnesota USLI CDR 2012 Page 77

78 Electronic Switch RC Receiver Camera Data Logger Servo Figure 39: Rover Drawing With Control Component Locations As of now the only component of the system that has been tested is the BoosterVision BVGM-1 wirelessly transmitting camera. A test in a corn field in North Branch Minnesota verified the camera can transmit over the required range of 2500ft. The battery pack, servos and RC receiver have been tested for functionality, but a range test has yet to be conducted. Upon completion of the rover, the system will be tested, as described in Section 4.1.4, to ensure the rover is capable of surviving all forces experienced during competition, as well as ensuring the rover is capable of working at the required range and for the required duration, as set out in the payload success criterion. 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 in Section 4.1.8). 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. University of Minnesota USLI CDR 2012 Page 78

79 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 18: Data Collection system Specifications 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 40: 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. University of Minnesota USLI CDR 2012 Page 79

80 C. Chassis The rover is 13.25in long and 7.3in wide and is expected to weigh six pounds. In addition to the electrical components described above, the rover will have several mechanical features to improve mission performance. One such feature is the outrigger arm that will be spring loaded for automatic deployment once released from the airframe. This arm will provide stability to the rover by supplying the counter-torque needed for forward movement and is designed to be as simple as possible to increase reliability. In addition to this, the rover drive wheels will compress inwards while loaded into the airframe tube and mate to six bracers on either side of the frame. This configuration will make the rover capable of handling the large loads experienced during the motor burn, parachute deployment and payload ejection by forcing loads around the fragile drive axial and into the strong rover frame. In order to allow for normal driving once deployed, a spring will be fitted into a spacing between the axial cog and the inner face of the wheel to force the wheel to move outward and separate from the bracers. The axial cog is fitted with grooves to allow torque to be transferred to the wheel that match grooved cut into the wheel itself. A model of the rover assembly is shown below: Design Verification A. Terrain Handling Figure 41: Labeled CAD Model of Rover Assembly 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 ribbed extensions of the tire on either wheel help provide traction on soft terrain, as well as providing a tighter fit with the interior of the University of Minnesota USLI CDR 2012 Page 80

81 airframe. This extra grip will aid the rover over the large bumps anticipated in the landing zone. 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 4.4 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. In the PDR it was shown that the steepest slope the rover could climb up, imagine the tilled rows of the field, is about 31º. By equating the force of gravity acting along the incline with the force supplied by the two servos we derived the following expression for required torque, W sinθr τ = 2 where W is the weight of the rover, theta is the angle of incline, and r is the wheel radius. To get an upper limit on the torque requirements we will assume the maximum weight of the payload, 10lbs=44.5N, an angle of 31 degrees and a wheel radius of 3.5 in.=8.89cm.solving for required torque, The S6020 Servos chosen for the rover supply 10.5 kgcm of torque, sufficient to give the rover power to climb over the obstacles presented in the field. B. Orientation Maintenance As discussed in Section 4.1.1, 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, University of Minnesota USLI CDR 2012 Page 81

82 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. Given that the radius of the rover is 3.65 in. and given that most of the weight is evenly distributed about the axis of rotation, the center of gravity should be well within 1.82 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 contacts at Spektrum have indicated the range should be at least 2500ft on the ground. Future ground testing will verify the exact range. The AR600 RC receiver has 2 antennas, one long and one short. By orienting these perpendicular to each other the range of the receiver is maximized. Additionally, in this orientation either one antenna or the other should be able to receive a signal from the RC transmitter, meaning the rover should be capable of receiving command inputs in any orientation. Additionally, by constructing the rover out of G10-fiberglass, the rover should be radio transparent, preventing 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 battery life, far exceeding the requirements of data collection once every minute for a minimum of five minutes. E. Ejection 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 University of Minnesota USLI CDR 2012 Page 82

83 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 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. 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 only rover component that will require a high level of precision and skilled craftsmanship is the cog interface between the drive axial and the main wheel. The team is considering purchasing a cog if one is found to have the correct dimensions, however the matching slot in the wheel will still have to be custom made. This interface must be made correctly or else the wheel may not be capable of rotating, and so the slot must be confirmed to work reliably during testing Testing A. Controls Testing The RC transmitter and receiver will be extensively ground tested in a variety of conditions for ranges extending to 2,640 ft. The primary test will ensure the rover receives and executes command inputs for the entire range, assuming proper orientation. The secondary test will analyze the performance of the rover in improper orientations. Specifically, the range under which the rover is capable of receiving and executing commands while in an improper orientation will be determined. An additional test will be conducted to determine the effects of various obstacles blocking the signal from the transmitter to the receiver. In particular objects that may be encountered during competition, such corn stalks, will be tested. The range testing for the BVGM-1 is complete. The camera signal was received for distances greater that the required 2500ft in simulated competition conditions, i.e. a cornfield in North Branch, Minnesota. University of Minnesota USLI CDR 2012 Page 83

84 B. 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. C. Data Acquisition Testing The HOBO H Data Logger and associated sensors will be tested prior to launch to verify the data acquisition system capable of taking accurate/valid readings. Tests will be performed and compared to expected results, ensuring functionality of the system. D. Ejection Testing The ejection of the payload on the ground is subject to many variables, chief among these is orientation of the payload bay. It is possible the payload bay may land horizontally, at an angle, or with some obstruction in the ejection path. In order to verify the ejection system, the ejection of the rover from the payload bay will be tested in several configurations, including the aforementioned situations. This testing will include the use of the electronic switch and therefore the ejection will be tested at multiple ranges to ensure ejection is possible in the entire 2500ft range Status The rover has not yet moved into the fabrication phase, however the team is nearly ready to order structural components, as well as the last of the electrical components. Range testing is already underway on the wireless camera and RC receiver and the team is becoming more familiar with these systems and their capabilities. All components are to be ordered by the end of January, and fabrication is expected to begin by mid-february and be completed by early March. The team has agreed that the construction of the rover is to be a secondary priority compared to the assembly of the airframe, and although the rover chassis is expected to fly inside the airframe during full-scale testing, priority is to be given to the rocket should manufacturing issues arise to ensure manufacturing stays on schedule Payload Integration As mentioned previously in Section 3.5, the rover is connected to the rocket by only the wire that allows it to fire the black powder charge, once cleared by the RSO. The connection between the rover and rocket must be broken once it is ejected and this is accomplished with a 2 pin molex connector. A short wire will come from electronic switch, discussed in section 4.1.8E, and will connect to a longer e-match which is University of Minnesota USLI CDR 2012 Page 84

85 connected to the black powder canisters. The reason the e-match is longer is because once the rover is ejected, the shorter cord will not interfere with the rover s movement. Figure 42: Molex Connector Manufacturer: SUZO-HAPP Item #: Cost: $.70/each Instrumentation and Measurement A. Precision of Instrumentation The accuracy of the instrumentation is given in section 4.1.8, 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 University of Minnesota USLI CDR 2012 Page 85

86 uncertainty. Given the bias uncertainties reported by the manufacturers, it is expected that total uncertainty will fall in the range of 5-8%. B. Repeatability of Measurement 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 Payload Electronics 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. 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. University of Minnesota USLI CDR 2012 Page 86

87 Figure 43: 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 44: 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 CDR 2012 Page 87

88 Figure 45: 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 19: 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 CDR 2012 Page 88

89 Figure 46: 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 20: Silicon Pyanometer Specifications D. Wireless Camera A BoosterVision BVGM-1 Camera, as specified in the PDR, 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 CDR 2012 Page 89

90 Figure 47: 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 21: BVGM-1 Specifications E. Electronic Switch As mentioned in Section 2.2.1, the ejection of the payload must be triggered after the all clear is given by the RSO. To accommodate this change, 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 AR600 receiver, see Section F. The other end of the switch will connect to a quick release mechanism, discussed in greater detail in Section 4.1.6, 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 University of Minnesota USLI CDR 2012 Page 90

91 payload will be ejected. The switch will be operated at 6V, receiving power from the receiver battery pack. The electronic switch will be tested by performing static ground deployment tests, pending construction of the rover and payload bay. The switch well be triggered remotely and successful deployment of the rover will be verified at various ranges and conditions. Figure 48: Electronic Switch Manufacturer: RCATS Model: RC-100X Power Requirements: 6V (Supplied by Receiver Battery Pack) Cost: $29.95 Table 22: RC-100X Specifications F. 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 CDR 2012 Page 91

92 Figure 49: 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 23: S6020 Specifications G. 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 University of Minnesota USLI CDR 2012 Page 92

93 RC battery pack. Figure 50: 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 51: SPMB4500NM Batter Pack H. 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 CDR 2012 Page 93

94 Figure 52: 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 24 PT-1B Specifications I. RC Receiver Control inputs will be sent to the rover from the ground station using the DX5e RC transmitter. The transmitter will emit signals at 2.4 GHz, to be received by the AR600 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 antennas, measuring 8.25 in. and 2.5 in. in length, which will be mounted perpendicular to each other to maximize the range of the rover. University of Minnesota USLI CDR 2012 Page 94

95 Two S6020 servos, used to drive the wheels as discussed in Section 4.1.1G, 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. Open Port Figure 53: AR600 RC Receiver The AR600 will be powered by a 5 cell NiMH battery pack, the Spektrum SPMB4500NM battery pack discussed in the PDR. 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. Manufacturer: Spektrum Model: AR600 Power Requirements: 6V (Supplied by Receiver Battery Pack) Number of Channels: 6 Band: 2.4 GHz Range: Full Range line of sight, 2640 ft. on ground Antenna Length: 8.25 in., 2.5 in. Dimensions: 1.18x0.85x0.49 in. Weight: lbs. Cost: Included with RC Transmitter ($100) Table 25: Spektrum SPMB4500NM Specifications University of Minnesota USLI CDR 2012 Page 95

96 S6020 Servo S6020 Servo AR600 RC - Receiver SPMB4500NM Battery Pack RC-100X Electronic Switch Ejection Charge Quick Release Figure 54: Electrical Schematic of Payload Control System 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. University of Minnesota USLI CDR 2012 Page 96

97 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 Safety Analysis 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 University of Minnesota USLI CDR 2012 Page 97

98 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. 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 Section Payload Concept Features and Definition The purpose of the rover is to explore the potential value in sending an inexpensive probe to potential landing zones on extraterrestrial bodies in order to gain detailed scouting information. The information gained by the scouting rover may be used to mitigate risks for future, large scale operations. 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. 4.3 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. University of Minnesota USLI CDR 2012 Page 98

99 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. 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 Experimental 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 University of Minnesota USLI CDR 2012 Page 99

100 experimental sense, but will instead serve as validation that the flight as deployment left 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, Variables 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 4.1.8, which includes the uncertainty error associated with each instrument. Additionally, a more detailed explanation of the error analysis appears in section The variables in the experiment will obviously 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 pre-launch 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 4.1.8, and the expected bias uncertainty of the measurements will be determined during testing, which is detailed in section As referenced in section 4.1.7, 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 the validity of collected mission data when after testing has been completed Preliminary Experiment Process Procedures University of Minnesota USLI CDR 2012 Page 100

101 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. 4.4 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 Consequence Mitigation Status Rover out of range of RC Controller Rover lands in unfavorable orientation. Rover is damaged during take off, flight, and or landing. Rover fails to receiver control inputs and is uncontrollable. Rover is unable to receive control inputs to correct orientation. Rover may be incapable of driving. Extensive ground testing will determine exact range for which rover will receive controls. Ground Stations and Control System will be improved if need be to boost signal. Development of passive/automatic techniques to correct orientation of rover without control inputs. Configure the wheels to move inwards during flight and meet sturdy bracers connected to Proposed In progress Complete University of Minnesota USLI CDR 2012 Page 101

102 the frame Rover electronics are damaged during take off, flight or landing. Unable to receive camera signal. Rover is unable to navigate over difficult terrain. Rover sustains damage during black powder ejection Rover ejection takes place before bystanders are clear Rover may be incapable of receiving control inputs or taking data. Impossible to navigate, driving blind. 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. Electronics will be protected by the chassis and rover will be wrapped in protective coating during flight Extensive ground testing will determine the exact range for which signal may be received. Antenna may be upgraded if needed to increase range Use of high torque servos to guarantee rover has sufficient power. The development of expanding wheels to give the rover greater ground clearance. Extensive static testing of rover deployment will ensure all rover systems can survive deployment prior to launch. An electronic switch has been implemented to prevent triggering of charges until all clear signal is given. Complete Complete In progress Proposed Complete Table 26: Summary of Payload Failure Modes 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 University of Minnesota USLI CDR 2012 Page 102

103 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. As stated in section 3.5.3, 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. Thorough testing of the deployment system (detailed in section 4.1.8D) will be necessary to determine the exact safety procedure involved with the finalized deployment procedure prior to launch. Additionally, 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 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. A potential environmental concern associated with the payload is the proposed spacer put in between the rover and the inner wall of the airframe in order to ensure a tight fit and a successful deployment. The team has not yet determined the details of this system and as such do not have a material selection nor a definite plan to ensure that the material will not be discarded into the field, however the easiest and most likely solution will simply be to tie the ends of the spacers to the airframe so they will not blow away and become unrecoverable. University of Minnesota USLI CDR 2012 Page 103

104 5. Activity Plan 5.1 Budget Our budget has changed since the preliminary design report. The following tables outline details of our project budget and include status, to indicate what has been completed to date. Our funding comes from three sources. We are planning on staying below our total allocated funds as much as possible. 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 Ammounts. 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 University of Minnesota USLI CDR 2012 Page 104

105 Half Scale System Component Unit Cost Qty Total Cost Status Nosecone Nosecone $ $ Purchased Nosecone Bulkhead $ $ 5.70 Purchased Tracking RF Tracker 1 $ - On Hand Payload Section Airframe Tube $ $ Purchased Piston Tube $ $ Purchased Piston Bulkhead $ $ 5.70 Purchased Ejection Canisters $ $ 5.50 Purchased Eyebolts $ $ 4.00 Purchased Separator Bulkhead $ $ 0.01 Custom Avionics Avionics Sleeve $ $ 7.43 Purchased Avionics Sled $ $ 1.07 Purchased Primary Altimeter $ $ On Hand Secondary Altimeter 1 $ - On Hand Primary Battery $ $ 5.00 On Hand Secondary Battery $ $ 5.00 On Hand Switches 2 $ - On Hand Threaded Rods $ $ 0.02 On Hand Inner Avionics Bulkhead $ $ Purchased Outer Avionics Bulkhead $ $ 5.70 Purchased Eyebolt $ $ 2.00 Purchased Upper Airframe Tube $ $ Purchased Booster Transition $ $ Purchased Main Recovery Piston Tube $ $ 7.43 Purchased Piston Bulkhead $ $ 5.70 Purchased Ejection Canisters $ $ 5.00 Purchased Eyebolts $ $ 4.00 Purchased Main Parachute $ $ On Hand Main Shock Cord $ $ 7.50 On Hand Coupler Outer Band $ $ 1.57 Purchased Coupler Tube $ $ Purchased Bulkhead $ $ 5.70 Purchased U-bolts $ $ 4.00 Purchased Drogue Recovery Drogue Parachute $ $ Purchased Drogue Protector $ $ On Hand Drogue Shock Cord $ $ 4.14 On Hand University of Minnesota USLI CDR 2012 Page 105

106 Cord Protector $ $ On Hand Lower Airframe Tube $ $ Purchased Booster Body Centering Ring 1 $ $ 0.01 Custom Body Centering Ring 2 $ $ 0.01 Custom Avionics Ejection Canister $ $ 5.00 Purchased Avionics Bulkhead $ $ 5.70 Purchased U-bolt $ $ 2.00 Purchased Avionics Sled $ $ 1.07 Purchased Primary Altimeter $ $ On Hand Secondary Altimeter 1 $ - On Hand Batteries $ $ On Hand Threaded Rods $ $ 0.02 On Hand Rear Av Bay Bulkhead $ $ 0.01 Custom Fins Surface Fin Mount 3 $ - Custom Fins $ $ 5.70 Purchased Motor Motor Mount Tube $ $ 5.10 Purchased CTI Pro54 3G Casing $ $ Purchased CTI Pro54 J1520 Reload $ $ Purchased MMT Centering Ring 1 $ $ 0.01 Custom MMT Centering Ring 2 $ $ 0.01 Custom Airbrake Door Section $ $ 4.70 Purchased Airbrake Centering Ring $ $ 0.01 Custom Motor Retainer $ $ Purchased Half Scale Subtotal $ Table 29: Half Scale Expenditures The costs for the scale test vehicle will increase by an unknown amount. We will salvage and retest all usable components from the first scale test vehicle, and purchase all damaged parts. We plan on rebuilding the half scale as soon as possible. University of Minnesota USLI CDR 2012 Page 106

107 Full Scale System Component Unit Cost Qty Total Cost ProRated Nosecone Nosecone $ $ $ Bulkhead $ $7.99 $7.99 Tracker RF Tracker $ $49.95 $49.95 Payload Section Airframe Tube $ $ $ Piston Tube $ $58.00 $9.67 Piston Bulkhead $ $6.95 $6.95 U-Bolts $ $7.54 $7.54 Piston Centering Ring $ $7.99 $7.99 Payload Bulkhead 1 $1.00 $1.00 Avionics Avionics Sleeve $ $42.50 $3.54 Avionics Sled 1 $2.00 $2.00 Primary Altimeter $ $ $ Secondary Altimeter $ $85.00 $85.00 Batteries $ $10.00 $10.00 Threaded Rods $ $2.29 $2.29 Inner Av Bulkhead $ $8.55 $8.55 Outer Av Bulkhead $ $8.55 $8.55 U-bolt $ $3.78 $3.78 Quicklinks $ $6.94 $6.94 Magnetic Switches $ $50.00 $50.00 Upper Booster Main Recovery Coupler Section Airframe Tube $ $39.50 $24.69 Transition CR 1 $1.00 $1.00 Body Transition $ $ $ Threaded Rods $ $2.30 $2.30 Piston Tube $ $42.50 $5.31 Piston Bulkhead $ $8.55 $8.55 U-bolts $ $7.56 $7.56 Quicklinks $ $10.41 $10.41 Main Parachute $ $ $ Main Shockcord $ $16.40 $16.40 Deployment Bag $ $34.00 $34.00 Outer Band 1 $1.00 $1.00 Coupler Tube $ $42.50 $11.51 U-bolts $ $7.56 $7.56 Quicklinks $ $10.41 $10.41 Bulkhead $ $6.56 $6.56 University of Minnesota USLI CDR 2012 Page 107

108 Drogue Recovery Lower Booster Drogue Parachute $ $80.00 $80.00 Drogue Shockcord $ $16.40 $16.40 Deployment Bag $ $25.00 $25.00 Airframe Tube $ $39.50 $18.10 Body CR 1 1 $1.00 $1.00 Body CR 2 1 $1.00 $1.00 RF Tracker $ $49.95 $49.95 Avionics Front Bulkhead $ $6.56 $6.56 U-bolt $ $3.78 $3.78 Quicklinks $ $6.94 $6.94 Avionics Inner Tube $ $18.50 $2.06 Primary Altimeter $ $ $ Secondary Altimeter $ $ $ Batteries $ $10.00 $10.00 Nylon Battery Holders $ $8.00 $8.00 Threaded Rods $ $2.30 $2.30 Magnetic Switches $ $50.00 $50.00 Rear Bulkhead 1 $1.00 $1.00 Motor MMT CR 1 1 $1.00 $1.00 MMT CR 2 1 $1.00 $1.00 Motor Mount Tube $ $16.50 $10.08 Motor Casing $ $ $ Motor Reload 4-G $ $ $ Motor Retainer $ $52.00 $52.00 Fins Surface Fin Mounts 3 $10.00 $10.00 Fins 3 $10.00 $10.00 Airbrake Hinge Assembly $ $40.00 $40.00 Airbrake Doors $ $21.00 $21.00 Airbrake CR 1 $1.00 $1.00 Airbrake Actuators $ $ $ Airbrake Relays $ $5.12 $5.12 Boat Tail Boat Tail $ $ $ Boat Tail CR 1 $1.00 $1.00 Full Scale Subtotal $ $ Table 30: Full Scale Expenditures University of Minnesota USLI CDR 2012 Page 108

109 All full scale parts are in the ordering process. Note the prorated column is our on the pad costs. Also, some parts are on hand from previous rocketry teams and the unit costs are not included. Payload System Component Unit Cost Qty Total Cost Status Chassis Chassis panel $ $60.00 Proposed Springs $ $15.00 Proposed Wheels $ $11.00 Proposed Outrigger $ $5.00 Proposed Bushings $ $14.00 Proposed Electronics Microstation $ $ Proposed Temp/RH Probe $ $ Proposed Pyranometer $ $ Proposed GearCam Wireless Cam $ $ Purchased Battery $ $20.00 Purchased RC Transmitter $ $ Purchased Wireless Reciever $0 1 $0.00 Purchased Battery Pack $ $73.00 Purchased Charger $ $30.00 Purchased Charger Adapter $ $3.00 Proposed Ground Servo $ $90.00 Purchased Recovery Transmitter $ $50.00 Proposed Electronic Switch $ $30.00 Proposed Payload Subtotal $1,492 Table 31: Payload Expenditures Note: The full cost of the payload can be added to the vehicle prorated cost for a final on the pad cost. Vehicle 2708 Payload 1492 Total On The Pad Cost 4200 This value satisfies the USLI requirement and allows leeway for design modifications of needed pending further testing. University of Minnesota USLI CDR 2012 Page 109

110 Testing and Supplies System Item Unit Cost Qty. Total Cost Status Launch Rail (per inch) $ $21 Proposed Tower Legs $42 4 $84 Proposed Base Post $20 1 $20 Proposed Blast Plate $40 1 $40 Proposed Leg Screw $48 1 $48 Proposed Leg Pin $20 4 $80 Proposed Leg Foot $20 4 $80 Proposed Hardware $20 1 $20 Proposed Safety Supplies Foam Ear Plugs $ $7.63 Purchased Latex Gloves (per box) $ $15.03 Purchased Safety Glasses $ $18.96 Purchased Respirator Masks (10-pack) $ $40.17 Purchased Portable First Aid Kit $ $50.00 Proposed Tools Dremel Rotary Tool $ $81.96 Purchased Calipers $ $21.11 Purchased Portable Weigh Scale $ $60.00 Proposed Multi-meter $ $50.00 Proposed 150 lb Strain Gauge $ $25.00 Proposed 50 lb Strain Gauge $ $20.00 Proposed Finishing Epoxy Resin $ $20.00 On Hand Epoxy Hardener $ $40.00 On Hand Epoxy Filler $ $20.00 On Hand Various Spray paint $ $ Proposed Cleaning Cleaning Wipes $ $30.00 Purchased Fastening Shear Pins (per 100 pack) $ $6.13 Purchased Shear Pins (per 100 pack) $ $7.70 Purchased Plastic Rivets $ $16.00 Purchased Other PVC Rocket Stand $ $24.11 Purchased Wooden Travel Crate $ $ Proposed Polymer Film $ $20.37 Purchased Full Scale Flight Test Motor $ $ Proposed University of Minnesota USLI CDR 2012 Page 110

111 Rocksim $ $ Purchased Supplies/Testing Subtotal $1,992 Table 32: Testing and Supplies Cost Travel Item Per Day Days Mileage Charge Total Miles Total Van Rental $51 5 $ $715 Truck Rental $47 5 $ $975 Per Room Rooms Number of Nights Total Hotel $ $696 Travel Subtotal $2386 Table 33: Travel Expenditures Currently the travel costs are estimates and we are in the process of booking our vehicle rentals and hotels. More detailed budget spreadsheets can be found in the Appendix listing the vendors and manufacturers of our parts. 5.2 Timeline The project timeline has changed since the preliminary design report. The school break over the holidays prevented effective team communication as many team members were no longer in the city. With the beginning of the spring school semester, we have re-evaluated our project schedule. It has come to our attention that to be successful in completing our project we must maintain a strict schedule. University of Minnesota USLI CDR 2012 Page 111

112 Given that the whereabouts of our scale vehicle is unknown, and we will require extra time to locate and/or rebuild a scale vehicle this also puts our project in jeopardy of failure to meet deadlines. The team is willing and motivated to put in extra hours to get the project back on schedule. We are planning on proceeding with full scale build and whatever testing is possible until the half scale is found and analyzed. We have also begun to rebuild components for a second half scale vehicle, which will be manufactured and assembled faster than the previous one given that there will be no design changes. The key factor hampering our build timeline is parts acquisition. Our new timelines are based on reaching the milestones set forth by USLI. The following tables outline key dates of our revised timelines. Detailed Gantt charts can be found in the Appendix integrating all process timelines USLI Report Timeline January 16 th, 2012 January 18 th, 2012 January 22 nd, 2012 January 23 rd, 2012 February 1 st 10 th, 2012 March 19 th, 2012 March 21 st, 2012 March 25 th, 2012 March 26 th, 2012 April 2 nd 11 th, 2012 April 30 th, 2012 May 7 th, 2012 Draft CDR Report Complete Draft CDR Presentation Complete CDR Posted CDR Due CDR Presentations Draft FRR Complete Draft FRR Presentation Complete FRR Posted FRR Due FRR Presentations Draft PLAR Complete PLAR Due Table 34: Reports and Design Timeline Summary Test Launch Timeline Date Event Status January 18 th, 2012 Half Scale Test Flight Complete February 9 th, 2012 Secondary Half Scale Test Flight Pending March 3 th, 2012 Full Scale Flight Test Pending March 10 th, 2012 Final Day for Full Scale Test Launch Pending University of Minnesota USLI CDR 2012 Page 112

113 Table 35: Table Launch Schedule Our first test flight has been completed, although the recovery was unsuccessful. We are planning another test flight with a second test vehicle. The dates for the full scale launch have been pushed back from tentative dates, and currently there is little margin for an unsuccessful full scale test flight. Manufacturing, Acquisitions and Assembly Timeline Date Event Status December 5 st 22 nd, 2011 Parts Acquisition (Scale vehicle Completed components, payload electronics) January 3 th - 7 nd, 2012 Machining and Parts Preparation Completed (Scale vehicle components) January 9 th 13 th, 2012 Scale vehicle assembly Completed January 14 th 16 th, 2012 Scale vehicle ground testing Completed January 17 rd 23 rd, 2012 Full Scale Parts Acquisition In Progress January 23 th 27 nd, 2012 Parts Acquisition (Scale vehicle Pending replacement) January 23 rd - 27 th, 2012 Fabricating Test Stands Pending January 28 th February Full Scale Parts Machining and Pending 20 th, 2012 Preparation February 20 th 26 th, 2012 Full Scale Systems Tests Pending February 27 th March 3 rd, 2012 Full Scale Vehicle Assembly and Integration Pending Table 36: Manufacturing and Assembly Timelines The manufacturing and assembly timeline is rigid, and depends heavily on acquisition lead times, which will be a major factor in keeping our project on time. Scale Vehicle Timeline November 7 th - 21 th, 2011 November 21 st, 2011 December 8 th 22 nd, 2011 Half Scale Design Parts Order Machine Work for build University of Minnesota USLI CDR 2012 Page 113

114 January 3 rd 7 th, 2012 January 7 th, 2012 Half Scale Assembly Half Scale Launch Table 37: Half Scale Timeline Summary Outreach Timeline October 17 th, 2011 November 19 th, 2011 December 8th, 2011 January 7th, 2012 Jan-April 2012 Jan-April 2012 March 22 nd, 2012 Urban 4-H Kickoff Family Fun Fair Farnsworth Schoool Science Fair (tentative) Davinci Fest (tentative) Teaching Smart (tentative) Farnsworth Aerospace School (tentative) Lincoln Center Elementary Science Night (tentative) Table 38: Educational Engagement Timeline Summary 5.3 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. Hands on activities for our event either given to us or made include, but are not limited to: - Straw rockets - Plastic cup air cannons - CD Hovercraft - Water hydraulic pet racers - Air pneumatic circuit kit - Water hydraulic excavator demonstrator - 1 foot tall rubber based, air propelled rocket - 4 inch water propelled plastic rockets At this time, we have performed 2 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 the CD hovercraft and straw rockets University of Minnesota USLI CDR 2012 Page 114

115 to emulate fluid power and fin design of rockets. In the second half of the event we went outside to shoot off the water propelled rockets as well as the larger 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. The next upcoming event is on Friday, January 27 th. It is at Marcy Open, a K-8 school located near campus. The event is through Teaching SMART (Science, Math and Research Technology), which is a volunteer-based student group which volunteers in local elementary and middle schools near campus. Each SMART event is an hour long University of Minnesota USLI CDR 2012 Page 115