Table of Content 1) General Information ) Summary of PDR Report ) Changes Made Since Proposal ) Safety... 8

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2 Table of Content 1) General Information Student Leader Safety Officer Team Structure NAR/TRA Sections ) Summary of PDR Report Team Summary Launch Vehicle Summary Payload Summary ) Changes Made Since Proposal Changes to Launch Vehicle Changes to Payload Criteria ) Safety Safety Officer Duties Use of Materials NAR Safety Code Compliance Communications of Safety Plan Risk of Assessment Compliance with Federal, State, and Local Laws Handling Rocket Motor Range Safety Regulation Safety Checklist ) Launch Vehicle Criteria Mission Success Criteria Selection and Rational of Launch Vehicle Launch Vehicle Design Motor Selection and Alternatives Recovery & Avionics Subsystem Mission Performance Prediction Launch Vehicle Team Derived Requirements ) Payload Criteria Rover Deploying Mechanism (RDM) Overview System Summary Control Electronics Preliminary Interfaces Testing Plan RDM Team Derived Requirements ) Payload Criteria Autonomous Rover Overview Rover Design Overall Rover Design

3 7.4 Wheel Design Material Motor Sensor Testing Plan Control Rover Team Derived Requirements ) Scientific Research Airbrake Overview Airbrake Deployment Design Material Control Applicable Equations Electronics Mechanical Design Airbrake Team Derived Requirements ) Project Plan Requirement Verifications Timeline Budget Funding Educational Engagement ) Appendices Acronym

4 Section 1: General Information 1.1 Student Team Leader The Team Leader is Nam Nguyen (California State University, Long Beach, class of 2019). The academic year is Nam s third year on LBR. Table 1: Student Leader Information Name Title of Long Beach Rocketry Contact Nam Nguyen Team Leader nguyentransinam1717@yahoo.com (714) Safety Officer The team Safety Officer is Shawn Everts (California State University, Long Beach, class of 2018). The academic year is Shawn s first year on LBR, and he has gone through all the mandatory safety trainings during the summer for engineering teams. Table 2: Safety Officer Information Name Title of Long Beach Rocketry Contact Shawn Everts Safety Officer shawneverts@gmail.com (951) Team Structure The Long Beach Rocketry Team will consist of approximately 20 students from a variety of backgrounds. The team consists of students from the aerospace engineering, mechanical engineering, electrical engineering, and computer science departments. 3

5 1.4 NAR/TRA Sections Figure 1: NASA Student Launch team structure The team will work with the NAR/TRA sections listed in Table 3 for purposes of mentoring review of designs and documentation, and/or launch assistance. Table 3: NAR/TRA sections Section NAR Number TRA Number Launch Field Location Rocketry Organization of California (ROC) Lucerne Valley, CA Friends of Amateur Rocketry, Inc. (FAR) Not NAR/TRA sponsored Mojave Desert, CA 4

6 Section 2: Summary of PDR report 2.1 Team Summary School Name: Organization: Mailing Address: California State University, Long Beach Long Beach Rocketry 1250 Bellflower Blvd. Long Beach, CA LBR will have a Team Advisor and a Team Mentor: Team Advisor Dr. Praveen Shankar Associate Professor Department of Mechanical and AeroSpace Engineering praveen.shankar@csulb.edu Team Mentor David Alexander Roy Visual Artist, High Power Rocketry enthusiast Otis College of Art and Design kj6zfv@gmail.com Launch Vehicle Summary The launch vehicle will be 6 inches diameter and 108 inches in length. The launch vehicle is planned to weight 42 lb. with 13% extra mass contingency. The launch vehicle will have a 75- mm diameter motor mount tube and launch on a Cesaroni L1350 motor. LBR s recovery system involves three separate sections landing with the drogue and the main parachutes. The Milestone Flysheet can be found on LBR s website at longbeachrocketry.com/documents. 2.3 Payload Summary The payload of the Long Beach Rocketry launch vehicle will consist of the deployable rover design experiment selected from the NASA Handbook. The team will design and build an autonomous rover capable of deploying solar panels after moving at least 5 feet, as well as a rover deployment mechanism which will deploy the rover after the team remotely activates the process. Several designs have been considered but have been narrowed down through trade studies to ensure the best outcome of the mission. 5

7 Section 3: Changes Made Since Proposal 3.1 Changes to Launch Vehicle The major changes in the launch vehicle are the result of adding the airbrake subsystem to the launch vehicle. With the addition of the airbrake, the motor has been changed from AeroTech L1390G to a Cesaroni L1350 CS-P because the predicted overall weight of the launch vehicle has increased since the proposal. The propulsion bay also increased by 8 inches to create space for the airbrake subsystem. The Recovery and Avionics section has undergone further analysis on proper materials and design to facilitate a successful recovery sequence. A few changes have been made since the Proposal submitted. These changes are necessary and will increase the probability for success throughout the mission. The first major change are the main and drogue parachute sizes. Initially, the drogue was meant to be the 24-in Fruity Chutes Classic Elliptical Parachute and the main was meant to be the 72-in Fruity Chutes Iris Ultra Parachute. These given parachutes were based off the previous intended weight of 32-lbs. Because of the new weight increase, a larger diameter main chute is required to compensate for the new kinetic energy upon landing that would have been declared unsafe by NASA Student Launch regulations if still using the previous main chute size; however, because of the larger diameter, the launch vehicle is inclined to drift more if the same drogue was used. Therefore, to compensate for the potential drift increase falling out of the required recovery radius, a smaller drogue is now being used. The new drogue chute will be the 20" Fruity Chutes TARC Low and Mid Power Parachute and the new main chute will be the 84" Fruity Chutes Iris Ultra Standard Parachute. As described, the new parachutes will provide enough descent velocity to be safe, drift under the recovery radius, and provide a kinetic energy upon landing that would be within NASA criteria. The secondary major change is the GPS placement. Initially, the GPS was intended to be located beneath the altimeter electronics for not having the GPS frequencies interfere with the altimeter signals for the electronic matches. Prior experience and through clearance of other professionals in the industry, the previous GPS location has yet to cause any dangerous or unsafe conditions before, during, and after launch; however, due to the possible cause of the GPS having an adverse effect on the altimeters being used in the launch vehicle, the GPS location has been moved and is now part of the Nose Cone s internal structure. Placing the GPS in this location increases the success rate of the launch vehicle as it eliminates the close distance between the GPS and altimeters, creating a barrier that shields the altimeters from the GPS transmitting frequencies. 6

8 3.2 Changes to Payload Criteria Since the proposal submission, LBR has considered several other routes in completing the mission which has resulted in design changes and further development for both the rover and the rover deployment mechanism. The rover has originally been designed as a triangular shape to use payload space efficiently and allow for deployment in any direction. After exploring additional options, the team has decided to settle on this design and proceeded with detailed development. A gearbox has been created which allows for the control of all three wheels per side using a single motor which gives additional space for electronics to create a more complex yet robust control system. Rubber tires have been chosen through a trade study due to their additional grip and reduced weight compared to the original 3D printed design. These slight changes have addressed the challenges that arose during the proposal which will increase the likelihood of mission success for the rover. In addition to the rover, LBR reconsidered the original overly complex rover deployment mechanism and found a much simpler solution to the problem which reduces complexity and weight. The original design of the RDM was to have three threaded rods coupled with motors attached to a bulkhead, which would spin and unscrew threaded nuts on an opposing bulkhead. The rover would sit in the middle of the RDM and would have a shaft running through the center of it, giving a rover a path to follow during deployment which would ensure the rover does not return into the payload bay once deployed. Though this design is feasible, due to weight and complexity the team has decided to opt for a different solution. The new approach is similar, but now instead of four total rods there will simply be one threaded rod in the center responsible of removing the nosecone and deploy the rover. Through testing, the team has concluded this system is sufficient for the task and is a fraction of the weight. 7

9 Section 4: Safety 4.1 Safety Officer Duties Shawn Everts is the official Safety Officer of the Long Beach Rocketry team. The Safety Officer will be responsible for ensuring the safety of persons participating in the NASA University Student Launch Initiative activities. The Safety Officer will assure that the team adheres to all regulations pertaining to the construction, assembly, testing, flight, and recovery phases of the launch vehicle. Required Training for Safety Officer The Safety Officer will receive training to gain a thorough understanding of the National Association of Rocketry (NAR) safety code, the Tripoli Rocketry Association (TRA) safety code, and the Federal Aviation Administration's (FAA) code pertaining to High Power Rocketry. The Safety Officer will obtain MSDS sheets for all chemicals used and will keep copies of them for reference for all team members. Safety Officer Responsibilities The Safety Officer will be responsible for ensuring all the following: Ensure that all team members understand and comply with the NAR high power safety code. Ensure that all team members will abide by the rules set forth by the FAA. Develop and maintain team's hazard analyses, failure mode analysis and MSDS/chemical inventory data. Prepare a team safety plan and communicate it to all team members. Develop risk assessment tables to determine a risk level for each threat and a proposed mitigation. Ensure proper PPE is supplied during construction of rocket 4.2 Use of Materials The Safety Officer will also organize and maintain MSDS records of all hazardous materials used during construction, assembly, testing, launch, operation, and recovery phases of the launch vehicle. All team members will be informed of the risks of the materials, concerning potential health and safety hazards. Briefings will be given by the Safety Officer for proper handling of all materials. Composite Materials During the construction of the launch vehicle will be using composite materials including fiberglass and carbon fiber. Cutting fiberglass or carbon fiber generates particles that can irritate the eyes, skin and respiratory system. To ensure safety of all team member s PPE will be required including masks, safety glasses, and gloves. 8

10 List of Chemicals Pyrodex P Smokeless Powder o MSDS - Acetone o MSDS - SYS=1&C001=MSDS&C997=C100%3BESDS_US%2BC102%3BUS%2B1000 &C100=*&C101=*&C102=*&C005= &C008=&C006=HON&C0 13 Epoxy Resin o MSDS - Epoxy Hardener o MSDS - Hot Glue o MSDS NAR Safety Code Compliance NAR Code Table 3: NAR Safety Code Compliance Certification. I will only fly high power rockets or possess high power rocket motors that are within the scope of my user certification and required licensing. Materials. I will use only lightweight materials such as paper, wood, rubber, plastic, fiberglass, or when necessary ductile metal, for the construction of my rocket. Motors. I will use only certified, commercially made rocket motors, and will not tamper with these motors or use them for any purposes except those recommended by the manufacturer. I will not allow smoking, open flames, nor heat sources within 25 feet of these motors. Compliance David, NAR/TRA personnel, has a level 2 certification from the TRA and he will solely handle rocket motors. Structures lead will be responsible for using the proper material to comply with the NAR safety codes. Motor will be purchased and shipped to Huntsville before Launch week. NAR/TAR personnel will be responsible for storing and handling the motor before, during and after launch week. 9

11 Ignition System. I will launch my rockets with an electrical launch system, and with electrical motor igniters that are installed in the motor only after my rocket is at the launch pad or in a designated prepping area. My launch system will have a safety interlock that is in series with the launch switch that is not installed until my rocket is ready for launch, and will use a launch switch that returns to the off position when released. The function of onboard energetics and firing circuits will be inhibited except when my rocket is in the launching position. Misfires. If my rocket does not launch when I press the button of my electrical launch system, I will remove the launcher s safety interlock or disconnect its battery, and will wait 60 seconds after the last launch attempt before allowing anyone to approach the rocket. Launch Safety. I will use a 5-second countdown before launch. I will ensure that a means is available to warn participants and spectators in the event of a problem. I will ensure that no person is closer to the launch pad than allowed by the accompanying Minimum Distance Table. When arming onboard energetics and firing circuits I will ensure that no person is at the pad except safety personnel and those required for arming and disarming operations. I will check the stability of my rocket before flight and will not fly it if it cannot be determined to be stable. When conducting a simultaneous launch of more than one high power rocket I will observe the additional requirements of NFPA This requirement will be followed. It is the responsibility of the Safety Officer and the team leads to ensure that the ignition system is in compliance with the NAR safety codes. The Safety Officer will be responsible for ensuring that this requirement is met. During a misfire the Range Safety Officer will have the final say. The launch safety requirement will be follow. The Safety Officer will ensure the minimum distance table is enforced. 10

12 Launcher. I will launch my rocket from a stable device that provides rigid guidance until the rocket has attained a speed that ensures a stable flight, and that is pointed to within 20 degrees of vertical. If the wind speed exceeds 5 miles per hour I will use a launcher length that permits the rocket to attain a safe velocity before separation from the launcher. I will use a blast deflector to prevent the motor s exhaust from hitting the ground. I will ensure that dry grass is cleared around each launch pad in accordance with the accompanying Minimum Distance table, and will increase this distance by a factor of 1.5 and clear that area of all combustible material if the rocket motor being launched uses titanium sponge in the propellant. The team will comply with this NAR code. At the launch field the Range Safety Officer will determine if it is safe to launch. Size. My rocket will not contain any combination of motors that total more than 40,960 N-sec (9208 pound-seconds) of total impulse. My rocket will not weigh more at liftoff than one-third of the certified average thrust of the high-power rocket motor(s) intended to be ignited at launch. The team will follow this requirement. Leads will be responsible for design the rocket with the constraint. Flight Safety. I will not launch my rocket at targets, into clouds, near airplanes, nor on trajectories that take it directly over the heads of spectators or beyond the boundaries of the launch site, and will not put any flammable or explosive payload in my rocket. I will not launch my rockets if wind speeds exceed 20 miles per hour. I will comply with Federal Aviation Administration airspace regulations when flying, and will ensure that my rocket will not exceed any applicable altitude limit in effect at that launch site. The Safety Officer and NAR/TRA personnel will ensure that this requirement is followed. 11

13 Launch Site. I will launch my rocket outdoors, in an open area where trees, power lines, occupied buildings, and persons not involved in the launch do not present a hazard, and that is at least as large on its smallest dimension as one-half of the maximum altitude to which rockets are allowed to be flown at that site or 1500 feet, whichever is greater, or 1000 feet for rockets with a combined total impulse of less than 160 N-sec, a total liftoff weight of less than 1500 grams, and a maximum expected altitude of less than 610 meters (2000 feet). The test site is NAR/TRA certified and the team will comply with the what the Range Safety Officer says. Launcher Location. My launcher will be 1500 feet from any occupied building or from any public highway on which traffic flow exceeds 10 vehicles per hour, not including traffic flow related to the launch. It will also be no closer than the appropriate Minimum Personnel Distance from the accompanying table from any boundary of the launch site. The launch field is set up with the appropriate distance from any buildings and highway. The team will comply with the minimum distance table and follow the instructions from the Range Safety Officer. Recovery System. I will use a recovery system such as a parachute in my rocket so that all parts of my rocket return safely and undamaged and can be flown again, and I will use only flame-resistant or fireproof recovery system wadding in my rocket. The Recovery lead and the Safety Officer will make sure that the design adheres to this requirement. Recovery Safety. I will not attempt to recover my rocket from power lines, tall trees, or other dangerous places, fly it under conditions where it is likely to recover in spectator areas or outside the launch site, nor attempt to catch it as it approaches the ground. The team will follow this requirement during every launch. Minimum Distance Table Installed Total Impulse (Newton- Seconds) Equivalent High- Power Motor Type Table 4: Minimum Distance Minimum Diameter of Cleared Area (ft.) Minimum Personnel Distance (ft.) Minimum Personnel Distance (Complex Rocket) (ft.) 12

14 H or smaller I , , , , , , , , , , , J K L M N O Communication of Safety Plan The Safety Officer, will be responsible for communication of the safety plan to every team member and will ensure that all team members follow the safety protocols at all times. Hazard Recognition and Accident Avoidance All members of the Long Beach Rocketry team will be given a safety briefing and are required to sign a safety contract. The team's Safety Officer will be responsible for holding team safety briefing before any construction to ensure that all team members are briefed on hazard recognition and accident avoidance. The briefings will cover: Procedures of construction Proper use of personal protective equipment Safety of materials used for that session Risks associated with project Laboratory safety Tool safety Manufacturing procedures If an accident occurs, the Safety Officer will be notified immediately, and efforts will be taken to mitigate the risk. After the situation has occurred the safety officer will meet to discuss the incident with the all the parties involved and discuss ways to prevent a second occurrence. Pre-Launch Safety Plan The Safety Officer and Team Mentor will give a safety briefing before each launch. The briefing will cover launch procedures, NAR high power rocket safety code compliance, and the launch 13

15 site rules. The briefing will remind all members that the Range Safety Officer has the final say regarding safety and flight readiness of the rocket. The goal of the rocket launch, and the expected outcome will also be discussed at the briefings. Compliance with Safety Plan Every member of the Long Beach Rocketry team must sign and abide by the Long Beach Rocketry Team Safety Agreement that is shown on the next page: Long Beach Rocketry Team Safety Agreement In signing this contract, I agree to follow the rules and safety procedures that have been outlined in this document. I agree to read and agree to follow the laws and regulations: Federal Aviation Regulations 14 CFR, Subchapter F, Part 101, Subpart C; Amateur Rockets Code of Federal Regulation 27 Part 55: Commerce in Explosives; and fire prevention NFPA 1127 Code for High Power Rocket Motors. I also agree to: Listen to instruction given by Long Beach Rocketry Safety Officer or NAR/TRA personnel. Adhere to all safety procedures given by the Safety Officer or NAR/TRA personnel. Refer to correct MSDS if I am unsure of any chemical's safety. Stay observant and aware of my surroundings when working with Long Beach Rocketry Team Notify the Safety Officer if I have any Safety concerns. I will agree to the following safety regulations regarding the range safety officer: Range safety inspections of each rocket before it is flown. Each team shall comply with the determination of the safety inspection or may be removed from the program. The Range Safety Officer has the final say on all rocket safety issues. Therefore, the Range Safety Officer has the right to deny the launch of any rocket for safety reasons. Any team that does not comply with the safety requirements will not be allowed to launch their rocket I understand that if I fail to follow the guidelines set forth above it will result in disciplinary action. I certify that I have a full understanding and have read this agreement entirely. Name Signature Date 14

16 4.5 Risk Assessment Risk Assessment tables are used to help identify potential hazards, and mitigations to ensure that they are avoided. To justify the likelihood of a hazard occurring a value has been assigned to each event corresponding to a percentage in that event will occur. This can be shown below. Table 5: Probability of Occurrence Likelihood Value Chance of Occurring Rare 1 1% to 5% Unlikely 2 6% to 30% Moderate 3 31% to 70% Likely 4 71% to 94% Almost Certain 5 95% to 100% Another method at looking at risk assessment tables can be by determining the harshness of an event occurring. A value of harshness will be assigned to each event to determine the potential of injury and damage to equipment. This can be shown below. Table 6: Assessment of Harshness Definition Value Injury and damage Insignificant 1 Result in minimal injury and minor damage to equipment Minor 2 Result in minor injury and minor reversible damage to equipment Moderate 3 Result in moderate injury and reversible damage to equipment Major 4 Result in serious injury and irreversible damage to equipment Disastrous 5 Result in death and irreversible damage to equipment Using the probability of occurrence and harshness, a risk level can be assigned through a risk assessment matrix. This matrix will identify the event as either a high, medium or low risk level. By the launch at the competition the goal is to have all risk levels at low. This will be done by changing design choices by any means to ensure the safety of the team and everybody around us. Probability Value (P) Table 7: Risk Assessment Matrix Harshness Value (H) Insignificant (1) Minor (2) Moderate (3) Major (4) Disastrous (5) Rare (1)

17 Unlikely (2) Moderate (3) Likely (4) Almost Certain (5) Risk Level Low (2-4) Medium (5-7) High (8-10) Personnel Hazard Risk Analysis Table 8: Personnel Hazard Risk Analysis Hazard Cause Effect P H Risk Mitigation Not following proper Injury to LBR team member, Injury from procedure and possible chemicals contact with injury s: burns, 3 3 Medium skin/eyes occurs skin irritation, eye irritation Injury from fiberglass dust or epoxy fumes Injury from machinery Injury during testing Proper procedure not followed when handling fiberglass or epoxy Incorrect use of tools, Not enough knowledge of machinery Proper procedure is not followed, team members being unaware of their surroundings Injury to LBR team member, possible injury s: headache, nausea, dizziness from inhalation of fumes Injury to LBR team member, possible injury s: cuts, hair getting caught in machinery, objects hitting personnel in the eye Injury to LBR team member, possible injury s: objects hitting personnel, cuts, scrapes 3 4 Medium 2 4 Medium 2 3 Medium Wear proper PPE including gloves, glasses, mask and clothing. Read MSDS of the chemicals you are using. Proper PPE must be worn always. When working with fiberglass or epoxy team members must work in a wellventilated area. Members will be briefed on proper procedure for handling fiberglass and epoxy. Proper PPE must be worn always. When working in the machine shop the all members will abide by the machine shop safety procedures. Safety briefings will discuss proper use of machinery and safety procedures in the machine shop. Safety Officer must document and review the procedure before any testing occurs. LBR team members are required to sign the safety contract ensure that they understand proper safety procedure. 16

18 Injury during launch, rocket unstable launch Injury during recovery, recovery system fails Team members not aware of the safety procedure, rocket not designed with proper stability, launch rail not proper length or high winds Team members not aware of the safety procedure during recovery of rocket, recovery system fails to deploy Injury to LBR team member, possible injury s: Rocket hitting personnel Injury to LBR team member, possible injury s: Rocket hitting personnel 1 5 Medium 2 5 Medium Safety Officer will brief members on the procedure to be followed during launch and are required to sign the safety contract ensure that they understand proper safety procedure. Team members will follow the minimum distance set by the NAR and follow all guidelines in the NAR launch safety and launcher section to lower the likelihood of injury. Safety Officer will brief members on the procedure to be followed during recovery and are required to sign the safety contract ensure that they understand proper safety procedure. Team members will follow rules set by NAR with regards to recovery and recovery safety Failure Mode Hazard Risk Analysis Table 9: Structure and Propulsion Risk Analysis Hazard Cause Effect P H Risk Mitigation Structure Rocket will lose Buckling of cannot control and airframe properly become unstable 3 3 Medium during handle stress and flight of flight unpredictable. Tail fins shear off during flight Structure Damage during transport Fins not properly secured to airframe Improper storage during transportation Rocket takes unpredictable flight path and becomes unstable Rocket will be unstable and unpredictable 2 3 Medium 1 2 Low Use proper material to ensure that it can handle the stress of the launch and flight. Ensure that adhesive used to secure fins is strong enough to handle the force of flight. Check the adhesive for cracks before launch. And confirm everyone is at the minimum distance set by the NAR and alert the Range Safety Officer. Check to ensure rocket is secure during transportation and that it is properly inspected before launch. 17

19 Motor fails to ignite Motor explodes on launch pad Fins not properly aligned Rocket Velocity not high enough leaving the launch pad Motor centering ring fails Delayed Ignition, Faulty motor Faulty motor Fins not assembled correctly Rocket to heavy. Thrust of the motor is not large enough Adhesive not properly applied to the centering rings Rocket could launch unexpectedly or not at all Rocket will be highly damaged and possible injury to spectators Rocket becomes unstable and spins uncontrollably Launch is unstable Motor to launch through the rocket. 2 2 Low 1 5 Medium 1 2 Low 1 4 Medium 2 4 Medium Motors will be purchased from vendors with a good reputation. Wait the 60 seconds required by NAR safety codes before approaching the rocket to ensure it is not a delayed launch. Use minimum distance table that is in NAR safety codes. When safe extinguish fire. Motors will be purchased from vendors with a good reputation. Ensure proper procedure is followed when assembling the fins. Run simulations to verify that the motor we have selected will provide a sufficient velocity leaving the launch pad. Verify that the constructions procedures are followed for apply adhesive. Table 10: Recovery Risk Analysis Hazard Cause Effect P H Risk Mitigation Parachute does not deploy Parachute has rip or tear Parachute gets tangled around rocket, Rocket does not split open Parachute gets rip while packaging in rocket, Parachute gets ripped while deploying Rocket will fall to the ground at high velocity and become damaged upon impact Rocket will descend quickly and become damaged upon impact 3 4 Medium 2 4 Medium Parachute will be properly integrated into rocket to reduce risk of getting tangled. Team mentor will ensure proper amount of black powder charge to split rocket open. Parachute will be carefully inspected before it is packaged. Team members will be careful during packaging of parachute. 18

20 Altimeter failure Rocket fails to separate Rocket descends to quickly Rocket descends to slowly Avionics malfunction Rocket separates from recovery system Faultily altimeter, Altimeter gets damaged during launch Black powder fails to ignite, Black powder fails to break shear pins Parachute not sized properly Parachute not sized properly Low power supply or incorrect assemble of avionics Parachute disconnects from U-bolt Parachute will not deploy, and rocket will fall to the ground at high velocity and become damaged upon impact Parachute will not deploy, and rocket will fall to the ground at high velocity and become damaged upon impact Rocket will descend quickly and become damaged upon impact Rocket will drift more than calculated Early or no deployment of parachute causing rocket to descend quickly and become damaged upon impact Rocket will descend quickly and become damaged upon impact 1 4 Medium 3 4 Medium 1 3 Medium 1 3 Medium 1 4 Medium 1 4 Medium Use more than one altimeter for redundancy. Which will reduce the chance of failure. The recovery system will not deploy if all the altimeters fail. Range Safety Officer will be informed, and action will be taken to keep everyone safe. Recovery system will be designed so that we use the correct amount of black powder to ensure the rocket separates. If the rocket does not separate, the Range Safety Officer will be informed, and action will be taken to keep everyone safe. Parachutes have been selected carefully to ensure that recovery is successful. Parachutes have been selected carefully to ensure that recovery is successful. Two avionics systems will be used for redundancy to reduce the chance of malfunction. Testing will be done before launch to ensure avionics is functioning properly Parachute cables and the U- bolts are designed to handle large loads. Table 11: Airbrakes Risk Analysis Hazard Cause Effect P H Risk Mitigation Airbrakes fail to deploy at desired altitude Programing failure or altimeter failure Rocket will fly higher or lower than desired altitude 2 1 Low Test the Airbrake system during the subscale launch to ensure that the airbrakes will deploy at the proper altitude. 19

21 Structural damage to airbrake system during launch or flight Flaps for airbrakes systems fly off during use Airbrake system does not retract during recovery Materials of airbrake system cannot handle the force of launch Flaps cannot handle the force of the air during use Programing or electronic failure Airbrakes will not deploy or become damaged Rocket will become unstable and flap will be uncontrollable falling from the sky Damage to the airbrake system 2 2 Low 3 5 High 2 2 Low Verify through testing that the airbrake system can handle the amount of force that is applied during launch and flight of the rocket. Verify through testing that the flaps for the airbrake system can handle the amount of force that the air will generate on them. Testing will be done to ensure that the airbrake system will retract at a certain altitude after they have deployed. Table 12: Rover Risk Analysis Hazard Cause Effect P H Risk Mitigation Rover damaged on landing Rover damaged during flight Rover flips over Rover gets stuck on rock or hill Faster than normal landing, Payload not secure in place Payload not secure in place, Rocket unstable during flight Rover not designed to handle the terrain Rover not designed to handle various objects that could be in the terrain Payload becomes damaged and inoperable Payload becomes damaged and inoperable Rover becomes stuck and unable to make distance requirement Rover will be stuck and potential damage itself making it unable to make the distance requirement 2 2 Low 1 2 Low 3 1 Low 3 2 Medium Table 13: Rover Deployment (RDM) Risk Analysis Ensure that the parachute is properly chosen to slow down the rocket. Make sure that payload is secure in place before the rocket is launched. Ensure that the rocket design is stable before we launch, and that the payload is secured in place. Ensure that the rover is designed to handle the type of terrain at the launch site. Verify through testing that rover is capable of handling terrain at the launch site. Design the rover in such a way that objects such as rocks and hills will not affect the rover s performance. Verify through testing that the rover can handle the terrain of the launch field. 20

22 Hazard Cause Effect P H Risk Mitigation RDM Materials for RDM will becomes RDM system become damaged cannot handle damaged and 2 2 Low during the force of rover will not launch or launch or flight deploy flight RDM does not deploy when activated RDM deploys during flight RDM deployment damages the rover Programing or electronic failure Programing or electronic failure Poor design of RDM system Rover will not deploy Nose cone will separate from rocket during flight causing rocket to become unstable and lose control Rover will become damaged and not function Environmental Risk Analysis 2 1 Low 2 5 Medium 2 2 Low Table 14: Environmental Risk Analysis Ensure that the materials chosen for the RDM system are capable of handling of force or launch and flight. Verify through testing that the electronics will deploy upon being activated. Verify through testing that the electronics will only deploy when they are wanted to. Test the deployment system with model of the rover to ensure that it can handle deploying the rover without damaging it. Hazard Cause Effect P H Risk Mitigation Low clouds low visibility Rocket launched in conditions of low clouds and low visibility Rocket is not visible so there no way to tell if rocket launch is going well 1 3 Low Rain High winds Rocket launched in conditions of rain Rocket launched in conditions of high winds Electronics of rocket could become damaged Rocket will become unstable 1 3 Low 2 3 Medium Check forecast for planned launch days to ensure that the weather will be ideal for launch. Check forecast for planned launch days to ensure that the weather will be ideal for launch. Keep electrical equipment safe from the rain to avoid water damage. Check forecast for planned launch days to ensure that the weather will be ideal for launch. If wind speed exceeds 20 miles per hour the launch will be aborted in compliance with NAR safety codes. 21

23 Trees Motor exhaust cause fire on ground Battery explosion breaks rocket and scatters parts Team members leave trash at launch site Rocket lands in area with many trees Exhaust allowed to flow onto the surrounding ground Batteries may explode unpredictably if not charged properly Team members not cleaning up after themselves Rocket could hit tree and damage components inside the rocket Fire can occur at the grounds around the launch rails Explosion cause parts of the destroyed rocket to be spread all over the launch field Potential harm to environment due to trash being left at launch site 1 2 Low 2 2 Low 1 4 Medium 1 2 Low Ensure that trees are out of range of the drift distance so that the rocket does not hit the trees or get stuck in the trees. As stated by the NAR safety codes in the launcher section a blast deflector will be used, and dry grass will be cleared from the area Team members will follow correct procedure when charging batteries. Team members will carefully check the launch site before leaving to ensure that no trash is left at the site Project Risk Analysis Table 15: Project Risk Analysis Hazard Cause Effect P H Risk Mitigation Team must spend time to Design of the determine Change of rocket is second design. design of determined to Effecting time, the rocket be unsafe or to budget and costly resources of the project 3 3 Medium Failure to meet deadlines of the project Team did not have time to complete all aspects of the rocket Project could be scrapped, or design may need to be changed 3 4 Medium Design has been chosen based on how feasible and effective it is. Testing and analysis will be done to ensure that our design is feasible. Communication will be necessary to ensure that the team understands the deadlines that have been given. Weekly meetings will be had to ensure progress of the project is being made. 22

24 Failure of parts of rocket during testing Team needs to purchase parts that go over the team s budget Damage to rocket after subscale or full-scale launch LBR is prohibited from continuing the project Parts of rocket determined as unusable during testing Change in design or failure of parts Rocket did not perform as intended Team fails to comply with NASA policies Team must spend time to determine an alternative to that part. Effecting time, budget and resources of the project Budget will need to be reevaluated which cost time and money Team will have to modify the design, which will take time and money depending on how much parts are needed to be replaced Team will not be allowed to complete project 2 2 Low 3 3 Medium 1 4 Medium 1 3 Low Parts will be inspected many times and analysis will be done to ensure parts are capable of handling testing. A margin will be used to ensure that the team is under budget. Parts will be chosen to ensure an ideal cost. Parts of the rocket will be tested and analyzed frequently to ensure that they will function as intended to. Team leads will ensure that all policies are followed. 4.6 Compliance with Federal, State and Local Laws The Long Beach Rocketry Team will comply with all federal, state and local laws on unmanned rocket launches and motor handling. The launch vehicle is classified as a class 2 high power rocket. The team will be launching with the Rocketry Organization of California at Lucerne Dry Lake launch site and with the Friends of Amateur Rocketry in Cantil, California. The Rocketry Organization of California ensures that all rocket launches comply with FAA regulations and in strict compliance with all regulations of the California State Fire Marshal. By launching with ROC, the team complies with Federal Aviation Regulations 14 CFR, Subchapter F, Part 101, Subpart C and fire prevention, NFPA 1127 Code for High Power Rocket Motors. To comply with Amateur Rockets, Code of Federal Regulation 27 Part 55: Commerce in Explosives, the mentor will either acquire a federal low explosives user permit or find a person 23

25 that has a federal low explosives user permit to purchase or transport the motor to Huntsville, Alabama. 4.7 Handling of Rocket Motors Purchase and Storage The Long Beach Rocketry Team recognizes that rocket motors will be purchased by the NAR/TRA certified team mentor. The team mentor will be in possession of the motor until the team's launch at the at the Lucerne Dry Lake Bed or the FAR (Friends of Amateur Rocketry) sites. Prior to the launch all motors will be disassembled and stay in the original packaging. Handling and Use The rocket motors will only be handled by the team members while under the supervision of the team mentor. The team mentor will oversee the preparation of the rocket motor for all launches. Transportation If the team mentor possesses a federal low explosive user permit he may purchase a rocket motor in Huntsville, Alabama. He will coordinate the purchases and shipping of a rocket motor to Huntsville, Alabama through a third party possessing a federal low explosives user permit, if the team mentor does not have said permit. 4.8 Range Safety Regulations All members of the Long Beach Rocketry Team will follow the range safety regulation as stated in the Long Beach Rocketry Team Safety Agreement. The team members understand and will abide by the following regulations: Range safety inspections of each rocket before it is flown. Each team shall comply with the determination of the safety inspection or may be removed from the program. The Range Safety Officer has the final say on all rocket safety issues. Therefore, the Range Safety Officer has the right to deny the launch of any rocket for safety reasons. Any team that does not comply with the safety requirements will not be allowed to launch their rocket. 4.9 Safety Checklist Safety Checklist: Payload System To be checked and signed by Safety Officer & Payload Lead. Safety Officer & Payload Lead Signatures: Note: If any damage is identified, immediately inform both the safety officer and the Payload 24

26 Lead. The launch pad will be approved safe only if no materials were damaged or an alternative action could be decided upon- by the safety officer and all Payload system leads. Required Equipment: Phillips head screwdriver Lithium Polymer Battery size TBD Screws, nuts, and fasteners Stepper motor Clay putty Lithium Polymer Battery Voltage tester Extra Arduino and Raspberry Pi for testing Prior to leaving for launch site: 1. Ensure all mounting shelves are tightly screwed into payload bay. 2. Verify that nose cone is flush with the payload bay. 3. Ensure that limit switches and sensors are in place and functional. 4. Inspect parafoil for tears or damage. 5. Fold parafoil for and prepare parafoil for deployment in nose cone. 6. Ensure cords of parafoil are correctly placed to avoid tangling. 7. Verify motors are operational and functioning. 8. Verify motor is operational by using Arduino to run a test. 9. Verify electronics are functioning as expected. 10. Verify camera operation using spare Raspberry Pi. 11. Close nose cone for transit. At launch site: 1. Unscrew bottom mounting shelf. 2. Use Lithium Polymer battery voltage tester to ensure functioning battery. 3. Connect electronics system to battery. 4. Place mounting shelf into payload bay and screw in place. 5. Place clay putty on seams on nose cone. Warning: If something is not functioning properly or meeting the requisites stated above inform the Payload Lead and Safety Officer. The launch pad will be approved safe only if no materials were damaged or an alternative action could be decided upon by the safety officer and all Payload system leads. Safety Checklist: Motor Preparation To be checked and signed by Safety Officer and Team Mentor. Propulsion Representative and Team Mentor Signatures:

27 Prior to leaving for launch site: Sustainer Propulsion Bay Assembly Checklist: Required Equipment: Grease Motor Forward Extended Plugged Tapped Closure Motor Casing Aft Closure 75mm Retainer End Cap Required Personal Protective Equipment (PPE): Nitrile Gloves CAUTION: Protective gloves must be worn when applying grease to the motor. Note: When damage inspection is conducted, any damage detected must be immediately reported to the team captains and safety officers. The affected areas will be analyzed, and the rocket shall be deemed safe to launch if the damage is rectifiable. Warning: Proper certification is required to handle the motor. The assembly must be stored in an area devoid of moisture and open flames. At Launch Site 1. The AeroTech K805G-P motor is taken out of its antistatic bag. 2. The instructions provided with the motor is followed to prepare the motor to be placed into the motor casing. 3. The forward extended plugged tapped closure and aft closure is screwed on the motor casing. 4. Motor casing is slid into the propulsion bay of the rocket mm retainer cap is screwed onto the end of the propulsion bay, securing the motor casing to the propulsion bay. 6. The eyebolt along with its bulkhead is screwed on the forward extended plugged tapped closure. Safety Checklist: Recovery & Avionics Bay Preparation To be checked and signed by a recovery team member. LBR Team Member Signatures: Prior to leaving for launch site: 26

28 Required Equipment: Recovery Avionics Bay Electronic matches (x4) Masking Tape Scissors Black powder Wadding Spoon Paper Required Personal Protective Equipment (PPE): Safety glasses Black Powder Charge Preparation 1. Place all required equipment and Required Personal Protective Equipment are place in a container to take to the launch site. 2. Make sure that Black powder container is properly closed. Safety Checklist: Recovery To be checked and signed Recovery Safety representatives. Recovery Representative Signatures: Prior to leaving for launch site: Parachute Packing Required Equipment: Main Parachute Drogue Parachute Main Deployment bag Drogue Blanket Shock Cord (2x) Shock Cord Protective sleeves (2x) Quick Links (6x) U - Bolts (4x) 1. Check for any damage including burns, cut, fraying, or any other visible damage for the following parts: a. Parachutes b. Shock Cords c. Blankets d. Deployment Bag 27

29 e. Quick Links f. Eyebolts Note: Any damages identified are to be reported to team leads and safety officer. 2. Lay drogue parachute canopy flat on the ground. 3. Check line for any entanglement, if entanglement is found, untangles the line as needed. 4. Fold parachute and place in bag. 5. Repeat steps 1 through 4 for main parachute. 6. Place all required equipment in a container to take the launch site. Avionics Bay: Avionics coupler Avionics wooden boards 3D Printed support tray Avionics battery holder (2x) Precision flathead screwdriver Standard Phillips head screwdriver Altimeter (2x) & associated hardware BRB900 GPS tracker 9 volt battery (4x) 9 volt battery clips/connectors (2x) E-matches (8x) Black powder Arming Switch (2x) Wires #2-56 nylon screws (6x) Zip-ties 1. U - bolts are place on each bulkhead with washers and bolts 2. Attach two PVC caps to each of the bulkhead with #4-40 screws 3. Batteries and battery connectors are placed inside the battery holder; check connector tips for proper soldering 4. The two wood boards are then screwed together with the 3D printed support tray in between the two boards (make sure the screw and the nut are loose enough to rotate the plastic holders.) 5. Slide the two 5/16 rods through each side of one of the bulkhead and place a washer and nuts on the side of the bulkhead facing outside the avionics bay. 6. Slide the two boards and battery the 3D printer support tray into the two threaded rods. 7. Tighten the screws holding the two boards together. 8. Check that both altimeters are working properly. Make sure all electrical wirings are connected correctly. Pull slightly on all electrical connection points to ensure a secure connection. 28

30 9. Check that the GPS tracker is working properly. 10. Check that all 9V battery has the proper voltage. 11. Check that arming switch is in the off position. 12. Check that all nuts & bolts are properly tighten. 13. Assemble the altimeters to the avionics bay sled and connect the batteries and arming switch to altimeters. 14. Set arming switch on and check that altimeters function properly. Turn off arming switch if altimeters work properly. 15. Place the GPS tracker on avionics bay sled (do not connect battery to transmitter). 16. Place all required material in a container to take to launch. Launch Day Procedures Parachute Assembly: Required Equipment: Blanket Deployment Bag Shock chords (x2) Shock Cord Protective sleeves (2x) Quick links (6x) Main parachute (60") Drogue parachute (18") 2-56 Nylon screws (12x) 1. Untangle any lines on the parachutes if any. 2. Link all shock cord segments with quick links. 3. Fold up both parachutes following the procedures from Fruity Chutes. 4. Connect the main parachute to the shock cord connecting the avionics to the motor section of the rocket 5. Attach the drogue parachute to the harness connecting the avionics to the payload. 6. Slide the main parachute and harness into the deployment bag. 7. Connect one quick link to the U - bolt that is connected to the payload 8. Slide the deployment bag with the main parachute inside into the main parachute bay. 9. Connect the other end of the shock cord to the U - bolts connected to the avionics bay. 10. Slide the avionics into the main parachute bay. 11. Screw in shear pin into shear pin holes for the main parachute bay. 12. Connect drogue parachute to harness and fold flame shield blanket over drogue. 13. Connect one end of the harness to the eye U - bolt that is connect to the motor. 14. Slide the covered drogue parachute into the drogue bay. 15. Connect the other end of the harness to the U - bolts connected to the avionics bay. 16. Slide avionics bay into the drogue bay. 17. Screw in shear pin into shear pin holes for drogue bay. 29

31 Avionics Bay: Required Equipment: Multimeter Precision flathead screwdriver Wire Strippers Optional Equipment: Masking tape Zip-ties 1. Slide the wood board with all the electronics into the coupler 2. Turn on GPS tracker by connecting battery to the tracker. 3. Connect wires form the terminal block of the second bulkhead to altimeters. (Make sure E-match are not connected to block terminals) 4. Slide in second bulkhead into the coupler. 5. Place a washer and nut to each of the rods on the second bulkhead and tighten 6. Connect wire sticking out of bulkhead to block terminals. (Not E-matches) 7. Shake Avionics Bay to make sure altimeters do not turn on. If altimeters do turn on during shake, loosen arming switches screws. Repeat until altimeters no longer turn on during shake. Ejection Charges: Required Equipment: Precision flathead screwdriver Digital Scale Spoon Masking tape E-matches Wadding CAUTION: Safety glasses must be worn during charge preparation process. DANGER: Electronic matches are explosive. To avoid misfires, the black powder charges and electronic matches must be kept clear of batteries and open flames. Warning: If case of a black powder leak, the resulting ejection charges may be weaker or fail to detonate. This would cause a fatal failure in the recovery subsystem. 1. Measure needed black powder for PRIMARY ejection charge for drogue and main a. Primary main ejection charge: TBD b. Primary drogue ejection charge: TBD 2. Cut off excess wire off the E-match as needed. 3. Place the E-match inside the PVC cap. Make sure the E-match is lying flat on the bottom of the cap 4. Place the specific measured amount of black powder into the PVC cap. 5. Place wadding on top of the black powder and compress the wadding until the PVC cap is full. 6. Place tape over the top and around the PVC cap so the wadding and black powder and wadding does not fall out. 30

32 7. Mark the PVC "P" and the amount of black powder in the cap. 8. Measure needed black powder for SECONDARY ejection charge for drogue and main a. Secondary main ejection charge: TBD b. Secondary drogue ejection charge: TBD 9. Repeat step Mark the PVC "S" and the amount of black powder in the cap. 11. Wait for avionics assembly to be complete and check that the altimeters are off. 12. Connect each the E-match to its corresponding terminal blocks. 13. Check that all connections are correct and secure. Safety Checklist: Overall Final Assembly Checklist Launch Vehicle Lead and Structures Lead Signatures: Required Equipment: Allen Wrench Set SAE Screwdriver Set (Large and Small) Socket Wrench Set Aluminum tape Socket Cap Screws Shear pins 1. Ensure the shear pins are fastened between the coupler and airframe, check for any flexing of the rocket between those sections. 2. Tape the igniter to the outside of the propulsion bay. 3. Visually inspect the rocket to ensure all systems are go. Safety Checklist: Clear to Leave the Launch Pad: All sections of the safety checklist prior to the At the launch pad checklist must be complete before leaving for the launch pad. All Lead's signatures are required prior to launch. General Preparations: Propulsion/Aerodynamics: Recovery/Avionics: Overall Final Assembly (LV & Structures): Team Lead: Safety Officer: Safety Checklist: At Launch Pad Checklist Required Equipment: 31

33 Pen or pencil Level 2 Certification card. Payload Featherweight Screw Switch (Arming Switch) (2x) Level Bubble Ruler 1. Verify that permission has been granted by the RSO to launch. 2. Place rocket onto the Payload launch rail. 3. Ensure that the Payload Launch Rail is pointed in the desired direction, or in direction ruled necessary by RSO. Use level bubble ruler to ensure launch rail is horizontal. 4. Arm all electronics. Check for correct LED readout, beeping pattern, etc. 5. Before leaving launch pad area, double check that all electronics are still operating correctly. 6. Clear and Leave the launch pad area. Safety Checklist: During and After Flight (DAF): Location: Temperature: Pressure: Wind Speed: Flight Events: First Event: Drogue deployment Observer Signature: Time: Second Event: Main deployment Observer Signature: Time: Landing Events: Propulsion Bay Observer Signature: Time: Nose Cone Observer Signature: Time: Video Recorder Signature: Photographer Signature: LV Retrieval Team Member #1: LV Retrieval Team Member #2: LV Retrieval Team Member #3: 32

34 Required Equipment: Phone timer Featherweight Screw Switch (Arming Switch) (2x) Screwdriver Camera 1. LV Retrieval Team members should be ready to move within a few seconds notice. 2. Start phone timer at liftoff and call out each event recorded time. 3. Maintain line of sight with rocket at all times. Indicate any obscurities. 4. While secure rocket, disarm all rocket recovery systems first 5. Take photo of how the rocket landed, recorded any damage 6. Disassemble the rocket looking for any signs of damage or fatigue. After Flight Checklist: To be checked and initialed by Recovery/Avionics Lead. Recovery Lead Signatures: Check for any damage on the shroud lines. 2. Check for any damage on the attachment section of the rocket. 3. Check for any damage on the parachutes. 4. Check for any damage on the blankets. Damage found on shroud lines? Y / N Notes: Damage found on attachment point? Y / N Notes: Damage found on blanket? Y / N Notes: Damage found on the parachute? Y/N If yes, photo approximate location below: Notes: Altitude Achieved: Motor Used: 33

35 Event #1 Success: Yes or No Event #2 Success: Yes or No Team Lead Approval: 1. 34

36 Section 5: Launch Vehicle Criteria The launch vehicle will be constructed primarily with fiberglass, carbon fiber, aluminum, plywood and ABS. It must safely house the interior components of the launch vehicle throughout the duration of the flight and withstand the forces induce during takeoff and ascent. In order to achieve an apogee of 5280 feet, the launch vehicle will be optimized to minimize mass. Furthermore, the launch vehicle must successfully deploy its recovery parachutes and land without causing any damaged or safety hazard. The launch vehicle can be divided into seven sections, which will outline in the figure below: 5.1 Mission Success Criteria Figure 2: Launch Vehicle Overview 1. The launch vehicle shall be successfully departing from the launch rail. 2. The launch vehicle shall carry a payload up to an apogee of 5280 feet ±100 feet. 3. All the recovery events shall successfully occur at the programmed altitude. 4. The launch vehicle shall have a stable takeoff and ascent. 5. The launch vehicle shall be successfully recovered in a reusable condition. 5.2 Selection and Rationale of Launch Vehicle 35

37 Material Selection There are many factors to consider when deciding what material should be used for what rocket component. It is essential that the mechanical properties, as well as characteristics such as weight and price are considered during this process, as the overall performance of the launch vehicle relies significantly on these factors. Below is a pros and cons of the material properties for carbon fiber, fiberglass, and blue tube. Table 16: Materials Comparison for the Airframe Pros and Cons Materials Carbon Fiber Pros Highest Yield Strength High Strength to Weight Environmentally Resistant Cons Very Expensive Very Conductive Less Fatigue Resistance Fiberglass Good Yield Strength Low Conductivity Affordable Price Moderate Strength to Weight Moderate Fatigue Resistance Blue Tube High Strength to Weight Inexpensive High Fatigue Resistance Low Yield Strength Low Environmental Resistance Below is a comparison of the material properties for carbon fiber, fiberglass, and blue tube, as well as a price and weight comparison. Each characteristic was measured on a scale of 1 to 5, 1 is poor and 5 is excellent. Table 17: Materials Properties Trade Study Material Properties Trade Study Carbon Fiber Fiberglass Blue Tube Yield Strength (MPa)

38 Fatigue Resistance Conductivity Environmental Resistance Adhesive Properties Strength to Weight Price Total When totaling the ratings, fiberglass is shown to have the best score for several reasons. While the yield strength of fiberglass is slightly less than carbon fiber, it is significantly greater than blue tube, and will outperform a blue tube airframe in terms of strength. Fiberglass is more environmentally resistant than blue tube, considering that blue tube is reinforced cardboard and will not stay intact if it were to experience rain or other weather conditions. Fiberglass has better fatigue resistance than carbon fiber, as fiberglass will bend and deform before it breaks, whereas carbon fiber will snap or fracture suddenly. Fiberglass has extremely low thermal and electrical conductivity, and as an airframe will provide good insulation to internal electrical rocket components. Fiberglass has better adhesive properties than blue tube, and will allow more ease when epoxying fins to the airframe. While fiberglass is more expensive than blue tube, it is much less expensive then carbon fiber. Airframe Diameter One trade-off was the diameter of the launch vehicle, particularly choosing a diameter length that would best allow LBR to accomplish its task. Below in Table X and Table XX, pros and cons tables for each consideration diameter can be found. 37

39 Table 18: Pros and Cons Table of 5.15 inches Diameter Pros Cons More space would be available for needed payloads and recovery systems. A larger diameter would allow a decrease in length to accommodate the same volume, reducing bending moments. The cost of fiberglass parts, motors, and most launch vehicle components would be greater than a 4-in. diameter launch vehicle. A larger size launch vehicle increases the mass of the launch vehicle, resulting in larger required motors and increased kinetic energy of independent sections on landing. A larger length of coupler would be required for coupler joints between sections (5 in.). Table 19: Pros and Cons Table of 6.17 inches Diameter Pros Cons More space would be available for needed payloads and recovery systems. A larger diameter would allow a decrease in length to accommodate the same volume, reducing bending moments. The cost of fiberglass parts, motors, and most launch vehicle components would be greater than in a 4-in. or 5.15-in. diameter airframe. A larger size launch vehicle increases the mass of the launch vehicle, resulting in larger required motors and increased kinetic energy of independent sections on landing. A larger length of coupler would be required for coupler joints between sections (6 in.). As shown from tables above, certain pros and cons result from increasing or decreasing the size of the airframe. Importantly, larger-diameter designs allow more space for components to be housed, such as the rover and RDM; however, it results in greater mass, and larger motors would be required. The current launch vehicle design consists of a 6.17-in. diameter fiberglass launch vehicle that is 108 in. long and has a mass of 42 lb. The 6.17 in. diameter airframe comes with most of the benefits of a larger diameter, e.g., more space for required payloads and a decreased length. The most promising alternate design would be a 5.15-in. diameter launch vehicle, which would reduce mass, cost, and other benefits shown above in Table X. The trade-off for the smaller diameter would mainly be a decrease in space for scientific payloads. The alternate design for a 38

40 5.15-in diameter launch vehicle would be very similar to the current launch vehicle design, with an appropriate increase in length to account for the decreased diameter. Nose Cone Profile The shape of a rocket s nose cone directly affects the rocket s performance as different shapes produce different drag coefficients for the launch vehicle. The drag of three common nose cone shapes, elliptical, power, and ogive were calculated using SolidWorks Flow Simulation. Four simulations were run for each nose cone shape at 500, 600, 700, and 800 ft./s. The drag coefficient of the rocket with each shape is calculated and averaged; the pressure surface plot of the rocket with each nose cone was compared. Figure 3: Pressure Surface Plot with Elliptical Nose Cone at 600 ft./s Figure 4: Pressure Surface Plot with Power (n=.4) Nose Cone at 600 ft./s 39

41 Figure 5: Pressure Surface Plot with 4:1 Tangent Ogive Nose Cone at 600 ft./s The coefficient of drag of the launch vehicle determined from the equation below. The drag force is obtained through the CFD simulation. C d = 2D ρav 2 Table 20: Simulated Coefficient of Drag for Each Nose Cone Profile at Four Different Velocity Power Series Elliptical Ogive Velocity (ft./s) Coefficient of Drag Velocity (ft./s) Coefficient of Drag Velocity (ft./s) Coefficient of Drag Average: Average: Average: As seen above, the Ogive nose cone has the lowest coefficient of drag when compared to elliptical and power series nose cone. The weight of each nose cone profile was also compared. The weight of each nose cone profile was determined with a density of 1.07 oz./in 3, which is approximately the density of fiberglass. The weight of each nose cone profile with a wall thickness of 0.1 inch is shown below in Table

42 Table 21: Nose Cone Profiles Weight Nose Cone Shape Weight Comparison Power Series Elliptical Ogive Weight (oz.) A trade study comparing the power series, elliptical, and ogive nose cone is shown below: Table 22: Nose Cone Trade Study Nose Cone Trade Study Power Series Elliptical Ogive Coefficient of Drag Mass Price Total When comparing the calculated average drag coefficients of the rocket with each nose cone shape, the results of the SolidWorks Flow Simulation show that the ogive nose cone has the lowest drag coefficient, and therefore will produce less drag than the power or elliptical nose cones. When doing a weight comparison of the entire rocket, the ogive also adds the least weight to the launch vehicle. When comparing the pressure surface plots of the rocket with each nose cone, the images show there is less pressure at the tip of the ogive nose cone, compared to elliptical or power. This is determined by the fact that the plot shows less red (representing higher pressure) at the tip of the ogive from Figure 3, 4, and 5 above. It is for these reasons the team has decided to use a 4:1 length to diameter tangent ogive shape for the nose cone. Fin Mounting System There are many factors to be considered when deciding how the fins will be attached to the rocket. The two fin attachment systems being considered are: 41

43 1. Epoxying the fins to the airframe and creating fin fillets with epoxy and epoxy additives 2. Screwing fins into an external fin can 3. Screwing fins into an internal fin can Below is a comparison between the different factors being considered for each option. They are rated on a scale of 1 to 5, 1 is poor and 5 is excellent. Table 22: Nose Cone Trade Study Epoxying Fins to Airframe External Fin Can Internal Fin Can Coefficient of Drag to the launch vehicle Integration Weight Stability Manufacturability Total The epoxy fins to airframe method was chosen for the fin mounting system to be utilized in the launch vehicle due to it modularity and low weight when compared to other fin mounting systems. In addition, weight has always been an issue in previous year competition because a heavy propulsion bay will cause the static stability of the launch vehicle to go down below 2.0 calibers. 42

44 5.3 Launch Vehicle Design Figure 6: OpenRocket Side View of Proposed Launch Vehicle Design An OpenRocket side view image of the launch vehicle can be seen in Figure 2. The launch vehicle will have a 6 inches diameter airframe because an additional space is needed for this year s scientific payload. To achieve an apogee of 5280 feet, the main goal of the launch vehicle this year will be to minimize mass while optimizing for maximum efficiency. The total length of the launch vehicle is 108 inches. Table 23: Launch Vehicle Section Lengths Section Length (in) Weight (oz.) Nose Cone Payload Bay Recovery Bay Propulsion Bay Total Length and Weight Using estimated payload masses and known material densities, the static stability margin of the launch vehicle in its current configuration is 2.65 calibers with a total weight of 33.6 lbs. An OpenRocket model was created to verify all the equations from section 4.1 which calculate the location of the center of gravity, center of pressure, and apogee of the proposed launch vehicle. The specifications of the OpenRocket Simulation of the launch vehicle are shown in Table 17. Table 24: Launch Vehicle Flight Specifications Specifications of the launch vehicle Numerical Value Center of Gravity (in. from nose cone) Center of Gravity after the motor burnout Center of Pressure (in. from nose cone) Static Stability Margin (cal)

45 Static Stability Margin after the motor burnout (cal) 3.76 Rail exit velocity (ft./s) 66.9 Max acceleration (ft./s^2) 240 Predicted Apogee (ft.) 5472 Nose Cone Design Figure 7: Imagine of a Proposed 6 in 5:1 Ogive Fiberglass Nose Cone with Metal Tip The ogive nose cone, as seen above in Figure 3, will be fiberglass. This choice is based on the successful use of this design in the previous years of the competition. It is necessary that the nosecone be high-strength because it will house part of the rover and the RDM system. The nose cone will be secured to the launch vehicle via four 2-56 SHCS nylon shear pins to ensure safety 44

46 during flight. The nose cone section will attach to the forward end of the payload bay. This shape was chosen due to its combination of low mass and relatively low efficient of drag when compared to other nose cone profiles. Figure 8: Nose Cone Simulation Results Eight SolidWorks flow simulations were performed on both shapes at four different velocities. A pressure surface plot for each shape was created to view the different pressure values on each part of the nose cone during a maximum velocity of 240 m/s. Each simulation computed the drag force which was then used to calculate the drag coefficient. Table 25: Flow Simulation Data Velocity ft./s) Drag Force (lbf.) Coefficient of Drag Payload bay Figure 9: The Payload Bay 45

47 The payload bay is responsible for housing the rover and RDM. The payload bay consists of a 24-inch fiberglass nose cone which will house a inch diameter, custom-made aluminum bulkhead. The bulkhead will be installed 3 inches into the nosecone for the RDM to attach to. The nose cone will be secured to the under section by the RDM s threaded rods further down the rocket. A 14-inch-long, 5.5-inch diameter payload section will be installed below the nose cone to house the rover. Another aluminum bulkhead with the same dimensions will be attached 5 inches from the bottom of this section. Under the second bulkhead, the motor s that will turn the threaded rods will be housed. The threaded rods will be connected through this aluminum bulkhead to these motors and a threaded housing unit. The bulkhead will be secured via 3/8 th inch screws installed from outside the rocket into the bulkhead. This section will end with another aluminum bulkhead that will be secured to the section below it using 3/8 th inch screws. An eyebolt will be installed in the center of the final bulkhead to provide an attachment point for the entire payload bay. The main parachute will use a 12-ft long ⅜-in nylon webbing shock cord that is attached with quick links to eyebolts from the forward side of the avionics bay to the aft side bulkhead located on the midsection payload bay. Recovery bay Figure 10: Recovery Bay The Recovery bay section functions as the launch vehicle s electronics casing (avionics bay); it also carries the main and drogue parachute sections, as well as the ejection charge deployment system. The avionics bay itself serves as the structural coupler to connect the propulsion and payload bays together. The avionics electronics bay will have a total length of 13 inches long 46

48 G12 coupler tube with the outer diameter of inches and an inner diameter of inches with the main chute accompanying an additional 16 inches and the drogue chute taking an additional 8 inches to make up the total length of the avionics section. In the center of the avionic bay, there is a 1 inch strip of G12 airframe with an outer diameter of 6.17 inches and inner diameter of 6 inches which will be epoxied to the avionic bay, centered 6.5 inches from either end of the avionic bay. Additionally, four holes of 1/5 inch diameter will be drilled around the switchband with equal spacing between them. These will allow the avionic bay to measure the pressure the launch vehicle is experiencing. To hold the avionic bay together, two holes will be drilled in both aluminum bulkheads and threaded rods will be put through the entire avionic bay. Washers will then be inserted around the threaded rods and steel hex-nuts will be threaded on. U- bolts will be inserted into the center of the bulkheads, secured with washers and steel nuts and hot glued in place. Four shear pins holes (4/40 inch diameter) will be drilled to the avionic bay. The avionics bay coupler that holds all recovery electronics will be tethered with 27 feet shock cords and quick links attached to U-bolts on the aft and forward sides of the propulsion and payload bays. The propulsion and payload bays will be assembled and connected with four 1/8 nylon shear pins on the forward and aft sides of the avionics bay. The black powder ejection charges will initiate the deployment of the main and drogue chutes. The ejection charges will be ignited by commercially-available electronic matches that will be energized and ignited at specific altitudes (apogee for drogue, and 500 feet for main) by commercial barometric altimeters. When ignited, the ejection charges create hot gasses and pressure that generate an external force that pushes on the forward and aft sections of the payload and propulsion bays, separating the sections for a successful dual deployment. Figure 11: 3D Printed Center Housing Case 47

49 The avionics sled will be modeled and printed using computer-aided design; it will be printed as a slab component with ABS plastic, with additional cut wooden plates attached to the front and back of the sled. This method of avionics construction decreases the amount of time taken for assembly increasing the overall durability of the avionics bay. Threaded rods will run through the sled with the premade holes during printing and will connect the sled in a fixed position to be secured with the respective bulkheads in the avionics bay. Using ABS plastic prevents any disturbance to the delicate electronics on the wooden plates acting as its own monolithic component in the avionics bay. Given its initial modeling with CAD, using 3D printed components allows for multiple iterations of the design. Any holes or possible openings within the avionics bay bulkheads will be sealed with industrial epoxy to prevent any corrosive damage by the heated gasses after the ejection charges ignite. Propulsion bay Figure 12: Propulsion Bay Assembly The propulsion bay serves the purpose of housing and retaining the motor, airbrake bay, providing fins for stability and a drogue bay. The propulsion bay airframe will be constructed from 6- inches diameter fiberglass tubing and consists of a 42 inches long section of G12 airframe. On the aft end of the booster section, a 19 inches long 75-mm diameter G12 motor will be secured by using three aluminum centering rings. The centering rings that secured the motor tube also compress four through the wall fins, adding strength to the fin attachments 48

50 Motor Tube Figure 13: Motor Casing The motor tube will be commercially purchased to accommodate an Aerotech L1150 engine. This casing is RMS-75/5120 casing with forward seal disk manufactured by Aerotech. This motor tube is made from Aluminum weight: 1018 g Length/Thickness: 23.72" Width: 3". Trapezoidal Fins 49

51 Figure 14: Fin Geometry The purpose of fins on a rocket is to provide stability and to maintain the intended trajectory of the launch vehicle. Placing fins on a rocket serves to place the center of pressure aft of the center of gravity, which allows the perturbing forces on the center of pressure (such as wind) to be balanced. The team has chosen to utilize four trapezoidal shaped fins, with airfoiled edges to reduce pressure drag and induced drag. Because the launch vehicle will fly in the subsonic regime, the leading fin edge will be rounded and the trailing edge will be tapered. The fins will be constructed of carbon fiber, and will be lightweight and extremely durable. They will be attached externally to the airframe using high temperature epoxy. Fin fillets will be integrated to reduce interference drag and to increase each fin s structural strength. The fillets will be created also using high temperature epoxy, and with an epoxy filler and additives. Centering Rings For the centering ring system, we plan to use three centering rings each separated 8 inches from each other and 5 inches from the top of the engine. LBR has made the decision to deviate from last year s design which used plywood, because plywood wasn t strong enough. Reduced stability caused by flex and misalignment of the plywood centering rings was a major cause of failure in last year s rocket. The centering rings will be attached using epoxy on the outside of the ring that will adhere it to the inner side of the propulsion bay, which was the same method used last year. The aftmost centering ring will be manufactured by the team using plasma cutters from 6061 Aluminum with an outer dimension of 6 and 3 ID. LBR plans to manufacture the aluminum ring using the plasma cutters so they can be built to their exact specifications. On the aft centering ring, there will be 6 holes drilled 60 degrees equilaterally apart. The holes will be made to fit a 3/8th inch screw. To attach to the rocket, there will be the same holes on the outside body of the propulsion bay so it can be screwed in from the outside. 5.4 Motor Selection and Alternatives The choice of propulsion system depends on the apogee requirements. Below is a comparison between different propulsion systems for the mission along with their thrust curves. Motor Cesaroni L1350 CS-P Cesaroni L1355- SS Motor Length (in) Table 26: Comparison of four motors Impulse (N) Max Thrust (N) Launch Mass (lb) Empty Mass (lb) Projected Apogee (ft) Velocity Off the Rod (ft/s)

52 AeroTech L1390G AeroTech L1170FJ Table X above shows relevant data on the planned motor along with some possible alternatives. The motor for the launch vehicle shall meet the apogee requirement of 5280 ft, as well as have a velocity off the rod at least 52 ft/s. Since there is an airbrake subsystem in the rocket, the launch vehicle must overshoot the apogee by a margin to compensate for the additional drag that caused by the airbrake. The Cesaroni L1350 CS-P not only satisfies both requirements, reaching apogee of 5467 ft and a velocity off the rod of 66.9 ft/s, but also gives extra altitude for the airbrake subsystem to work with. 5.5 Recovery & Avionics Subsystem Avionic Components Altimeters Introduction Figure 15: Diagram of Launch and Recovery System Deployments 51

53 Using the dual deployment method to minimize maximum drift distance (in order to remain in the 2,500-ft radius recovery zone) requires two separate ejection charges: one charge to eject the drogue parachute occurring at apogee and another ejection charge to eject the main parachute occurring at a lower altitude after drogue deployment (500-ft). To deploy the main and drogue parachutes in their respective sequences, an altimeter is required to ignite the ejection charges. The following three are altimeters chosen for the selection process: MissileWorks RRC2+, MissileWorks RRC3 Sport, and the PerfectFlite StratoLoggerCF. Table 27: Altimeter Comparison Altimeter Vendor Model Cost Weight Features Integration Average Score Eggtimer Eggtimer Quark PerfectFlite Stratologger CF MissileWorks RCC MissileWorks RCC3 Sport Adept AltS2-50k Altus Metrum Easy Mega Altimeter #1 Figure 16: MissileWorks RRC2+ Barometric Altimeter 52

54 The MissileWorks RRC2+ altimeter is a barometric altimeter capable of operating at 40,000-ft above mean sea level, which satisfies the desired altitude of 5280-ft the launch vehicle is constructed to reach. It is programmed using a straightforward DIP switch configuration and eliminates the use for a user manual on the field, as all switch settings are printed legibly on the altimeter itself. The barometric altimeter transmits audible beeps to report the peak altitude directly from the device for the user to successfully retrieve data after launch. The unit is capable of reading its stored peak altitude from its previous flight and the current battery voltage. The RRC2+ is a simple and easy-to-use altimeter for beginners and experts alike at a very low cost, meeting most rocketry budgets. The RRC2+ unit is capable of dual deployment, providing a drogue event at apogee, a main event at an adjustable setting as low as 300-ft above ground level, and as well as a time delay drogue event as a back-up altimeter option. The device utilizes a precision pressure sensor and 24 bit delta sigma analog to digital converter to obtain an accurate measurement of the surrounding air pressure, before and after launch. The velocity data is also analyzed during flight so that mach dips, or sudden drops of altitude, are not incorrectly interpreted as apogee. All of the calculations are done inside the altimeter and no conversions or adjustments are necessary. The altimeter can be mounted in any direction on the sled, making the device omnidirectional. It also measures at 2.28 L x W x 0.5 H, being able to fit into a 24-mm diameter tube and weighing as low as 0.35 oz., making it a very lightweight and optimal instrument for the launch vehicle. The dimensions are optimal in design since it is compact enough to fit in most sub-scale and full-scale Avionics Bay designs. Although relatively inexpensive, the MissileWorks RRC2+ altimeter proportionately performs basic operations and telemetry. It does not report anything more than peak altitude, omitting useful data such as velocity. The altimeter heavily relies on audible beeps to report basic readings, lacking a variety of trajectory recordings such as ejection times and flight duration, and storing only one flight at a time. Altimeter #2 53

55 Figure 17: MissileWorks RRC3 Sport Barometric Altimeter The MissileWorks RRC3 (Rocket Recovery Controller 3) is a barometric altimeter with high-end features at a modest cost. The device is capable of operating up to 40,000-ft above mean sea level. This satisfies the desired altitude of 5280-ft the launch vehicle is designed to reach. The barometric altimeter transmits audible beeps of both the peak altitude and velocity directly from the device for the user to successfully retrieve data after launch. The RRC3 unit is also capable of executing dual-deployment with an adjustable launch trigger of 100-ft to 300-ft. The altimeter comes with two outputs; one each for the drogue parachute and the main parachute. It is a simple altimeter that can be used to perform right out of the box. In addition to its dual deployment capability, the RRC3 is a multi-flight recording unit, allowing the user to fly without the inconvenience to stop and download data between flights. In addition to altitude and velocity, the altimeter logs temperature and battery voltage. It captures up to 15 consecutive flights with approximately 28 minutes of data per flight, giving the user the opportunity to download and plot the data at the end of the day. This proves useful for the team to analyze a certain number of flights on and off the field. The RRC3 unit can also be used as a backup altimeter, providing the option for a backup mode and time delay for redundancy. The device can also support real-time data streaming capabilities in flight for telemetry with other custom control systems, allowing the user access to constant updated sources of data, such as time to apogee or current altitude. Along with transmitting audible beeps, the RRC3 unit has three sockets readily available for connecting to accessories such as a LCD screen display for better data retrieval. Where other dual deployment altimeters only have two pyro ports, the RRC3 has an additional auxiliary terminal block also capable of being programmed based on flight triggers for executing other missions. The RRC3 Sport altimeter measures 3.92 L x W x 0.5 H, being able to slide into a 24-mm tube and weighs at 0.59 oz. The dimensions are optimal in design fitting into most sub-scale and full-scale Avionics Bay designs. 54

56 The MissileWorks RRC3 Sport altimeter proves to be a high-end barometric altimeter, but is overqualified for the purposes of the mission at an unnecessarily higher cost. The unit provides various extra features that may prove to be redundant or unnecessary and includes additional methods to execute a certain function, such as the third pyro port or its two operation indicators. Altimeter #3 Figure 18: PerfectFlite StratoLogger CF Barometric Altimeter The PerfectFlite StratoLoggerCF altimeter is the Compact Footprint, more condensed version of the widely used and respectively popular PerfectFlite StratoLogger altimeter. The barometric altimeter is capable of operating up to 100,000-ft above mean sea level. This satisfies the required altitude as the launch vehicle is designed to reach a peak altitude of 5,280-ft. The StratoLoggerCF offers an easy to use altimeter that is capable of reporting peak altitude and maximum velocity after flight through an audible progression of beeps, making it optimal for comprehending information after flight as the user can listen for the audible beeps detailing the recorded flight data. The altimeter comes with two outputs that allow for dual deployment: one output is for the drogue parachute deployment and the other output is for the main parachute deployment. The altimeter also has a separate terminal block for the arming switch. This provides the competitor with the necessary outputs to satisfy a dual deployment recovery system. With the StratoLoggerCF altimeter, the main parachute deployment altitude is adjustable from 100-ft to 9,999-ft giving extreme flexibility to the user by having incremental values of 1 foot with the computer interface or 100 feet without the computer interface; the nine user-defined presets capable of being programmed and adjusted make on-the-field changes feasible and relatively simple should these changes be necessary when at the launch site. The StratoLoggerCF altimeter provides the user with a variety of useful flight data such as the altitude, temperature, and battery voltage. 55

57 With the altimeter and provided software, the user is equipped with a rate of 20 samples per second over 18 minutes and retains data for the most recent 16 flights; this is especially useful as a competitor because it allows for further analyzation of flight data for both sub-scale and fullscale launches. The altimeter also measures at 2.0"L x 0.84"W x 0.5"H and weighs just 0.38 oz. The dimensions are optimal in design since it is compact enough to fit in most sub-scale and fullscale Avionics Bay designs. According to PerfectFlite, this StratoLoggerCF Altimeter has a measurement precision of +/- 0.1% Reading + 1 foot; the advertised measurement precision is extremely accurate, making it an excellent data recording instrument. Along with the features this altimeter provides, multiple members of the team are knowledgeable in maneuvering through the software and altimeter; because of the reasons described, the PerfectFlite StratoLoggerCF Barometric Altimeter is an optimal altimeter and is the current lead component as the competition altimeter for the launch vehicle. GPS Units Introduction As part of NASA Student Launch, an electronic tracking device must be installed in the launch vehicle and must transmit the position of the tethered vehicle or any other independent sections of the launch vehicle. As to not interfere with the altimeters being used in the Avionics Bay upper tray compartment, the GPS is currently intended to be located in the Avionics Bay lower tray compartment, housed in its proprietary housing case that will be 3D printed from ABS thermoplastic. There following three are chosen for the selection process: Altus Metrum TeleMetrum and BigRedBee BRB900 Table 28: GPS comparison GPS Unit Vendor Model Cost Weight Dimensions Integration Average Score Transolve BeepX Eggtimer Eggfinder BigRedBee BRB900 TX/RX Altus Metrum TeleMetrum Altus Metrum TeleMega GPS Unit #1 56

58 Figure 19: Altus Metrum TeleMetrum GPS System At approximately 0.71-oz, the Altus Metrum TeleMetrum GPS System was one of the priority contenders for the multiple features that the model has integrated within it. The primary factor is the Dual Deployment Recovery System integrated with GPS tracking software which allows the user to integrate dual deployment directly with the GPS unit. This is highly ideal in that the allotted space required for a system of this size is approximately 1.068"L x 2.75"W x 0.62"H making it easy to relocate if necessary within the Avionics Bay. It is efficient on location but lacks efficiency on cost which is a major factor in choosing altimeters or GPS units. Lacking on cost efficiency is the only factor into the overall decision-making process. Another important but lesser factor is the uncertainty of using an unknown system when LBR has past experience with other GPS systems that have been proven to work. Furthermore, LBR plans to keep the redundancy in altimeters the same maker and model as it would decrease the amount of risk if the recovery electronics housed two different systems of deployment; it would cause a higher probability for error. With that in mind, it would not make sense to have two GPS systems although it would demonstrate redundancy in the GPS tracker portion of the electronics. Therefore, while lacking in cost efficiency but having the qualifications for dual deployment and GPS altogether, the Altus Metrum TeleMetrum GPS System remains as a contender as the possible GPS unit. 57

59 6GPS Unit #2 Figure 20: Big Red Bee BRB900 GPS Transmitter Figure 21: Big Red Bee BRB900 Receiver The Big Red Bee GPS Transmitter and Receiver functions as a sole GPS tracking system without any extra features integrated within it. The GPS Transmitter is based off of an out-of-the-box consumer system, reducing any risk for setup and operation; this allows the user to setup and install the GPS transmitter in the launch vehicle and use the receiver accordingly to successfully find the launch vehicle with ease. Although other GPS units mentioned may have similar functions, this unit acts solely as a GPS unit and is therefore efficient on cost which is a crucial factor in designing the overall launch vehicle along with its necessary components. The transmitter and receiver work up to six miles which is extremely favorable towards the user and Long Beach Rocketry overall as the maximum allotted recovery radius (2,500-ft) is minimal to the amount of distance that this unit covers. The transmitter is about 2.85 L x 1.25 W, making it easy to maneuver around the launch vehicle regarding location and placement if necessary. Because of cost efficiency and because this unit features Long Beach Rocketry s exact needs, the Big Red Bee BRB900 GPS System will be the competition GPS unit. Housing Trays and Casings Introduction 58

60 To attach the necessary bulkheads onto the Avionics Bay aft and forward ends, threaded rods must run through the internal structure of the bay and attach with the respective connections; to keep the threaded rods aligned and to hold the front and lower housing trays that the electronics will rest on, a casing is required in between the trays. The casing will also be the compartment for the battery placement. As said along with the casing, trays will sit on both sides of the apparatus to provide a flat platform for the GPS and altimeters to rest on. The plate will match the maximum dimensions of the apparatus to conserve space. There are three types of materials to consider when withstanding the stress of the launch vehicle during flight and recovery: metal, plastic and wood. Material Table 29: Housing Casing Material Comparison Housing Casing Material Cost Weight Strength Integration Average Score 3D Printed ABS Aluminum Plywood Metal Plate Aluminum 6106 would be the leading candidate for a metal plate since its properties would ensure the safety and durability of the avionics system. Aluminum 6106 possess a yield strength of roughly 240 MPa and a tensile strength of 290 MPa which is significantly overqualified for the purposes of the housing trays. Furthermore, an exact dimension of aluminum 6106 cannot be purchased which would prove to be considerably expensive as one would have to purchase an entire slab of commercial aluminum 6106 to use only a small portion of the material. 3D Printed Plate A 3D printed tray made of ABS plastic would be convenient for the customization of its dimensions, since an entire commercial slab of plastic would not have to be purchased. ABS plastic also carries moderately strong properties with a tensile strength of approximately 33 MPa. However, the material is more expensive than such material as wood. Wooden plate Similarly, to 3D printed ABS plastic, plywood provides an approximately 13.8 MPa yield strength and 31 MPa tensile strength which is potent enough to withstand the external forces of the flight. Plywood is also proportionately inexpensive and highly accessible for the team since a commercially stocked slab of plywood would not have to purchased. For the reasons described, 59

61 plywood will be the leading component for both the front and lower housing trays in the avionics bay. Battery Figure 22: Standard 9V Battery Batteries are necessary to feed power to the altimeters in the recovery system. Because each altimeter requires its own separate battery, two batteries are required. However, the dimensions and type of the battery must be considered in order to comply with the circuit and avionics system. The PerfectFlite Stratologger CF altimeters that must be powered, operates from 4V to 16V at 9V nominal. Since dimensions and weight are both sacrificed with greater voltage, a 9V battery balances all aspects of minimizing dimensions and weight while maximizing voltage. One 9V battery measures at 1 L x 0.68 W x 1.9 H and a AA battery at, D x H. Since a AA battery only outputs 1.2V, it would be required to operate in packs of at least four which would violate the boundaries of the avionics bay. The 9V alkaline battery serves optimal to power the altimeters while conserving space in the avionics bay. Power Switch Introduction Power switches are manual switches that contain an on and off position to power a circuit. These switches are necessary for the team to directly manually control the operation of the power supply that feeds the altimeters. This is especially useful as the user can save power for the purposes of the altimeter functions when necessary. By using the switch, the unit closes a continuous circuit containing a battery and altimeter, allowing the current to pass through along the circuit. 60

62 Table 30: Power Switch Trade Studies Power Switch Vendor Switch Cost Weight Accessibility Safety Average Score Schurter Rotary Electroswitch Rotary RAFI Emergency Stop NKK Switches Pushbutton Telodyne Relays Coaxial Figure 23: Circuit with Power Switch, Battery & Altimeter There are various models of switches that each utilize a specific method to close a circuit or turn on an operation system. Research has been done to determine a suitable power switch for the purposes of the recovery system and launch. Among the various types of power switches, the rotary power switch serves most practical for the purposes of the recovery system. Since the power switch will not only sit inside the avionics bay, but be accessible to the team to operate the recovery system, part of the switch must be exposed and will sit on the coupler piece. To minimize any potential debris during flight, the exposed switch should be as flat as possible and for this reason, a rotary switch is optimal. A rotary switch closes a circuit by rotation and only requires a simple twist to operate. 61

63 Figure 24: Schurter Rotary Power Switch The Schurter Rotary Power Switch is a rotary power switch that is more than suffice to fulfill the duties of the operating system with high voltage and low current. The unit works with the dimensions of the bay and with minimal weight and low cost. In addition, multiple members of the team are knowledgeable with the switch as it has been employed in previous years. For the reasons described, the Schurter Rotary Power Switch is an optimal power switch and is the current lead component as the competition power switch for the recovery subsystem. Avionics Bay Introduction The Avionics Bay is the compartment within the launch vehicle that houses all of the avionics electronics; it serves as the central coupler piece that attaches the Propulsion Bay and Payload Bay together. The purpose of having the avionics bay is to hold all the main electronics that will make the launch vehicle capable of having a dual recovery system as required by NASA regulations for the Student Launch Competition. In general, for most hobbyist, the avionics bay can house a single or dual deployment; for NASA Student Launch, the avionics bay being designed will have a dual deployment recovery system to minimize the maximum drift distance in order to stay within the 2,500-ft recovery radius. Single Versus Separate Avionics Bays Separate Avionics Bays For dual deployment recovery systems, there are two types of Avionics Bays available: Single or Separate. Using separate Avionic Bays for each event (main and drogue) allows freedom over the location of each bay and therefore parachute. This optimizes the safety of the launch vehicle 62

64 as it would offer the necessary protection of any of the crucial sections relative to the launch vehicle s design. Having Separate Avionics Bays also reduces the total amount of wiring in each Bay. Having a reduced amount of electronics and wiring in each bay increases the safety factor by significantly decreasing the risk of any electronic malfunctions during flight. While having separate bays seems favorable on certain systems, it does increase the overall weight as there would be a separate section to the launch vehicle. Increased weight could be beneficial or detrimental in certain cases as it would alter the center of gravity and center of pressure of the launch vehicle. Single Avionics Bays Using a single Avionics Bay is desirable in many cases, especially for the LBR s current design and mission. With better wiring techniques, it is possible to keep all the electronics for avionics on both the main and drogue side together without causing any risk for confusion and possible malfunction. The Payload Bay is still protected by having the main open between the Payload Bay aft end and the Avionics Bay forward end. This will cause the Propulsion Bay to land first instead of the Payload Bay. Having a single Avionics Bay also reduces the need for another section on the launch vehicle, maintaining efficiency on weight. Because there is only one Avionics Bay, there will be limitations on placement which could hinder the success of the launch vehicle s mission depending on what that mission is. Another area that is lacking is that since all electronics and wirings are together, there still is a possible risk for misplacement or miswiring which would be at the fault of the user leading the Recoveries and Avionics systems. Evidently, it will be up to the user to understand how to void this area of risk by understanding ease of access with electronics and wiring. Therefore, depending on the mission and purpose of the launch vehicle, there may be different aspects with having separate Avionics Bays that is more favorable than using a single Avionics bay. Likewise, there are aspects of having a Single Avionics Bay that is more favorable in other cases as well. For LBR s case, having a Single Avionics Bay is the most optimal for the mission and current launch vehicle design. With three main compartments to the launch vehicle including the Avionics Bay, the Payload Bay will utilize the Nose Cone, applying said method earlier where the main will deploy between the Avionics Bay forward end and the Propulsion Bay aft end. Trade space The Avionics Bay compartment is located between the Propulsion Bay and Payload and Nose Cone Bay. The purpose of having the Avionics Bay between the propulsion and payload bay is to create a workspace of ease for setting up dual deployment. By having the Avionics Bay between the Propulsion and Nose Cone bay, there is sense of ease for setting up the ejection charges with the wire connections. Having the avionics electronics in the middle allows for said electronics to connect to the main and drogue ejection charges with ease. From past launches and prior testing, the GPS has not experienced any malfunctions due to specific location and used materials that would potentially block or weaken the GPS signal. Using fiberglass airframes for the Student Launch year does not pose any issues for the electronics being used currently; however, if any carbon fiber material was to be used around the avionics electronics 63

65 such as for the airframes, then the location of the GPS would have to be placed at another location such as the nose cone. This is because the carbon fiber material can potentially act as a radio frequency shield, blocking or weakening the GPS signal which would hinder the success of the mission in terms of the recovery process. Thus, the current design poses for a fiberglass airframe surrounding the avionics bay so that the GPS may be within the same compartment as the rest of the avionics electronics. External and Internal Coupler Piece Properties As stated, the Avionics Bay holds all avionics electronics and acts as the central piece that connects the Propulsion Bay and the Payload Bay together; these avionics related components will be housed in a 6 fiberglass coupler piece. There will be two plywood rectangular plates that will serve as the front and lower trays to hold specific avionics components. The front plywood tray will hold the two (primary and backup) PerfectFlite StratologgerCF barometric altimeters and wire connections that will deploy the ejection charges and parachutes. The lower plywood tray will carry the BRB900 GPS unit with the 3D printed GPS housing case. In between both plywood trays, there will be a 3D printed centered housing case which will allow two 5/16-18 threaded rods to pass through and connect to the aluminum bulkheads on the forward and aft ends of the avionics bay. Both the forward and aft ends of the coupler will be secured with two manufactured 6061 standard aluminum bulkheads and will be fastened with nuts and washers. Each aluminum bulkhead will have a U-Bolt integrated within it to allow shock cord attachments and tethering. Recovery Subsystem Introduction The launch vehicle will be designed for a dual deployment recovery system. The purpose of having dual deployment over single deployment is to minimize the drift distance and stay within NASA s required 2,500-ft recovery radius. The system uses various parts and tetherings connected with the Avionics Bay to successful discharge the respective parachutes at apogee and then at 500-ft. This method of recovery is fundamental in dramatically decreasing the drift distance and ensures that Long Beach Rocketry meets NASA guidelines in the recovery process. Ejection System Introduction During launch and descent, the launch vehicle requires a deployment system to successfully eject the drogue parachute and the main parachute at the correct altitudes. In order to There will be two deployment system methods to initiate the recovery sequence: CO2 ejection system or Black Powder ejection. CO2 Ejection System 64

66 Using a CO2 Ejection System is plausible for the purpose of our launch vehicle regarding recovery. A method of pursuing CO2 ejection is to use the Peregrine CO2 system, a commercially available deployment system that utilizes the puncturing of a CO2 canister to create a pressurized volume of CO2 gas forces that would act against the avionics bay front and aft ends, releasing into the parachute compartments and separating the shear pins, deploying the recovery system successfully. The apex section of CO2 canisters will be punctured after reaching an altitude as specified by the user. The CO2 gasses are cold which poses as a major advantage of the CO2 ejection system as the parachutes and shock cords will not suffer any heated corrosive damage. The disadvantage of the CO2 ejection is that it requires a lot of components, which contributes to a high mass value; this detail is crucial as any attributed mass will affect the center of gravity as well as the center of pressure and therefore, must be accounted when designing the overall launch vehicle. Black Powder Ejection System The Black Powder Ejection System is another plausible method of deployment and at the moment is the most favorable method due to it requiring less components, making it efficient on weight and overall construction. Black powder ejection charges may be considered a staple in the rocketry industry when deploying a parachute system. Black powder will require an electronic match (commercial available) that will ignite based on a signal received by the altimeters used. Once the desired height has been reached the electronic matches will ignite, causing combustion and creating hot gasses that will form air pressures within the airframes that act against the forward and aft ends of the Avionics Bay, shearing the pins and pushing out the parachutes. Using Black powder charges are favorable for the purpose of NASA Student Launch due to it having a low mass system. The system requires minimal components compared to CO2 deployment when setting up and throughout operation; another advantage of this method is the relatively low cost. The system is eminently effective and reliable but does have its disadvantages. After the electronic match ignites the black powder, the powder will instantaneously combust and excrete high temperature expanding gases. The recovery components such as the tethering cords (shock cords) or parachutes are at risk if not properly protected and shielded with their respective blankets and covers. If not properly protected the hot expanding gases from the igniting black powder can potentially cause the shock cord lines to fray as well as produce burn holes in the parachutes; ultimately, this would render the parachutes or shock cords defective; however, these issues may be absolved through protective kevlar blankets and shock cord protective sleeves. Thus, for the purpose of NASA Student Launch and because of the cost and weight efficiency, black powder ejection charges will be used for the parachute deployment system of the launch vehicle. Preliminary Analysis on Parachute Sizing The projected weight of the launch vehicle is 42-lbs; after launch vehicle has reached apogee, the composites in the motor will be expelled along with the drogue chute deploying, creating a new weight during descent of approximately lbs. After the second event, the main chute 65

67 deployment will create a new weight of approximately lbs. In selecting parachutes and calculating projected descent velocities, the new weights during descent and at each event must be accounted for. Selecting the required parachute requires knowing the descent velocities at different parachute diameters. In order to calculate descent velocity, the following parameters: mass of launch vehicle, parachute diameter, parachute drag coefficient, and the air density taken at sea level. The drogue parachute will be the 20 Fruity Chutes TARC Low and Mid Power Parachute which has a drag coefficient of 1.5. The main parachute will be the 84 Fruity Chutes Iris Ultra Standard Parachute which has a drag coefficient of The equation for descent velocity is given by: Descent Velocity = 8mg πρc d D 2 With the chosen parachute models, five drogue parachutes and six main parachutes with varied diameter sizes were chosen for calculations of descent velocities. Calculating the launch vehicle descent velocity under differing parachute sizes of both the main and drogue is essential to choosing the necessary parachute for LBR s launch vehicle as it will illustrate the optimal descent required to stay within the 2,500-ft recovery radius as well as not exceed the required maximum kinetic energy load of 75 lb-ft upon landing. Shown below are two graphs depicting those calculations of descent velocities under different parachute sizing as a function of weight. Weight values greater than or less than 33.8-lbs were chosen as contingency mass. 66

68 Figure 25: Descent Velocities Versus Weight with Various Drogue Parachute Sizes Figure 26: Descent Velocities Versus Weight with Various Main Parachute Sizes Based on multiple sources of experienced professionals in the rocketry industry, a good rule of thumb for projects similar to Long Beach Rocketry s launch vehicle is to keep the drogue parachute descent speed between 70-ft/s and 95-ft/s; the main parachute descent velocity should also be kept under 25-ft/s. Based on the calculations, the points highlighted in yellow diamonds on the graph depict the parachute sizes with the most optimal descent velocities for the following conditions set. Therefore, the drogue parachute will be the 20 Fruity Chutes TARC Low and Mid Power Parachute and the main parachute will be the 84 Fruity Chutes Iris Ultra Standard Parachute. The drogue parachute has an optimal descent velocity of ft/s and the main parachute has an optimal descent velocity of ft/s. Type of Parachute Drogue Parachute Table 31: Parachute Sizing Parachute Size and Model 20" FC TARC Low and Mid Power Parachute Relative Descent Velocity (fps) Main Parachute 84" FC Iris Ultra Standard Parachute

69 Recovery System Figure 27: Recovery Layout Once the launch vehicle reaches its highest point of altitude (apogee), it needs a recovery system to land safely without drifting an unreasonably protracted distance. Based on prior research and multiple forums with experienced professionals in this industry, the dual deployment method recovery system is the most optimal form of recovery when compared to single deployment. This is mainly due to dual deployment s exceptional capability of reducing the maximum drift distance during the recovery process; therefore, the launch vehicle will utilize a dual deployment recovery system. The drogue parachute will deploy at 5,280-ft once the ejection charges ignite and separate the Propulsion Bay from the Avionics Bay aft end. The main parachute will follow afterwards and will deploy at 500-ft; it will deploy once the secondary ejection charges ignite and separate the Payload Bay from the Avionics Bay forward end. The entirety of the launch vehicle will be tethered together with nylon shock cords in which the shock cords will be protected from the combustion and heated gasses of the black powder via protector sleeves. Shock cords will have a minimum requirement of 1400-lb breaking strength to qualify its 68

70 usefulness and reliability as the tethering cords. The black powder ejection charges will shear the four ⅛ shear pins connecting the Avionics Bay to the Propulsion and Payload Bay. To ensure maximum safety during recovery, both parachutes will be housed in a deployment bag that is given additional protection through protective fireproof blankets. 5.6 Mission Performance Predictions Applicable Equations To successfully launch a rocket of any size, it is important to have an understanding of how to calculate center of gravity, center of pressure, static stability, and peak altitude of the rocket. The static stability is a dimensionless number found by dividing the distance between the center of gravity and the center of pressure by the body tube diameter. 1. S = x cp x cg d where x cp is the center of pressure and x cg is the center of gravity. The center of pressure can be measured from the nose cone tip by using 2. x cp = (C N) n X n +(C N ) f X f (C N ) n +(C N ) f where (C N ) n is the ogive nose cone center of pressure, X N is calculated by 3. X n = 0.466L n Where L n is the nose cone length. (C N ) f is the fin center of pressure coefficient can computed by 4. (C N ) f = (1 + R S+R ) S 4n( d ) ( 2l a+b )2 where S is the radius of the body between the fins, S is the fin semi-span, and n is the number of fins, l is the length of the fin mid-chord line, a is the fin root chord length, and b is the fin tip chords length. And finally, X f is calculated by 5. X f = x f + m(a+2b) 3(a+b) + 1 ab (a + b ) 6 a+b where x f is the distance from the nose cone tip to the front edge of the fin root, m is the distance between the fin root and fin tip. The center of gravity is computed by 69

71 6. x cg = d nw n +d r w r +d b w b +d e w e +d f w f W where W is the sum of the whole rocket weight, d is the distance between a specific center of gravity (nose cone, recovery, body, engine, and fins, respectively) the the aft end of the rocket. The peak altitude is found through the sum of the boost phase while the motor is burning and the coast phase from the motor burn-out to peak altitude. 7. PA = y b + y c To find the altitude of the boost phase, y b, the average mass is first calculated using 8. m a = m s + m e m p 2 where m s is the structural mass of the rocket, m e is the motor mass, and m p is the propellant mass. The aerodynamic drag coefficient, k, is 9. k = 1 2 ρc da where ρ is the air density, C d is the drag coefficient, and A is the rocket cross-sectional area. Knowing the average mass, m, and the aerodynamic drag coefficient, k, the burnout velocity coefficient, q b, is then computed by 10. q b = T m ag k where T is the motor thrust, and g is the gravitational constant. Equations 2, 3 and 4 are then used to compute the burnout velocity delay coefficient, x b, using Then the motor burn time, t, is calculated using 11. x b = 2kq m a 12. t = I T Where I is motor impulse, and T is the motor thrust. Equations 4,5, and 6 are then used to calculate the burnout velocity, v b, using 70

72 13. v b = q b 1 e x b t The altitude at burnout, y b, can finally be computed by 14. y b = m a 2k 1+e e b t ln 2 (T m ag kv b ) T m a g After the boost phase altitude is calculated, the coast phase altitude can be determined. At burnout, the new mass during the coast phase is 15. m c = m s + m e m p With the new coasting mass comes with the coasting velocity coefficient, q c, and the coasting velocity delay coefficient, x c 16. q c = T m cg k 17. x c = 2kq c m c With equations 10 and 11, the coasting velocity, v c, can be computed by 18. v c = q c 1 e x ct 1+e x ct The coasting phase altitude can then be calculated by using 19. y c = m c ln 2 2k (m cg+kv c m c g With the burnout altitude and the coasting altitude, the peak altitude can be determined using equation 7 above. ) 71

73 Flight Simulations Figure 28: Motor thrust curve for an Cesaroni L1350-CS The motor thrust curve for an Cesaroni L1350-CS motor from OpenRocket can be seen above in Figure X. The motor burnout occurs at 3.17 seconds. Figure 29: Graph of launch Vehicle motion vs. time with 0 mph wind conditions 72

74 Open Rocket was used to simulate launches at various wind speeds from 0 to 20 mph wind in increments of 5, using an Cesaroni L1350-CS motor. A sample simulation can be seen above in Figure X, these simulations were used to predict apogee at various wind speeds, with a rocket mass of lbs with a loaded motor. Table 32: Projected Apogee at Different Wind Speeds Using Openrocket Wind Speed (mph) Projected Apogee (ft) Stability Figure 30: Openrocket side view of launch vehicle The stability of the launch vehicle was projected by Openrocket to be 2.65 calibers with a center of gravity of inches from the nose cone and a center of pressure inches from the nose cone as shown in Figure X above. Openrocket was cross referenced for redundancy using the following equations. CP CG Static Stability = D where S is the stability in calibers, CP is the length from the top of the nose cone to the center of pressure, CG is the length from the top of the nose cone to the center of gravity, and D is the diameter of the airframe. Calculation of launch vehicle stability inches inches 6.17 inches = 2.67 calibers 73

75 This stability for our rocket is the result of careful consideration in weight distribution and fin design. LBR went for a much stable rocket in order to prevent the mishaps of last year rocket launching off course due to a low of a stability. With a higher stability this year, that should no longer be an issue. The stability is also sufficient for NASA s minimum stability requirement of LBR s stability is perfect in maintaining a steady launch while not becoming over stable where weather-cocking becomes a bigger issue. Weather-cocking is the result of the a too high stability leaving the launch vehicle more susceptible to the movements of the wind. This is most commonly seen with a stability over 3.00 calibers which LBR is safely under. Kinetic Energy The launch vehicle will separate at various stages during its flight and will have multiple parachutes to facilitate a safe decent. It is required for all the components of the launch vehicle land with no more than 75 ft-lb of kinetic energy. Using the equation below, the kinetic energy of each section can be calculated. KE = 1 2 mv2 Sample Calculation for descent velocity and kinetic energy of payload bay: v = 8mg πρc d D 2 KE = = 68.4 lb ft Table 33: Kinetic Energy for each Sections Kinetic Energy for Each Independent Sections Upon Landing Section Weight (lb) Mass (slugs) Descent Velocity (ft/s) Kinetic Energy (lb-ft) Payload Bay Avionics Bay (After Event 2) Propulsion Bay LBR has noted that the landing kinetic energy for the payload is alarmingly large. The payload bay is housing all the electronics for the RDM system and rover, any misalignment of the electronics after landing can cause a mission failure. However, LBR was required to choose 74

76 drogue descent velocities that would prevent sections of the launch vehicle from drifting out of the 2500 ft limit NASA requires for launch vehicle sections. These parachutes were chosen with that criteria in mind. As a result, LBR will explore options and do more intense drop testing of the payload system to simulate this very kinetic energy in order RDM will still function after landing. Drift Calculations To ensure the launch vehicle recovers within the 2,500-ft recovery radius set by NASA guidelines, the main deployment altitude of the main parachute must be optimized to decrease the maximum drift distance during recovery. The drift distance must be calculated for five different wind conditions: 0-mph wind, 5-mph wind, 10-mph wind, 15-mph wind, and 20-mph wind. The calculations shown below are based on the assumption that the launch vehicle is directly above the launch rail during apogee and that the drift speed of the rocket is the same as the wind speed; with these conditions set, the following equation is used to calculate drift distance: Drift Distance = (t d S) + (t m S) where t d = Drogue Descent Time, t m =Main Descent Time, and S=Wind Speed Table 34: Drift Calculation for Various Wind Speed Drift Calculations Using 84" FC Iris Ultra Standard Parachute Main with 20" FC TARC Drogue Wind Speed (mph) Wind Speed (ft/s) Drogue Drift (ft) Main Drift (ft) Total Drift (ft) Once the launch vehicle reaches apogee, the drogue deployment system will eject the drogue parachute. In the most optimal case where there is 5-mph or less wind speeds, the maximum total 75

77 drift during the first event to the second event will be approximately 377-ft. At 500-ft the main parachute deployment system will eject the main parachute; in the most optimal case with 5-mph speeds or lower, the maximum total drift once the main ejects and creates an overall descent velocity of ft/s is approximately 206-ft. Altogether, this creates a maximum total drift distance of about 583-ft which is highly favorable as it remains in the 2,500-ft recovery radius. In another case, at 20-mph wind speeds, the drift speeds during both events is significantly increased. With the drogue deployed at a descent velocity of about ft/s, the maximum drift speed is about 1,509-ft. After the main is deployed and creates an overall descent velocity of ft/s, the maximum drift distance then becomes 824-ft. Altogether, this creates a maximum drift distance of about 2,332-ft during descent. Even in the worst-case scenario with 20-mph wind speeds, the total drift distance given by the launch vehicle s weight and parachute selection are still favorable regarding the maximum drift distance as the launch vehicle remains within the 2,500-ft recovery radius set by NASA Student Launch guidelines. Figure 31 Openrocket simulation of drift in 5mph wind The drift calculation at 5mph wind is cross referenced in figure 31 above, which shows a drift distance of approximately 900 feet. This is contrast to our previously calculated drift distance of 582 feet leaving a 318 feet difference. This a result of the previous assumptions that each section of the rocket will descend at the same velocity and that the rocket s flight path will go straight up unaffected by the wind. Making Openrocket s calculations a more accurate result as it does not take these assumptions into account. 76

78 5.7 Launch Vehicle Team-Derived Requirements Team-Derived Requirements Table 35: Launch Vehicle Team Derived Requirements Verification Method Verification plan Altimeters unable to deploy ejection charges. Avionics bay will be easily accessible as well as precisely organized. Drogue or main parachute becomes tangled during flight. Verify altimeters are unable to deploy ejection chargers Inspection Testing Two altimeters will be used for redundancy. Each altimeter will be individually tested in a vacuum chamber to simulate launch and will carry LED lighting that turns on at apogee and at 600-ft to confirm successful operation of the ejection charges. Each electronic match will be physically inspected before every launch for any possible defects. All wiring will be checked prior to launch to ensure there are no loose wires. The avionics bay will house multiple electronics that will be physically placed after being modeled with computer-aided design software. All wiring will be tightly sealed and wrapped to ensure maximum organization. Enough wiring length and airspace will be provided to ensure easy access upon removal of the avionics bay before and after launch. The drogue and main parachute will each be meticulously inspected to ensure that the chutes do not tangle during flight. Each drogue and main chute will be inspected thoroughly and folded with precision to give the highest probability for successful deployment. 77

79 Drogue or main parachute is unable to deploy during flight. Recovery electronics become nonfunctional due to defective wiring. Testing Testing Numerous ground ejection testing will be executed to ensure that the drogue and main parachutes successfully deploy once the ejection charges fire. All wiring ends will be soldered to ensure a secure connection between each input and output sources. Wiring connectors and adapters will be used where applicable to ensure easy accessibility and stable connections. Avionics bay bulkhead caps will can withstand the shock cord forces. Testing The caps will consist of fiberglass bulkheads which will take the load. The launch vehicle will reach apogee while carrying a scientific payload. Testing Appropriate motors for the mass of the launch vehicle with scientific payloads will be selected. The launch vehicle will successfully integrate the rover deploying mechanism and the autonomous rover. Testing The diameter of the launch vehicle will be increased to provide space for the RDM and the autonomous rover. Transparent section of the launch vehicle will not hinder the structural stress resistance of the rocket The launch vehicle must house electronics and parachutes. Analysis and Testing Testing A transparent material at least as strong as fiberglass will be selected. Launch vehicle tubes will be large enough to house avionics bays for the electronics and enough space for the parachutes and shroud lines. 78

80 The launch vehicle must be able to separate and deploy parachutes with black powder charges. The launch vehicle must provide adequate initial thrust to maintain stability Motor tube must be able to securely hold motor in place. Testing Testing Testing There will be adequate black powder to separate the sections but not enough to cause damage. Approximately three grams of black powder will be used to separate each section, and ground testing will be used to verify this design. Calculations will be used to determine an appropriate motor to give a stable thrust-to-weight ratio and test launches will confirm this. The motor mount is designed to have a motor retainer that will hold the motor in place Fin alignment must be symmetrical, so rocket path is accurate. Testing A fin alignment mechanism built by LBR will be used to align the fins. Bulkheads must securely hold shock cords and be able to withstand deployment forces. Testing Bulkheads will be epoxied and screwed in place and eyebolts will be secured to them to hold the shock cords. 79

81 Section 6: Payload Criteria Rover Deploying Mechanism (RDM) 6.1 Overview Rover Deploying Mechanism s objective is to separate the nose cone from the launch vehicle after launch, and then allow the rover to deploy from the inside of the launch vehicle. RDM will utilize rotary to linear mechanism to move the rover and nose cone out and deploy the rover. RDM Mission Requirements 1. LBR team will launch a rocket which carries a rover that fits within the constraints of the payload bay. 2. The deployment process will be remotely triggered following a safe landing of the rocket. 6.2 System Summary Figure 32: Tilted view of chosen RDM option that demonstrates the separation of rocket components The Rover Deployment Mechanism (RDM) will deploy the rover from the internal structure of the launch vehicle. The rover deployment mechanism must comply with the constraints of the rover design to allow the most efficient use of space while remaining within the weight restrictions. Since the rocket will be landing in an arbitrary position, the deployment mechanism must can work in all directions and be powerful enough to open regardless of the terrain or obstacles. Long Beach Rocketry team has considered various options to extract the rover, and has decided to go with an electro-mechanical system of disconnecting rocket components to expose and extract the rover from its internal structure. This system must can secure the rocket frame together during flight and rugged enough to withstand the impact of the rocket landing. Since the rover will be housed in the upper payload bay of the launch vehicle, RDM will extract the rover by disconnecting parts of the exterior frame of the rocket. With this general design principle in mind, LBR has considered multiple design solutions with different components. The commonality between all discussed solutions however, is a system of threaded rods and nuts secured to the bulkheads. The bulkheads with attached nuts are secured to exterior frame components. Motors in the airframe spin the threaded rods, and the bulkhead nuts along with 80

82 their attached frame component will move linearly along the rods. Specifics of the design however vary with the different design options. Two design options have been narrowed down for the opening of the nose cone. For Option 1, the team considered separating the rocket into three components: airframe, coupler, and nose cone. Three threaded rods in a triangular configuration will run through all three components. An unthreaded rod will run through the center. The rover will have a hole in the center, and it will sit in the coupler with the unthreaded rod running through the center. Aluminum bulkheads with threaded nuts built in will be secured to the coupler and nose cone so that when the rods are twisted, these components can move linearly along the rods. Two bulkheads within the airframe will create a compartment for three high-torque motors. The motors will spin and first move the nose cone and coupler away from the airframe. The rover will follow the coupler since it is contained within. The nose cone will fall off the threaded rods so that it is completely separated from the launch vehicle. The coupler will approach the end of the threaded rod, and as it does, the rover will fall off the center unthreaded rod. The back end of the rover will catch the tip of the unthreaded rod, and as the coupler is retracted, the rover will be left behind on the ground which will allow it to be deployed. Option 2 is a simplified approach to the first option, with the use of a single threaded rod running through the center of the airframe and rover. Additionally, an unthreaded rod will also be secured parallel to the threaded rod, but its job will be to prevent the rover from spinning rather than to push the rocket components apart. A threaded nut will be attached to each bulkhead, as well as the rover. When the middle-threaded rod rotates, it will move both the bulkheads with attached rocket components as well as the rover which will allow the rover to be deployed. LBR plans on successfully deploying the rover from the airframe and ensure that the motor has enough torque to detach the nose cone from the airframe to break the shear pins after the launch. Option 1 Using three high torque brushed DC motors, three 20 long threaded rods with a diameter of 0.25 and thread sizes of ¼ -20, one 20 unthreaded rod with a diameter of 0.25 and three bulkheads with 6 diameter and 0.25 thickness, LBR, would connect the three DC motor shafts and the threaded rods using three 1.04 couplers, one for each motor. The coupler would rest behind one 6 bulkhead with threaded nuts attached before the middle and nose cone bulkhead. Using an H-Bridge, an Arduino board, a 11.1 V battery, and a controller, the H-Bridge would read input from the digital pins and cause motors to spin the threaded rods when a switch is flipped on the controller. The threaded rods spin until they detached from the nuts inside the nose cone bulkhead. After, the coupler and rover are left on the threaded rod. The threaded rod continues to spin until the end of the coupler is away from the end of the threaded rods. The rover, which has a linear bearing inside of it, will be free to move along the rod and will eventually be pushed off the rod by the coupler. After, the motors will spin backwards causing the coupler to retract. This means that the rover will sit only on the unthreaded rod and not inside the coupler, allowing it to fall off the rod when the coupler is removed. To ensure that during 81

83 flight the components will not detach, while closed, the motor will spin the threaded rods backwards to tightly secure them. Figure 33: Schematics for RDM Option 1 82

84 Figure 34: Front view RDM option 1 Figure 35: Tilted view RDM option 1 Table 36: Threaded Rods: Option 1 Requires Three Threaded Rods Pros Cons Three rods ensure that the component load is shared on top of three rods instead of one. Three rods cause more possible points of failure because the success depends on all rods working correctly. Any deformations in the rod can affect how the nuts, attached to the bulkhead, travel along the rod. If one nut fails to travel along the rod, the whole system fails. 83

85 Table 37: Motor: Option 1 requires three threaded rods that will connect to three DC motors Pros Cons The torque required to break the force of shear pins is distributed over the three motors. There is more opportunity of failure because there are three motors. If one motor fails, the whole RDM system fails. Table 38: Rocket Divisions: Option 1 requires that the rocket will be divided into three sections Pros Cons This design allows for the use of three threaded rods. There are two points of detachment, so if either of them fail, the rover will not be able to leave the payload bay. Option 2 For Option 2, LBR will only separate the rocket into two components: the airframe and the nose cone. A single 20 threaded rod with a ¼ -20 thread size will run through the center of the two components and two 20 unthreaded rods on opposite sides of the airframe. A high-torque brushed DC motor with a built in rotary encoder will connect to the threaded rod using a inch coupler. The base of the motor will be secured to an aluminum bulkhead that is 0.25 inches thick and has a diameter of 6 inches. Another bulkhead of the same dimensions with a 0.25 hole through the center will block the motor from the rest of the airframe. The rover will have a hole in the center and a threaded nut attached to it that will spin along the threaded rod. A potential problem the LBR team noticed was that the rover had a possibility of spinning with the rod rather than moving along the rod. To solve this problem, the team added an unthreaded with the same dimensions as the threaded rod to the design. It will run parallel to the threaded rod and rest against the rover to prevent it from spinning. This contact force will translate the spinning of the rover into linear motion along the rod. The final aluminum bulkhead will attach to the nose cone. A nut and washer will secure to this third bulkhead as well so that it has the capability of translating along the threaded rod. Shear pins will attach the nose cone to the bulkhead so that during flight, they do not separate. When the motor spins, the nutted rover and the nutted bulkhead will move along the threaded rod. The motor will supply enough torque to exceed the 250-pound force required to break shear pins and free the nose cone. The nose cone will eventually unscrew off the rod. The rover will follow until it reaches the end of the rod. The rover will keep spinning along the rod until it pushes the loose nose cone out of the way. Then it will fall to the ground and drive. 84

86 Figure 36: Schematics for RDM Option 2 Figure 37: Front view of RDM Option 2 85

87 Figure 38: Expanded view of RDM option 2 Table 39: Threaded Rods: Option 2 Requires a Single Threaded Rod Pros Cons One threaded rod minimizes points of failure. There is more space within the airframe if there is only a threaded and unthreaded rod. One threaded rod means that the system load is placed solely on the single motor and rod pair. Table 40: Motor: Option 2 Requires One Single DC Motor with a Built in Rotary Encoder Pros Cons One motor requires less coding and synchronization making deployment simpler. One motor and battery reduce weight of the RDM bay. This motor has a rotary encoder included, so information about the motion of the motor will be fed back to the user. If problems arise, the user will know. Less motors and smaller battery size yields more stress placed on the components. The motor must supply a large amount of torque to break the shear pins. Table 41: Rocket Divisions Option 2 requires that rocket will be divided into two sections Pros Cons There are less opportunities for the rocket to fail since there is only one point of separation. The motor must have a lot of torque, so it can overcome the force of the shear pins. 86

88 Shear pins can be used, so programming becomes simpler. The motors no longer need to spin backwards during flight. Comparison Between Option 1 and Option 2 LBR considered two main options for the design of the rover deployment mechanism. Below is a numerical comparison of the efficient use of space, weight, complexity, ease of coding, torque capability and price between Option 1 and Option 2 since these are important factors in the choice of design. Each characteristic was measured on a scale of 1 to 5, 1 being poor and 5 being excellent. Table 42: Comparison between Two Options Rover Deployment Mechanism RDM Mission Requirements Option 1 Option 2 Launch a rocket which carries a rover that fits within the constraints of the payload bay. Yes Yes The deployment process will be remotely triggered following a safe landing of the rocket. Yes Yes Comparison of Features Efficient Use of Space 1 5 Weight 3 4 Complexity 2 3 Coding 2 4 Torque Capabilities 5 3 Price 2 5 Total

89 After evaluating the pros and cons of each rover deployment mechanism, Option 2 was found to be the best option. This design utilizes less motors and threaded and unthreaded rods, yielding less points of failure than Option 1 while still supplying sufficient torque to successfully deploy the rover from the rocket. The design had to be light and simple while keeping a high quality of functionality with few points of failure. Reduced motors ensures less weight and requires a smaller battery and fewer technical components. The decreased amount of motors also has the advantage of less required coding and synchronization that would be otherwise be needed through Option 1 s method of rover deployment. The one threaded rod of Option 2 decreases the weight of the RDM in contrast to three threaded rods with one unthreaded rod through the center. With these changes the weight of the rover deployment mechanism can be reduced while still ensuring the same level of practicality and successful deployment. 6.3 Control The RDM must be remotely activated once the launch vehicle has safely landed. The wireless communication between the team and the rocket is essential in order to begin the rover deployment process, so a reliable connection between the transmitter and receiver is necessary. There are many ways to send a signal to communicate with the microcontroller but due to the high risk of losing signal between the systems the team has limited the choice of communication systems between the XBee RF module or a hobby-grade FlySky 2.4 GHz radio transmitter. Due to the familiarity, high range, and small size of the transmitter/receiver combination, the team has decided to use the FlySky radio transmitter to wirelessly interface with the onboard electronics. In order to begin the deployment of the rover, a switch on the transmitter will be triggered which will send a signal to the receiver on board the launch vehicle. The receiver sends a pulse width modulated signal which will be an input to the microcontroller, which will begin the drive of the motor. The motor used to drive the threaded rod is a high-torque brushed DC motor with encoder, so the motor will be connected to an H-Bridge motor driver which will allow the microcontroller to drive the motor in both directions. In the circumstance that the RDM gets stuck and begins to stall, the encoder will recognize the change in movement of the motor shaft and will notify the Arduino to stop and retract the motor for a predefined period of time. This will allow for any object interfering with the deployment mechanism to be omitted and it can continue extracting the rover. 88

90 Table 43: Pros and Cons of Signal Transmitters Pros Cons DIGI XBEE S There is only one button that needs to be pushed to start the RDM Limited transmitting range Needs a shield which takes up more space Complicated to program because of all the components FlySky 2.4 GHz Radio Transmitter/Receiver System Easy to program Up to 6 Channels that can be utilized for control The receiver takes up little space The remote control is very large and takes up a lot of space 6.4 Electronics The electronic components that make up the RDM is consist of geared motor with encoder, H- Bridge motor driver, Arduino Nano Microcontroller, 2.4GHz Digital Transmitter and Receiver Radio System, and an 11.1 V LiPo Battery. To be able to initiate the RDM process remotely, LBR team will use the 2.4GHz Digital Transmitter and Receiver Radio System. Using an Arduino Nano Microcontroller, the team will validate the values coming from the digital receiver and program validation statements that will determine whether the H-Bridge will get an input that would make the motor turns on and off or rotates clockwise/counter-clockwise. The team will also utilize the rotary encoder attached to the motor for feedback regarding the real time speed and distance traveled by the rover attached to the threaded rod. For instance, when load becomes higher, the speed of the motor will decrease despite at 100% duty cycle. The encoder reads this change in speed, then the Arduino reduces the speed of the motor at appropriate duty cycle to gain higher torque and avoid motor burnout. 89

91 Figure 39: Electric Schematic of RDM Option 2 The chosen design of Option 2 provides balance of weight and component number. Below is a table of all the components that make up the RDM and their expected masses. Table 44: Estimated Mass of the RDM Estimated Mass of Components (oz) Motor 4.48 Arduino Nano Microcontroller 0.21 L298N H-Bridge GHz Digital Receiver V LiPo Battery 7.94 Rod/ Coupler Pair 6.24 Threaded Rod 2.89 Unthreaded Rod Bulkheads

92 Total Preliminary Interfaces The payload will be housed within the internal structure of the launch vehicle. The connection between the rover and the launch vehicle is the threaded rod that runs through the center of the rover and the launch vehicle. The rover contains a threaded nut that will secure it to the launch vehicle until it is screwed off. Additionally, the unthreaded rod will make contact with the rover and prevent it from spinning while also helping to secure the rover. In order to connect the threaded rod to the motor, LBR used 2 High-Parallel-Misalignment Flexible Shaft Coupling, one had a ¼ diameter and the other was a 4 mm diameter, and a durometer 98A Spider for 3/4" OD for Clamping Vibration-Damping Precision Flexible Shaft Coupling. During flight the bulkheads will keep the rocket deployment mechanics from moving and ensure their ability to operate after the landing of the rocket. The bulkhead in the nose cone has a nut trapped inside to connect the nose cone to the threaded rod, which will also provide stability and keep the RDM intact. The motors will be securely screwed into the base bulkhead, and the other bulkhead will seal off the motor compartment and provide stability. Shear pins will secure the nose cone to the airframe during launch, flight, and landing. The high-torque brushed DC motor will break the shear pins and allow for the separation of launch vehicle components. To secure the bulkhead to the airframe and the nose cone, the LBR team will fasten them in from the outside using ⅛ screws. In the custom 3D printed bulkhead holding the motor, the LBR team created a detachable component to securely mount the motor. 91

93 Figure 40: 3D Printed Bulkhead Between the motor bulkhead and the main parachute bulkhead is a 5-inch space that houses the RDM electronic components and hardware. To secure all the electronic components during flight, the team will have a 3D printed avionics housing. 6.6 Testing Plan Testing Plan Breaking Shear Pins Drop Test Maximum Load - Current Drawn Maximum Load - Nose cone and Rover Deployment Response in Difficult Terrain Transmitting Distance Table 45: Testing Plans Description Connect the airframe, nose cone, and motor bulkhead together and drill holes into the airframe for the shear pins. Then run the motor to see if the motor is strong enough to break the shear pins and separate the nose cone from the airframe. Connect the airframe, nose cone and motor bulkheads together and secure them together. Then drop it from different heights, (heights) and then test to see if nose cone can still separate after impact. To ensure that the motor is strong enough to push off the 250 lb nose cone, knowing that the maximum current the motor can safely handle is 4.5 Amperes, through calculations, LBR will determine the maximum load that will draw that current and connect the motor to an ammeter and monitor the current drawn from this load at different times. Using the nose cone, LBR, will repeatedly run the motor to ensure that the nose cone and airframe will separate and that when the two separate, this allows enough room for the rover to separate from the internal compartment. To ensure that the RDM will work in different terrain, LBR will test the RDM mechanism in terrain with lots of potential room for failure. For example, LBR will bury half of the rocket in sand or dirt and activate it to see if the mechanism will deploy. In addition, the RDM will be tested in terrain with lots of rocks and uneven surfaces and monitor how the RDM responds in these situations. To test the transmitting distance, LBR will test the RDM at different distances between ½ to 1 mile to determine the maximum distance for which the RDM will respond to the signals. 92

94 6.7 RDM Team - Derived Requirements RDM Team Derived Requirements Table 46: RDM Team Derived Requirements Verification Method Solution Terrain blocking the nose cone from extending Securing the bulkheads Securing the nose cone to the airframe during launch Applying enough torque to break the shear pins Responding if the RDM gets stuck Ensuring that the RDM components are small enough so that the rover has space within the launch vehicle Controlling the activation of RDM remotely Testing Inspection Inspection Testing Testing Inspection Testing Strong motors will push the nose cone out of the way to ensure that the nose cone won t be prevented from opening due to the terrain. The bulkheads will be fastened by screws from outside the structure. The nose cone will be secured to the airframe by shear pins so that the components will stay intact during launch, so the vehicle can fly safely. High-torque brushed DC motors will have enough torque to break the shear pins. The rotary encoder was added to the motor so that it will recognize changes in movement of the motor shaft. It will send a signal to the Arduino to stop and retract the motor for a predefined period. Any object interfering with the deployment mechanism to be omitted. Limiting the design to two rods, one threaded and one unthreaded, frees up space internally. The RDM will be remotely triggered by a hobby-grade FlySky 2.4 GHz radio transmitter. 93

95 Preventing nose cone from rotating after the nose cone separates from the airframe. Testing The team plans on putting two unthreaded rods on the inside of the airframe that are the same length as the threaded rod to prevent the nose cone from rotating as it separates from the threaded rod and the airframe. 94

96 Section 7: Payload Criteria Autonomous Rover 7.1 Overview The team has chosen to design the autonomous deployable rover. The design must fit within the constraints of the rocket payload section while complying with the rover deployment system in order to fit and function properly. In order to find the best solution to complete the required task, the team has developed multiple rover designs with different advantages which will be implemented and tested until a final design can be concluded. Rover Mission Requirements 1. The rover will be a custom designed autonomous vehicle that will be stored within a payload bay on the launch vehicle. 2. Once the launch vehicle has landed, the LBR team will remotely initiate an autonomously deployment of the rover through the RDM which will traverse across a rugged terrain until it is at least 5 feet away from the rocket. 3. When the rover has surpassed the required distance, it will stop and deploy a set of solar panels which will be used to charge the power supply onboard the rover. System Summary Team members were considering two main designs for the rover located inside of the rocket. The challenge for the rover design is to create a vehicle that is efficiently stored within the payload bay of the rocket that maximizes the use of available space, so it must be designed with the internal structure of the deployment system in mind. This vehicle must be capable of driving in rough terrain regardless of the orientation the rocket lands due to the rover being deployed in an arbitrary manner. It must also withstand forces of being pushed out of the payload against the nosecone yet light enough that it doesn t inhibit the performance of the rocket. The team intends to use a microcontroller in the rover which will be used to control the motors and actuators, while reading positioning sensors and a wireless transceiver module which will be used to communicate back to the rocket to provide a real-time location system to judge the distance between the rover and the rocket. The team intends to use a microcontroller in the rover which will be used to control the motors for the drive mechanism and actuators which will be used to deploy the solar cells. The rover will know its position and distance with respect to the launch vehicle by using an inertial measurement unit to know the heading and orientation, as well as a wireless transceiver module which will be used to communicate back to the rocket to provide a real-time location system. This will ensure the rover surpasses the minimum 5 feet of distance it needs to travel from the rocket. 95

97 7.2 Rover Design Option 1: Wheg Wheel Rover The first design uses a more conventional chassis system, housing the rover servos, electronics and the motors within the frame. The solar panels will fold on the top of the rover to protect the solar panel surface from damage while driving, and when it is time for deployment they will fold out to double the surface area. A downside to this design is the usage of the space available in the rocket. This design does not fully utilize the space inside the rocket because of its rectangular design. The triangular rover design fully utilizes the available space inside the rocket. The wheels were carefully chosen knowing that the surface the rover will traverse could potentially be extremely rough terrain. Wheg wheels will be designed and used on the rover. This specific wheel design is able to climb over multiple types of rough obstacles. It can climb over twigs, rocks and other rugged terrain. Under the circumstance that it is deployed upside down, the wheg wheels would still be able to move the rover in the direction desired. The orientation of the rover would not matter after it is deployed from the rocket due to the design of the wheg wheels. Solar panels would also be deployed from the top of the rover. In the case that the rover is moving upside down, the solar panels could be used to flip the whole rover over after it reaches its deployment location. Servos will be utilized to deploy the solar panels. The servos will have more than enough torque and power required to be able to flip the rover if needed. All of the electronics will be controlled by an Arduino Nano. The motors will have a motor shield and the battery would also be located inside the rover body. Bogie System Figure 41: General Design of Wheg Wheel Rover Overview LBR considered a bogie system combined with wheg wheels for the rectangular rover, but decided against it. A bogie system takes up more room within rover, is more complex, and needs more materials. It is also unnecessary since body rigidity is needed to accompany wheg wheels 96

98 for wheg wheels to perform the best. In fact, a bogie system would actually make a wheg wheeled rover less controllable since the bogies will bounce up and down to compensate for the design of the wheg wheels. Without a bogie system, the rectangular rover will have a simpler design, more room within rover, and would weigh less. With more room inside the rover, it would result in more room for other systems. A rectangular rover with wheg wheels will be more stable without a bogie system. Wheg wheels combined with a bogie system will cause the bogie system to bounce and pivot uncontrollably since it would compensate for every bump or step that the wheg wheel makes. Wheg wills will work better with a rigid drivetrain because not all wheg wheels will contact the ground at the same time, some will be in the air while others will touch the ground causing the rover to be just balanced if the wheels are timed correctly. A bogie system will make a wheg wheeled rover be too flimsy. Option 2: Triangular Rover The second design is a triangular shaped rover that fits sideways in the payload bay. There is a wheel on each end of the triangle so that it it can be deployed in any orientation. Because it is deployed sideways in the payload bay, when deployed and exposed to the environment it is already facing the correct direction to move away from the rocket. Since the rover is wide it will be difficult to flip on its side and can still continue trekking if flipped frontwards or backwards. The triangular design allows for maximizing the size of the interior of the rover while being able to fit in the payload bay, which allows for additional space to store the electronics. Once the rover reaches a desired distance from the rocket it will deploy its solar panels using a single servo on a hinge mechanism. A triangular rover design presents a much more volume efficient craft over a rectangular design. The three-wheel layout allows the rover to be deployed from any orientation and still be able to move and deploy solar panels; flaps of two of the three sides will open up to deploy the panels. If one the flaps is parallel to the ground upon deployment, the opening process will rotate the rover upright. The triangular rover design has no orientation so it can deploy on any side. It uses space more efficiently since it is fitted sideways into the payload, and its triangular shape allows it to comply with the rover deployment mechanism. The size of the wheels can be maximized since the wheels curve along the walls of the payload bay, and the larger wheels result in an elevated ground clearance to go over rough terrain. The three-wheel design allows for a single gear in the middle to drive all three wheels simultaneously, which will make the rover only require two motors to drive all six wheels. Since it is triangle shaped, when the solar panels deploy there is more surface area exposed to the sun for the solar panels to collect. Due to the design having a higher center of gravity it is more prone to flipping frontwards or backwards, especially when going up a steep surface, but the rover is symmetrical and when the solar panels deploy they will orient the rover back in its upright position. 97

99 Figure 42: General Design of Triangular Rover This rover has a minimalistic and simple design to minimize possible failure on the field. The rectangular design allows the electronics to efficiently fit inside the chassis to maximize the use of the internal space, while keeping the rover symmetrical without protruding components which might inhibit the performance of the rover. The solar panels will be deployed based on a hinge system, which will deploy from the top of the rover controlled by a single servo. This design has a low center of gravity which will keep the rover stable, and in the case that the rover is deployed on its top the rover is capable of flipping itself back to its correct orientation using the solar panel deployment system. While a simple design, the width and length of the rover does not allow for sufficient ground clearance while traversing the rugged terrain unless the team decides to use larger wheels, which would call for a narrower chassis in order for it to fit within the rocket payload bay. This does not maximize the available space located inside the rocket and would make it more difficult to house all the necessary electronics. Table 47: Estimated Mass of the Rover Estimated Mass of Components (oz) 2x Motors: 116 RPM Premium Planetary Gear Motor 6.46 Arduino Nano Microcontroller 0.21 L298N H-Bridge

100 11.1 V LiPo Battery 2.89 Dorito chassis x Gearbox covers x 1/2" Hex Bore Bearings 14.4 Scrambler Offroad 1.0" Scale Tires x 25t planetary gears x 35t large center gear x pinion gears 0.5 2x ½ wide center hex shafts 2.56 hardware 1.7 total 55.5 Center Driveshaft Overview The triangular rover will sit inside the payload bay with a thread rod going through the middle for easy storage with and deployment from the payload bay. At the same time, the triangular rover will have a bogie system which will need a pivot in the center of a triangular panel. As a result, it is imperative to find a driveshaft that will double as a thread, a pivot point for the bogie system, and an axle to transfer drive power from the inside of the rover to the gearbox. The result is the need for a hollow driveshaft. Table 48: Pros and Cons of Hex Coupler Hex Coupler as a center shaft Pros With the hex coupler, it has higher tensile strength. The hex shape acts as a key and locks the center of the gears with each other when transferring power. Unlike the tube, there s no need to a clamping hub to lock the gear to the center shaft. Cons Because of the hex shape, stress concentrates at the corners. It s more challenging to find matching parts to fit such a system such as gears or bearings with a hex bore center. 99

101 The center of a hex coupler also doubles as the thread to push the rover out of the payload bay. Hollow Aluminum Tube as center shaft Pros Stress is uniformly dispersed along the surface of the cylinder. It is easier to find parts for, such as bearings and gears with circular bores. Cons The gear hub clamps may slip with the hollow tube design. With this design there will be no thread through the middle. It inherently makes the gearbox wider since gears must be attached to hubs which will attach to the hollow aluminum tube, versus hex centers with don t need to a clamp to lock onto the shaft. Table 49: Pros and Cons of Hollow Aluminum Tube LBR decided to go with the hex coupler center driveshaft because it is hollow with threads in the middle to allow the rover to be unscrewed from the payload bay. Also, because the center driveshaft is hexagonal shaped, the driveshaft can transfer power in more efficiently. In order to transfer power with a hollow round tube, the gear from the inside of the rover will need to be clamped onto the tube with a hub adapter, which takes too much space, and won t guarantee power transfer since there might be slip with a round clamp. On the other hand, hex center bored gears can lock onto a hex coupler to transfer power without the need or a hex hub. Besides, with the use of a hex bore bearing, the hex coupler can be doubled as a pivot for the bogie system. Bogie System Overview The terrain at the launch site has very rough terrain. Besides, the rocket can land anywhere, and the rover must be ready for any kind of terrain, no matter how rough. While a bogie system is advantageous, it adds complexity and takes away interior space from the rover. It is not imperative to use a bogie system for the rover, since the rover most likely will be able to traverse rough terrain without a bogie system and ground clearance is more important for traversing rough terrain. Therefore, there is a need to consider whether a bogie system is needed or not. Table 50: Pros and Cons with/without Bogie System With Bogie System 100

102 Pros All 4 wheels can contact the ground - All the wheels touching the ground allows greater grip and more thrust. Bogie system acts as the suspension- Bogie system gives greater stability to the drown allowing more accurate data from sensors that rely on orientation. This allows for greater maneuvering across difficult terrain. Cons With this design the rover will take up more room within rover - The housing needed for the bogie system will increase the space needed and reduces spaces for other systems. The bogie system will increase complexity to the project and the devoted time and resources that could be used elsewhere. More parts needed - More resources and parts needed to implement this system than would be use without it. It will be unnecessary if little to no difficult terrain is encountered. Without Bogie System Pros Simple - The overall design without the Boogie system will be simpler. More room within rover - The bogie system uses a high amount of volume and not including it would result in more room for other systems. Less weight- Not having the boogie system would decrease the overall weight of the design. Cons Less flexibility over terrain - The bogie system allows the design to have a greater range of terrains that the rover is capable of traversing. No suspension- Data with respect to axis from sensors have a potential to be inaccurate due to vibrations induced by rough terrain. In rough terrain the not all four wheels can contact the ground. LBR will go with a bogie system allows for more flexibility over terrain. With a bogie system, all 4 (of 6) wheels can contact the ground always for more traction. While the bogie will add much more complexity and room, it is advantageous if the rover is equipped with a bogie system. With team member interest in tackling the complexity of the drivetrain and the challenges of packaging electronics within the rover, a bogie system is feasible on the rover, therefore LBR concluded to equip the rover with a bogie system. 101

103 Gear Selection Overview A gearbox is needed to transfer power to the triplet of wheels that each motor will power. There are many possible combinations to arrange motors in a triangular pattern. Since the triangular rover will traverse rough terrain and the process is not timed, speed is not an important factor. Therefore, a torquey setup is preferred. But, a torquey setup will require a small center gear and large planet gears. Since large planet gears will take away ground clearance, a compromised same sized gear all around setup will do. Table 51: Pros and Cons with Various Gears Big Center Gear, small surrounding gears Pros With the big center gear, the rover will be more energy efficient. This design will also help with the ground clearance. Cons A bigger center gear means there will be less torque to the surrounding gears. Less torque means the gears will be faster. Same sized gears all around Due to the bigger gear the movement of the rover will be less precise. Pros Easier to find matching gears to meet specific requirements (needs to be same pitch, center needs to have ½ hex bore). Cons Not as much ground clearance compared to the big center gear design with small surrounding gears. This gives decent torque to the wheels. Optimized for gearing size for both torque and ground clearance. Small center gear, big surrounding gears Pros This design is the most ideal for torque to the wheels compared to the other design choices. Cons Less ground clearance, bad angle of attack from wheels to gearbox body 102

104 With this design the rover can be controlled with more precision. LBR decided to go with same sized gears all around as a compromise to torque and ground clearance, but since using a hex coupler as the driveshaft, gear choices were limited. As a result, matching gears were difficult to source up to the design specifications. The center gear must have a ½ inch hex bore size. Since 48P and 32P gears weren t available in ½ inch hex bore size, a 20P gear was selected. As a result, all other gears need to be 20P, but don t have to have a ½ inch hex bore. Since 20P gear choices are limited, they usually come with a ½ inch hex bore if a hex bore is required, which it is to power the wheels since the wheels and planet gears need to be synchronized. As a result, 20P ½ inch hex bore gears were selected to be used all around. Equal sized gears could not be found, resulting in one gear being bigger than the other. Therefore, the team opted to put the bigger gear in the center and assign the planet gears to be the smaller ones. Figure 43: Triangular Rover Gearbox The entire gearbox pivots on a bogie system. A motor mounted inside the rover powers the gear inside the rover, which will transfer the torque through a hex coupler (that also doubles as a center thread to deploy the rover), powering the center gear in the gearbox which powers the 3 surrounding gears to turn the wheels. The gears are as similar in diameter as possible. Expandable wheels design With a bogie system that involves a gearbox, interior space of the triangular rover is limited. There may not be enough interior space to fit an H-bridge, two motors, an Arduino Nano, a gear on both sides, and a battery, in addition to fitting a threaded rod through the rover. Therefore, the need for an expandable wheel design was considered. Table 51: Pros and Cons of the Wheels Expandable wheels Pros Cons 103

105 This potentially means more rover body volume is available in the RDM. The space within the payload bay will be used more efficiently/maximized. A more complex gearbox design resulting from installing servos within the gearbox or actuate the wheels to expand out. Non expandable wheels It is heavier because of more components. This introduces possibility of objects blocking wheel deployment Pros The wheels are simpler to use. Less components are required so the rover will be lighter overall, Cons This limits the body size for ground clearance. The body size would likely contain a larger volume and provide less room to design around. A servo would be needed that is strong enough to lift the rover yet small enough to fit within the rover s gearbox. A linkage mechanism would also be needed to expand the wheels outward, while not interfering with the center where a threaded rod would go through. In addition, a gearbox needs to be designed in a way that the wheels will be powered when it is expanded. All of this complexity will add weight. The expandable wheels will not fix the triangular rover s stability problem, where the triangular rover tends to fall forwards or backwards on a steep incline. While an expandable wheel design was taken into consideration by the team to expand volume and interior space within the rover, the team ultimately decided against an expandable wheel design because it would increase complexity and points of failure to the design. Option 3: Cylindrical Two Wheeled Rover Figure 44: Two-Wheel Rover Design 104

106 This rover design is the most volume efficient. It s cylindrical cross section allows it to fit easily into the rocket while allowing to incorporate all the electronics needed into the carriage. Solar panels are located on the inner body cover which will be deployed after reaching the final displacement. Incorporating a cover also allows all the electronics and the solar panels to be stored safely into the carriage during the launch sequence of the rocket. While this design is ideal in terms of incorporating the rover into the RDM, stability is sacrificed. It is much more difficult to keep the rover upright during travel; additionally, the rover can only be deployed from one orientation. This rover design also has limited ability to overcome obstacles in its path. An incline grade of over 5 degrees could cause the rover to lose its balance and topple. 7.3 Overall Rover Designs All design choices for the rover stem from the overall rover structure. LBR considered three different rover designs: Rectangular, Triangular, and Cylindrical. The team approached the decision-making process with the primary goal of the mission in mind. A successful rover in the context of this mission must be able to effectively fit in the RDM; additionally, the rover must house the required electronics and solar panels and be able to incorporate these elements into the design effectively. ` Overall weight and complexity are two other important considerations. If the rover design doesn t allow for ideal displacement of its load, the craft could easily topple while traveling. A simpler design generally allows for more intuitive packing and makes the construction process much more ideal. While considering all these choices, the team had to justify each of these choices with the overarching idea that the rover must be strong and stable. Having ideal load distribution is important, but if the rover is destroyed or unable to function after the launch sequence, the rover has failed the mission. Taking all these variables into consideration, a chart has been made listing the pros and cons of each rover structure to simplify the decision-making process. Table 52: Pros and Cons on Rover Design Rectangular Rover Design Pros This design is simpler. The overall rover will be lighter. Cons The rover can only be deployed from a specific orientation. Solar Panel deployment requires flaps to rotate at least 120 Degrees. 105

107 A rectangular rover has higher stress resistance. The design does not maximize the space inside the rocket. A wheg wheeled design gives it best ground clearance and all terrain performance. Triangular Rover Design Pros The triangular rover design can be deployed in any orientation. Solar panel deployment is more intuitive thus requiring less rotation. The design allows more volume inside the rover for needed electronic parts. Cons The rover is not stable and could flip over on steep slopes. Since there are more parts needed for the triangular rover design, there is a bigger chance of part failure. It is challenging to manufacture triangular shapes with round edges on the available machines. The shape when placed parallel to rocket gives more room. Cylindrical Rover Design Pros The cylindrical rover design makes the most use of payload space. The cylindrical rover design is the simplest rover design. Easier to model and manufacture compared to the triangular rover design. Cons This design is not unique and there is not much team member interest. It is not the most stable forwards and backwards, the stabilizer can lift up and possibly flip to the other side under braking or a huge bump or under reverse. Table 53: Rover Trade Study Rover Trade Study Rectangular Rover with Wheg wheels Triangular Rover Triangular Rover with tank treads Cylindrical Rover 106

108 Team Member Interest Control System Complexity Cost Volume efficiency within the rocket Ground Clearance Stability Weight Total The team has decided to go with a triangular rover design. It is balanced in stability, ground clearance, and maximized the space within the payload bay. A bogie system will help the rover cope with rough terrain better, and it faces the correct direction upon exit from the payload bay to travel perpendicularly away from the launch vehicle. The triangular rover can also deploy in any orientation since it will always land on a set of four wheels. Overall, it is the best design with the least compromises. 7.4 Wheel Design One of the defining features of any rover is the design and material choice of the wheels. With a plethora of different material and design approaches available, each with varying pros and cons, the team decided to do a trade study to compare the strengths and weaknesses of each and decide based on how compatible each is with respect to the mission objective. Wheels and tires Failure to make the correct design choice regarding the wheels of the rover could result in a failed mission. If the wheels are unable to overcome the burdens of the terrain, the rover will be unable to move from the rocket and, consequently, be unable to deploy its solar panels. The choices have been narrowed down to two possibilities: 3D printed wheels or Beadlocks with RC Crawler tires. Table 54: Pros and Cons of Wheel Design 107

109 3D Printed The wheels are lighter. Pros They can be custom made to meet specific design requirements They are cost effective- The cost will be greatly reduced on the tires since no specialized tools will be needed and the materials are cheap. It will be difficult to attach the tire to wheel securely. Cons Less grip - These tires will have a smooth texture unless the tires are made with treads which will increase which will increase the difficulty for the tire to print. Not flexible - The tires will be made of a plastic that has little flexibility causing the tires not to grip the terrain as effectively. It is challenging to design an effective tread design The wheels could need treads in order to increase grip across steeper environments or loose terrain. 3d printing wheels would present a design challenge both for designing and for the printer itself to manufacture. Not precise enough - 3d printed parts are not as precise as part manufactured by other machines and have a lower range of designability. They Allow structures such as hex grids to be made to maximize strength and minimize materials and weight. Beadlocks and RC Crawler Tires Pros Strong treads, more grip. These tires will have stronger treads which will allow a greater grip. More flexible -The material used will allow greater tire flexibility allowing greater grip. Cons More mass- These tires will contain more mass than other tires. More expensive -These tires will have a greater expense than other design models. 108

110 Figure 45: Rover Wheel Design Selection Table 55: Wheel Trade Study Wheel Trade Study 3D Printed Wheels and Tires 3D Printed Wheels and RC Rock Crawler Tires RC Aluminum Beadlock wheels and RC Rock Crawler Tires Tanks Treads Wheg Wheel Team Member Interest Complexity to design Complexity to manufacture Grip Cost Compatibility with Rectangular Rover Rectangle Rover Total

111 Compatibility with Triangular Rover Triangular Rover Total Compatibility with Cylindrical Rover Cylindrical Rover Total Since LBR has decided to use a triangular rover design, round wheeled tires have been determined to be the optimal setup. The main reason for the exclusion of the wheg wheels or tank tracks is the instability that will be resulted when they are combined with the triangular rover design. 3D printed tires will not provide enough grip and is more challenging to design; it is also easier to buy RC rock crawler tires, which provide a lot of grip over rough terrain. Although aluminum beadlock wheels are stronger, 3D printed wheels are slightly lighter much more cost effective. Ultimately, the team has decided to use 3D printed wheels with RC rock crawler tires. 7.5 Material The rover needs to be strong yet light. The launch and landing of the rocket will put high stress on the craft, so the team will focus on this property when choosing the material. A rocket launch of over 3Gs could cause the rover to compress and deform. LBR started rover research looking at materials with high tensile strength. Carbon Fiber has a high ultimate tensile strength and has a great strength to weight ratio; however, it is expensive and not easy to make properly. Aluminum is much easier to machine and work, but would be sacrificing strength with this material. ABS plastic is another material to consider. The team can quickly print out new rover parts for testing and using the material is much cheaper than the alternatives. ABS has the lowest tensile strength of only MPa, and could cause the rover to deform prior to deployment. Table 56: Pros and Cons Materials for the Rover Carbon Fiber Pros Carbon Fiber has a higher ultimate tensile strength (3.5 GPa) Cons Expensive- Carbon fiber will be expensive since the cost of the materials is high combined with possible tools needed to mold or manufacture it properly. 110

112 High strength to weight ratio - This material is exceptionally light while also providing a high amount of strength. More difficult to properly form due to less control of process and greater difficulty forming. Aluminum Pros Aluminum is relatively light (D = 2700 kg/m^3) It is cost efficient (strong while being relatively cheap compared to carbon fiber) Cons Aluminum is heavier than carbon fiber or 3D printed ABS - Aluminum will have a higher mass than 3D printed ABS or carbon fiber. It has a lower ultimate tensile strength (290 MPa) Easy machining- The machining required for aluminum is relatively simple and does not require specialized tools. 3D Printed (ABS) Pros The material can be easily acquired. Low cost- Due to the machine for constructing the rover already acquired, this material will require no special tools and cost for materials is lower compared to other materials. It has excellent impact resistance Cons Not precise- There is less precision from constructing the rover out of this material, which could interfere with components that need a perfect fit. ABS is not as strong as aluminum or carbon fiber. The tensile strength is low ( MPa) ABS has great aesthetic properties If the budget permits, carbon fiber is the preferred material for the rover. It is light yet strong. Since the budget is limited, the team has decided to go with 3D printed ABS. While total weight of aluminum in the rover isn t significantly heavier than 3D printed ABS, every single ounce counts. 3D printed ABS is light but strong enough to withstand the stresses of being inside the payload during the launch and landing. If budget was not an issue, carbon fiber is the preferred material because it is light and stronger than 3D printed ABS. 111

113 7.6 Motor Most of the space in the rover will be occupied by the motors and battery. The rover will be deployed from the rocket with a threaded rod through its center, so these components must still be able to effectively fit without causing this mechanism to jam or cause a blockage. Balancing power with weight/volume is crucial when approaching motor choice. Brushless motors were first considered to move the rover. They generally are high rpm and would allow the rover to overcome obstacles with ease; however, they use more power. The motors themselves are light, but the batteries that accompany a high power motor are heavy. With the limitations of a brushless motor in mind, other motor possibilities were taken into consideration to see if they could overcome these weaknesses. Table 57: Pros and Cons on the Motor Brushless Motor Pros Brushless motors have extremely high rpm. Lighter than a brushed motor- It has less mass compared a brushed motor. Brushless motors are efficient in lower rpm because at lower rpm it saps less electrical energy than at high rpm. The torque produced by a brushless motor is higher. Cons A faster motor means more power consumption. Due to higher power consumption, there will be more heat generated. Higher power consumption also means a bigger capacity battery will be used, and a larger battery will carry more weight. Brushless motors are more difficult to program. Higher speed is not necessary Brushed Motor Pros The dimensions of a brushed motor tend to be smaller compared to brushless motors. Cons Brushed motors are slower than brushless motors. Brushed motors have simpler power controls. The brushed motors have lower rpm rating. Due to the smaller dimensions, brushed motors have less weight. Brushed motors also produce less torque compared to brushless motors. 112

114 Brushed motors have less jitter in lower rpm, therefore more precise control. These motors generate more heat due to the contact between brushes and the commutator. Stepper Motor Pros Stepper motors will allow highly precise means of controlling the speed of the wheel and the circumvention. Cons The stepper motor will increase the difficulty of programming the system. The stepper motor has a low amount of torque making it harder for it to navigate rough or steep terrain. Stepper motors are too slow. It does not output as much torque as a brushed or brushless motor. Continuous Servo Pros It is more compact and easier to work with. The motor will be easier to program. Cons These motors output less torque. It will give less speed. LBR choose brushed motors with gearboxes for torque, low rpm control, and ease of programing. They are a better value, simpler to control, and have enough torque for the required application. Since high RPM is not required for this application, a brushless motor isn t necessary. In comparison, servos and stepper motors may be more accurate but do not provide adequate torque. 7.7 Sensors Sensors allow the rover to detect and react to environmental conditions faced by the rover. With correct application, the craft can avoid obstacles that it would otherwise be unable to overcome on its own. If the rover senses a large object in its path, it could adjust its course to avoid the obstacle entirely. Sensors are also great for providing the rover with a sense of direction and balance. The rover can adjust its speed based on its orientation and rotational acceleration to avoid toppling. The rover also needs a way to tell how far from the rocket it is. It would be a disaster if the rover deployed its solar panels at the wrong time/place or had objects next to it that could prevent them from deploying properly. 113

115 While all these potential applications are valuable, the team had to consider what complexity is necessary for this mission. The terrain is rough, but it is unlikely that there will be an obstacle that would be worth going around entirely. Slight adjustments to its course should suffice. The rover is designed to can overcome typical field obstacles in its path. Speed control, on the other hand, is much more important to consider. If the rover is to overcome most of the obstacles in its path, it needs to approach and tackle them with care to avoid toppling. Table 58: Pros and Cons of the Sensors RangeFinder Pros It occupies a smaller volume than the ultrasonic sensor (44mm x 7mm x 1.5mm). This is imperative due to the rovers limited housing space for electrical parts. Cons The range of the rangefinder gives an accurate reading of mm with. Any value outside this range is a constant value of 255, allowing the detection lack of objects within the range. The rangefinder will output values with an accuracy within 1mm, making it the more accurate of the 2 sensors. Ultrasonic Sensor Pros The ultrasonic sensor gives range values between 2 cm and 400 cm. However, the for objects or lack of objects outside of this range are inaccurate and give jumping values. A program would need to be made to filter out these out-of-range values. Cons The Ultrasonic Sensor uses a larger volume of space within the rover(45mm x 15mm x 20mm). Inertia Measurement Unit The values given by the ultrasonic sensor have an accuracy within 3mm. However, the 2mm difference between the accuracy of the sensors are a secondary concern. Values outside of range give inaccurate reading that have no consistency and a filter must be programmed to give accurate readings. 114

116 Pros The sensor is capable of detecting acceleration and inertia across all three direction giving values for acceleration. It is capable of detecting slight shifts in direction in order to move across a more linear path. Cons Since it is only giving data on acceleration and inertia, it has not object detection until collusion. It does not give the value for velocity or position and other sensors would be needed to overcome these shortfalls Barometer Pros The sensor is capable of detecting pressure and with correct initialization, the ability to derive altitude from it. Cons Errors from this sensor have an expected range of up to 1 meter meaning it can only be accurate over larger distances. Unless traversing large distances, the Barometer will be of little use. Table 59: Trade Study on the Sensors Sensor Barometer IMU Range Finder Sonic Category Weight Value Score Value Score Value Score Value Score Size 15% Precision 30% Data Utility 40% Cost 10% Weight 5% Total Score As shown by the trade study above, the IMU highest scoring and will be implemented. This is partly due to the sensor providing data for several different utilities that increase its versatility. On 115

117 the opposite end of the spectrum is the sonic sensor, which failed due to its sensors having little utility for the purpose of the rover and its size and weight being larger than the other sensors. The rangefinder and barometer are both in yellow because although they did win in overall score the barometer can be useful for function that the IMU cannot provide such as elevation and the rangefinder can help with object avoidance like the sonic sensor but are more useful overall. 7.8 Testing Plan The rover will be subjected to extreme forces upon launch, flight, and landing. It must also be light, but strong, capable of traversing over extreme terrain. To ensure the rover will survive these conditions and complete its mission, the team will perform launch tests, shake tests, drop tests, subject the rover to extremely rough terrain, and subject it to abnormal conditions. For example, the rover needs to withstand a drop of 20 ft/s. To ensure it to survive such as drop, the rover will be subjected to a 30 ft/s drop. During flight, the rover will be subjected to vibration. To ensure the rover will be able to withstand flight, the rover will be put it in a testing apparatus that will shake it for an hour continuously. To ensure the rover will be able to withstand launch force, the team will complete the rover in time for test launch where it will be placed in the payload bay of the subscale rocket. The launch, shake, and drop tests will make sure the structure of the rover can withstand extreme forces, wires not to tangle, unplug, or break, and sensors with not be damaged. The rover will also be subjected to abnormal conditions, for example, if the thread has not completely detached from the rover, or the nose cone isn t completely pushed off, or a rock is obstructing the pathway of the rover trapping the rover between the payload bay and nose cone, or dirt contaminating the sensors, obstructing. Addressing one of the main concerns the team had with the rover designs, an incline test will be performed by making the rover climb a 45-degree dirt slope that has ruts big enough to engulf the rover, simulating extreme rough terrain. Only after the same rover has passed all of these tests consecutively will the rover be deemed mission ready. 7.9 Controls All three of the rover designs will use differential drive and a control system to instruct the rover to follow a linear path. A three-dimensional accelerometer and gyroscope will allow for any change in angle to be detected along the x and y axis. Yaw suppression will be used for when the rover moves too far from one side, the corresponding motor on that side will increase RPM to guide the rover back to its original path. There will also be several other sensors to be able to sense object that range in difficulty from steep to impassible. Through testing, LBR can determine constants the responsiveness of the system. It needs to be designed to maintain a path, but needs to be able to divert from its path if necessary in order to avoid objects. However, it must remember the diversion in order to move back on a linear path. 116

118 Figure 46: Control Schematic For the rover system, the set point will be the direction the rover starts after it is deployed by the RDM. The drone will start to move forward and will look for any source of disturbance that would disrupt its path. If something causes a disturbance, the sensors will discover either discover the object in its path or the three-dimensional gyroscope detecting turning. The controller is a printed circuit board (Arduino Nano) that will be able to read the information and send an appropriate voltage to the motor shield. The motor shield will delegate how much power each of the motors will get and the motors will affect the state of affairs. Electronics Each one of the three rover designs will include similar electronics, including DC brushed motors, a motor driver, inertial measurement unit, and servo motors to deploy the solar panels. They will all use an Arduino Nano microcontroller. The Nano will be a Master Writer and Master Reader since it will be both receiving data and delegating commands based on the data. The sensors will be slave writer since they will input data for the Master Reader to process. The Master Writer will then delegate commands to the motor shield as a Slave Receiver to then change output sent to the motors. Microcontroller: 117

119 Figure 47: Arduino Nano The Arduino Nano is the microcontroller chosen due to it best fitting the limited housing constraints of all the possible rover designs. The Nano gives contains enough pins that can receive and transmit analog values. The Nano will also be paired with a motor shield since the shield allows the motors to receive voltages that would normally short-circuit the Nano. A 11 volt LiPo battery will then be connected to the motor shield and the shield will power the Nano through a port on the motor shield that will consistently output 5 volts. This is done so the LiPo can be used to its full potential without endangering the Nano. Motors: Figure 48: Motor The rover will have two motors that work independently of each other. The Nano will write to the motor shield which will write to the motors to control the movement. The sensors will make sure the rover will follow a straight path. Sensors 118

120 As the rover is moving forward, an inertial measurement unit will detect any potential shift in direction and will tell the Nano about any such changes. The Nano will then increase or decrease power to corresponding motor in order to correct the shift. The Nano will not attempt to a specific value but instead a possible range so the Nano will not make minor shifts that could result in recursively overcorrecting. In addition, the rover will use object detection to detect obstacles that could impede the movement of the rover Rover Team-Derived Requirements Rover Team Derived Requirements Table 60: Rover Team Derived Requirements Verification Method Verification Plan Rover must climb 45- degree dirt slope. Rover must weigh less than 4 pounds. Rover must survive a drop of up to 30 ft/sec. Rover must be no longer than 10in long and fit in a 6in tube. Rover must function in any random orientation. The rover will have low CG The rover will use lightweight materials. Testing will be implemented to meet the requirements. Rover will be designed to fit the restrictions. Due to the rover s design, it can travel in any After computer aided design and manufacturing of the rover, the rover will be tested to make sure it meets the requirements. One of the tests will be having the rover traverse up multiple different terrains and in varying slopes. The rover will be manufactured using lightweight materials. 3d printed filament such as ABS will be used. Rover will be dropped from a parking structure to reach its minimum success requirement of 30ft/sec. The height needed to reach the minimum success requirement will be calculated before the drop test to ensure accuracy in the test results. On the stage of computer aided design, the overall rover length and height will be considered for the rover to fit inside the given restrictions. Before moving to the manufacturing process, a full assembly of the rover will be finalized with the CAD to finalize the rover dimensions. The triangular rover design will be used because it can deploy in any orientation and would still be able to 119

121 orientation after deployment. traverse through the landing site s terrain. Because of the three-wheel design it does not need a specific landing orientation. 120

122 Section 8: Scientific Research Airbrake 8.1 Overview For this year s competition the LBR team decided to add an extra level of complexity to the rocket by adding air brakes to assist in the rocket reaching the desired altitude. The team intends to slightly overshoot the one-mile mark but use an air brake system with sensor feedback to control the speed of the rocket to ensure the most accurate elevation for apogee. Using instantaneous velocity, acceleration, and altitude readings, the team will create a control system to add or reduce drag accordingly. 8.2 Air Brake Deployment Designs For the air brake deployment mechanism, the team decided to go with a vertical deployment mechanism (VDM) over the horizontal deployment mechanism (HDM). Horizontal air brake deployment mechanism Figure 49: Horizontal Deployment Mechanism Table 61: Pros and Cons of the Horizontal Mechanism Pros Cons Maximum drag is produced by the same surface area at 90 degrees Small vertical space consumption The drag produced at such a steep angle will cause high stress to a very small area of the fin, which greatly increases the likelihood of damage to the fin The fins deploy by spinning out of the rocket, meaning there are only 3 small segments 121

123 around the fins that will hold the rocket together. This will lead to higher stress on those points, greatly increasing the likelihood of damage. Fast deployment speed due to the use of a single motor with a large gear reduction More complex design with the fins mounted to gears and the servo gear mounted inside to the main gear. High shearing force on the deployment mechanism due to the sharp increase of forces on the fin. The horizontal deployment mechanism (HDM) uses a singular servo placed in the center of the mechanism with a small gear attached to it with 3 flaps with teeth on one end of the flap that rotate to deploy the flaps. The team decided not to go with this design despite it deploying faster and taking up less vertical space. The horizontal deployment mechanism creates high amounts of pressure, putting stress on the flaps. Structurally, it would leave very little support material on the body of the rocket, which increases the likelihood of damage. Vertical air brake deployment mechanism Figure 50: Vertical Air Brake Deployment Mechanism Pros Table 62: Pros and Cons of Vertical Mechanism Cons Simpler design with parts that are easier to manufacture Less drag due to smaller angle of deployment 122

124 Larger surface area of support structure Less stress on the fins due the lower drag Higher vertical space consumption Much slower deployment time of 2-3 seconds Low shearing stress on the deployment mechanism due to a gradual increase in angle The VDM on the air brake relies on a singular linear actuator mounted to a bulkhead on the top of the subsystem that pulls up on a singular square piece of metal that is attached to 4 arms that attach to the flaps. As the linear actuator pulls up it causes the flaps to deploy out. This method is preferable due to its low and gradual increase in stress on the flaps versus the horizontal deployments sudden and rapid application of stress. This method also is simpler to manufacture and also provides greater surface area for support structure on the sides of the rocket. This method also allows for and easily implemented design for a feedback loop into the design. Air Brake Deployment Mechanism Trade Study Option: Vertical Deployment Horizontal Deployment Table 63: Airbrake Trade Study Category Weight Value Score Value Score Deployment speed 5.00% Simplicity 10.00% Manufacturability 25.00% Price 5.00% Weight 15.00% Potential Drag 15.00% Structural Integrity 25.00% Total Score As seen in the trade study above, the VDM is the optimal choice for the air brake deployment mechanism. Its manufacturability, potential drag and structural integrity are superior to the HDM. Despite the HDM s superiority in deployment speed, simplicity, price, and weight the VDM won out in the more heavily weighted categories. Additionally, the slower deployment speed for the VDM will be compensated by the deployment control loop. 123

125 8.3 Materials For the air brake LBR is planning on using a combination of materials to make up the entire subsystem. For the actual flaps of the assembly a strong yet lightweight is needed so carbon fiber was decided for the flaps due to the robustness of the material and its low weight mass. For the rest of the major components of the subsystem a robust material was still needed, however to keep a reasonable budget, machined aluminum was selected to make up the bulkheads, the arms, hinges, support bars, and centerpiece. These parts do not need the same amount of strength as the flaps so a lower quality material could be afforded in these areas. Finally, to hold the linear actuator in place along with the rest of the electronics LBR will be designing and 3-D printing a custom mount to hold all the electronic components together. 8.4 Control The air brake mechanism is designed to slow down the speed of the rocket in order to reach the desired apogee. The deployment timing and duration affects the drag which in turn controls the velocity of the launch vehicle. If the air brake deploys too early, it will result in prolonged drag which will make the rocket undershoot the desired altitude. If deployed too late, the rocket could be going slow to the point that the drag created by the air brakes gives little effect to velocity resulting in the rocket overshooting the desired altitude. Some parameters the team must consider when designing the control system of the air brakes include the current velocity and acceleration of the rocket, altitude, speed of flap deployment, and the amount of drag produced by the flaps depending on the distance deployed. Through simulation, the LBR team has information of the necessary velocity of the launch vehicle depending on the altitude in order to get to the desired apogee. Using sensors that read velocity and altitude of the launch vehicle, the team intends to create a control system that adds or reduces drag in response to the error in velocity. Figure 51: Controller Schematic The reference input is the desired velocity of the launch vehicle, where the controller will find the error in velocity and move the linear actuator for increased drag as the system input in order to reduce the error to get closer to the desired velocity. Onboard sensors will provide feedback to the system in order to keep correcting the speed of the launch vehicle until the desired speed is 124

126 achieved. Due to the high velocity of the rocket and relatively slow response of the linear actuator, the system will increment the response in sections where it will treat each incremented elevation block value as a checkpoint which will allow the actuator time to adapt accordingly. For the system to function as desired, the team must design the rocket to overshoot the one-mile apogee since the control system cannot speed up the rocket. This will allow for correction in speed by slowing the launch vehicle down, but it must be considered that the rocket must not be slowed down below the desired velocity since the control system cannot compensate for a lack of speed. To find the appropriate transfer functions for the system block of the control loop, LBR is using state space equations to calculate stability. x (t)=ax(t)+bu(t) y(t)=cx(t)+du(t) The controller must have readings of altitude and velocity, including drag factors at certain velocities. To calculate drag on the air brake system the rocket will be placed in a wind tunnel, where force will be measured depending on the deflect ion angle of the flaps. LBR has created a backup open-loop solution to control the air brake deployment mechanism which uses the same hardware and can be uploaded on the microcontroller under the circumstance that there is an error in the closed-loop system. This solution will not deliver as accurate of an apogee as the control system approach, but will still significantly decrease velocity of the rocket if it is overshooting the calculated values. The system will use data from the altimeter to find altitude and velocity readings and perform a response accordingly. If the rocket is reaching an altitude of 5000 feet above the required velocity, the air brakes will deploy at 25% deflection. If the rocket exceeds the desired 5280-foot apogee, the air brakes will deploy at full deflection for maximum drag to slow down the rocket. 8.5 Equations: 1. Drag: The drag force on the flaps will be calculated using D = C daρv 2 2 where Cd is the coefficient of drag, A is the reference area of the air brakes, p is the density of the air, and v is the velocity. As the projected area will change over time as the brakes are deployed, the coefficient of drag may also change, so the team wants to increase the speed of deployment to minimize the time it takes for the coefficient of drag to reach its final value. 2. Acceleration: 125

127 The acceleration of the rocket when the air brakes are fully deployed can be found by using the equation a = g D total m where a is the vertical acceleration, g is the acceleration due to gravity, D total is the total drag force on the rocket once the air brakes are deployed, and m is the total mass of the rocket after burn. 3. Velocity: Velocity can be represented as a function of height, derived from the acceleration equation. In the equation, c is the drag characteristics of the vehicle found by c=cdpa2m where Cd is the coefficient of drag of the whole rocket, p is the density of air, and A is the cross-sectional area of the rocket. 8.6 Electronics The air brake system must rely on the proper functionality of the electronics in order to be effective at slowing down the rocket for the desired apogee. This requires reliable sensors to collect rocket data, a powerful linear actuator that can deploy and retract the air brakes, and a capable microcontroller to compute the data. Linear Actuator: The vertical deployment mechanism extends the air brake flaps by sliding the square hinge along the three central linear rails. The system that provides the linear motion must have enough torque to handle the drag created by the flaps, and must be capable of deploying and retracting in compliance to the control system. LBR originally considered pneumatics due to the high speed and potential torque, but this adds weight and overall complexity to the system. LBR has decided on using an electronic linear actuator due to the high torque it provides and the small amount of space it consumes. This actuator will be controlled by the same microcontroller which reads the sensor inputs. Sensors: Sensors that determine the altitude of the launch vehicle rely on atmospheric pressure, considering that pressure decreases with increase in altitude. A barometric altimeter will be used to collect pressure data, which will be supplied with a nonlinear calibration to match the pressure readings with height. Due to pressure changes due to weather and location, the altimeter must be recalibrated at a known altitude in order to ensure accurate data readings. 126

128 Microcontroller: In order to read the inputs of the various sensors and perform an output function the air brake system will rely on a microcontroller. The team has decided on using the ATmega328 controller on the Arduino Nano platform due to the integrated electronics, ease of use, small footprint, and adequate number of general purpose input-output pins. I2C will be utilized to communicate between the microcontroller and the sensors, and it can provide a reliable pulse width modulated digital signal in order to control the speed of the linear actuator for air brake deployment. The Nano includes a built-in voltage regulator supporting an input of 7-12 volts, allowing it to be powered by the same 11.1V LiPo power source as the RDM. 8.7 Mechanical Design To make the air brake effective and plausible, LBR took several factors into consideration: 1. Weight 2. Potential Drag 3. Manufacturability 4. Structural integrity Weight Weight was a key deciding factor in determining which design will be used to deploy the air brakes, as more weight will decrease the amount of thrust the rocket can achieve. In the designs, the HDM had a lower weight than the VDM due to the fact the VDM requires additional materials as well as a high-torque linear actuator. Although this is the case, the additional weight of the VDM is marginal and as such, the greater weight from the VDM did not outweigh the other benefits it had to offer. Potential Drag The air brakes main function is to slow down the launch vehicle, therefore their effectiveness is measured by the ability to create drag. The potential drag of the air brakes is largely based on the chosen design s deployment system and fin shape. If a larger surface area is extended from the body of the rocket, a greater force of drag will be created, therefore the design for HDM was not ideal due to its use of smaller fins. To maintain the structural integrity of the rocket, the fins in the HDM could not be made any larger, therefore the secondary VDM option was considered. The potential drag created by the VDM when fully deployed is much larger due to the placement of the fins on the side of the rocket compared to inside the body of the rocket which allows for the fins to be larger without consuming excess space from the frame of the rocket. Manufacturability 127

129 Manufacturability was taken into consideration when choosing the VDM over the HDM. Though the HDM would be simpler to design, the VDM is more effective and still plausible to manufacture. The VDM will rely on structures that can be 3-D printed or machined. The VDM uses 4 rods as structural supports around a square shaped connector that moves up and down with the help of a linear actuator. The square has simple hinge openings for a rod to slot into with the other end slotting into another hinge opening on the flap. As the square moves up on the stabilizer rods, the deployment arms shift out, causing the flaps to deploy outwards to the desired angle and easily be retracted as per the control loop. Figure 52: Hinge Hub Center Piece Figure 53: Center Piece-Arm-Hinge-Flap Assembly 128

130 Structural Integrity LBR made several modifications to the original design to add additional structural integrity to the air brake deployment mechanism. The design has 4 rods that are used to support the centerpiece that the deployment arms are connected to prevent the mechanism from moving around during flight which could cause the deployment arms to become misaligned and potentially bind up or cause severe damage. Air Brake Flap Design There are three main designs considered for use as the flaps themselves. They are a slotted design, a long and thin design, and a short and wide design. For the rocket, the short and wide design was chosen due its balance of drag and structural integrity when compared to the other designs as explained below. For the slotted design, the flaps would be given angled slots along the length of the flap, similar to those of air vents. The slots would produce much more drag than a smooth surface as well as reducing the weight of the flaps themselves. However, because they are slotted, the flaps can only be as thick as the outer wall so there would be no slots in the inner wall. The flaps would also need to be made of a tougher material to avoid breaking because of the openings and increased drag. The slotted design would also be the most complicated and difficult design to make. For the long and thin design, the flaps would be completely smooth and longer going along the rocket body than across it. Due to their long length, they require less movement at the top of the flap for greater movement at the bottom, allowing for a greater angle of deployment for the air brake. The rocket also has more structural integrity since there will be more material between the flaps connecting the rest of the body. The size and shape of the flaps reduces their structural rigidity and increases frontal surface area. In the short and wide design, the flaps are smooth like the long and thin design but are instead longer going across the body of the rocket than going along the body. This gives these flaps the largest surface area and second most drag of the designs. The short and wide flaps are also the least likely to break because of their length and shape. However, because the flaps are wide across the rocket s body, there is less material keeping the rocket connected, meaning this design will provide the rocket with the least structural support. Pros Table 64: Pros and Cons of Slotted FlapDesigns Cons Produces the most drag Lowest weight Can only be as thick as outer wall of rocket or else the coupler wall must be removed, leaving payload bay open during flight Must be made from a from a tough material or risk snapping off. 129

131 Most complicated Pros Table 65: Pros and Cons of the long and thin flap Cons Small movement near the top translates to a lot of movement lower down on the flap Creates highest structural integrity for the rocket due to a large amount of horizontal space left over for the rocket Much lower structural integrity for the flap due to the length of the flap Wastes the most space Figure 54: Rendering of Short and Wide Flap Design Pros Table 66: Short and Wide Cons Least likely to break due to large amount of material in the fin Least structural support for the rocket Largest surface area Second most drag Options: Long and Thin Short and wide Slotted Table 67: Air Brake Flap Trade Study 130

132 Category Weight Value Score Value Score Value Score Drag 30.00% Weight 5.00% Flap Durability 20.00% Simplicity 15.00% Manufacturability 10.00% Total Score As shown in the trade study, the short and wide flap design won out due to its significantly higher drag, durability, simplicity of design, and manufacturability. Despite the slotted designs higher drag and lower weight, it is significantly less durable and significantly more complex in design and manufacturing which lead to LBR s decision to go with the short and wide design for the flap. 8.8 Derivation of Requirements To make certain that LBR accomplished all the necessary goals to achieve the highest score possible, a list of requirements was established as follows: Requirement Number Table 68: Airbrake Derivation of Requirements Requirement Method of verification 1.1 The vehicle shall deliver the science or engineering payload to an apogee altitude of 5,280 ft. A.1 The air brake shall be implemented in the vehicle to deliver the vehicle to an apogee of 5,280 ft. ± 40ft. Test Flight altimeters will record the apogee of the vehicle during test and competition launches. Test Flight altimeters will record the apogee of the vehicle during test and competition launches. A.1.1 The air brake shall be designed to vary in actuation to allow for a responsive deployment of the air brake as per the control loop. Test A small video camera will be placed in the air brake segment of the rocket to be able to determine if the air brake actuate properly. 131

133 A.1.2 The air brakes control loop shall be able to quickly and accurately respond to errors between target and actual velocity and deploy the air brake accordingly. Test Review of the feedback from the altimeter and accelerometer will determine if the control loop is functioning properly 132

134 Section 9: Project Plan 9.1 Requirement Verification Table 69: General Requirements Minimum Success Requirement 1.1. Students on the team will do 100% of the project, including design, construction, written reports, presentations, and flight preparation with the exception of assembling the motors and handling black powder or any variant of ejection charges, or preparing and installing electric matches (to be done by the team s mentor) The team will provide and maintain a project plan to include, but not limited to the following items: project milestones, budget and community support, checklists, personnel assigned, educational engagement events, and risks and mitigations Foreign National (FN) team members must be identified by the Preliminary Design Review (PDR) and may or may not have access to Verification Method Verify all project design, construction, written reports, presentations, and flight preparation are performed 100% by students. Verify that assembling the motors and handling black powder or any variant of ejection charges, or preparing and installing electric matches is performed by team mentor. Verify the team will maintain a project plan that will include project milestones, budget community support, assigned personnel, educational engagement and risk mitigation. Verify that all foreign nationals are identified before PDR by provided having all team members provide proof of US citizenship. Verification Plan LBR is a 100% studentrun organization. Team leads will ensure all written and physical work is performed by student members of LBR. Safety Officer will ensure assembling the motors and handling black powder or any variant of ejection charges, or preparing and installing electric matches is performed by team mentor. LBR utilizes and accessible team calendar to track assignments due dates and event dates. LBR has also created an online storage filled with budget, personnel assigned, and education engagement event. The safety officer ensures daily that every LBR follows the posted lab safety rules. Team leads will verify every team member's citizenship and submit the necessary documents to the NASA representative. 133

135 certain activities during launch week due to security restrictions. In addition, FN s may be separated from their team during these activities 1.4. The team must identify all team members attending launch week activities by the Critical Design Review (CDR). Team members will include: Students actively engaged in the project throughout the entire year One mentor (see requirement 1.14) No more than two adult educators The team will engage a minimum of 200 participants in educational, hands-on science, technology, engineering, and mathematics (STEM) activities, as defined in the Educational Engagement Activity Report, by FRR. An educational engagement activity report will be completed and submitted within two weeks after completion of an event. A sample of the educational engagement activity report can be found on page 31 of the handbook. To satisfy this requirement, all events must occur between Students must verify to the team advisor that they will be able to attend launch before CDR is submitted. One mentor and adult educator must also commit to the launch week. Verify team will engage a minimum of 200 participants in educational STEM activities. An accurate activity report will be completed and submitted two weeks after the event. Verify for an educational event to count it will be completed between project acceptance and FRR due date. LBR will require that every team member that can attend launch week sign up before CDR is submitted. LBR will keep a record of every student who will be attending launch. Team members are required to attend at least one outreach event to maintain team engagement. Outreach are planned out week by week by the outreach chair who will be in charge of planning each event and submitting each educational engagement report.. 134

136 project acceptance and the FRR due date The team will develop and host a Web site for project documentation Teams will post, and make available for download, the required deliverables to the team Web site by the due dates specified in the project timeline All deliverables must be in PDF format In every report, teams will provide a table of contents including major sections and their respective sub-sections In every report, the team will include the page number at the bottom of the page LBR will host and update their team website Verify LBR will make all necessary documentation available for download on the team website by the due dates of the project. Verify that all deliverables are in PDF format Verify every report provides a table of contents including major sections along with their respective sub-section. Verify in every report the page numbers are included at the bottom of the page. LBR has specified a specific webmaster that will handle all social media and website posts. Team leads will check for the their quality and inform the webmaster what information they want to be posted LBR will prepare and finish all documentation 2 weeks before the due date to ensure everything is posted before their due dates. All links will be tested to ensure that there will be no errors on the due date. Team leads will use PDF reader to compile all documents to a PDF Team leads will structure each of their sections to follow NASA s report guidelines. They will check that any subsection that they add to their system is added in the Table of contents. LBR s PDF compiler will add the page numbers to the bottom of the documents. Team members will check and re read every document to ensure no errors have occurred. 135

137 1.11. The team will provide any computer equipment necessary to perform a video teleconference with the review panel. This includes, but is not limited to, a computer system, video camera, speaker telephone, and a broadband Internet connection. Cellular phones can be used for speakerphone capability only as a last resort All teams will be required to use the launch pads provided by Student Launch s launch service provider. No custom pads will be permitted on the launch field. Launch services will have 8 ft rails, and 8 and 12 ft rails available for use Teams must implement the Architectural and Transportation Barriers Compliance Board Electronic and Information Technology (EIT) Accessibility Standards (36 CFR Part 1194) Each team must identify a mentor. A mentor is defined as an adult who is included as a team member, who will be supporting the team LBR has access to the CSULB college of engineering rooms which have access to all of the computer equipment necessary to perform a video teleconference. LBR launches off standard launch rails that will provided at the FAR launch site. To ensure that standard launch rails are used. Verify that Architectural and Transportation Barriers Compliance Board Electronic and Information Technology are implemented LBR will verify a team mentor who is in possession of an NAR certification and in good standing to handle all of the motors. Who also has a minimum of 2 flights with motor class or higher LBR will get in contact with the college of engineering in advance of the teleconference to reserve the rooms necessary for the conference. LBR will confirm on launch day with the FAR advisor that the launch rails available are of standard sizes. LBR will thoroughly read and acknowledge to the Architectural and Transportation Barriers Compliance Board Electronic and Information Technology (EIT) Accessibility Standards. LBR has identified David Roy as their team mentor. He is currently in possession of an NAR certification and in good standing. David Roy is 136

138 (or multiple teams) throughout the project year, and may or may not be affiliated with the school, institution, or organization. The mentor must maintain a current certification, and be in good standing, through the National Association of Rocketry (NAR) or Tripoli Rocketry Association (TRA) for the motor impulse of the launch vehicle and must have flown and successfully recovered (using electronic, staged recovery) a minimum of 2 flights in this or a higher impulse class, prior to PDR. The mentor is designated as the individual owner of the rocket for liability purposes and must travel with the team to launch week. One travel stipend will be provided per mentor regardless of the number of teams he or she supports. The stipend will only be provided if the team passes FRR and the team and mentor attends launch week in April. and capable of traveling with the team to launch week. also available to attend every launch and go to Huntsville, Alabama with the team. Table 70: Launch Vehicle Requirements Minimum Success Requirement Verification Method Verification Plan 137

139 2.1. The vehicle will deliver the payload to an apogee altitude of 5,280 feet above ground level (AGL) 2.2. The vehicle will carry one commercially available, barometric altimeter for recording the official altitude used in determining the altitude award winner. Teams will receive the maximum number of altitude points (5,280) if the official scoring altimeter reads a value of exactly 5280 feet AGL. The team will lose one point for every foot above or below the required altitude Each altimeter will be armed by a dedicated arming switch that is accessible from the exterior of the rocket airframe when the rocket is in the launch configuration on the launch pad 2.4. Each altimeter will have a dedicated power supply Each arming switch will be capable of being locked in the ON position for launch (i.e. cannot be disarmed due to flight forces) The launch vehicle will be designed to be recoverable and reusable. Reusable is defined as being able to launch again on the same day Based on simulations LBR will be reaching an apogee of 5,400 feet AGL Verify that the launch vehicle will carry one commercially available, barometric altimeter to record the altitude. Verify that each altimeter will be armed by a dedicated arming switch that is accessible from the exterior of the rocket. Verify each altimeter will have a dedicated power supply. Verify each arming switch is capable of being locked in the ON position for launch. Verify the launch vehicle is designed recoverable and reusable. LBR will utilize an air brake system to ensure that the launch vehicle will reach an apogee as close to 5,280 feet as possible. LBR will utilize two Perfectflite StratoLogger Altimeter in order to recorder the launch vehicle's altitude The second recorder will be used for redundancy to verify the first altimeters results. LBR will only be able to arm every altimeter with arming switches located outside of the launch vehicle. LBR will use separate lipo batteries to power each altimeter separately. LBR only purchases arming switches from apogee rockets that are capable of being locked in the ON position. LBR has designed the launch vehicle to be recoverable and reusable. 138

140 without repairs or modifications The launch vehicle will have a maximum of four (4) independent sections. An independent section is defined as a section that is either tethered to the main vehicle or is recovered separately from the main vehicle using its own parachute 2.8. The launch vehicle will be limited to a single stage The launch vehicle will be capable of being prepared for flight at the launch site within 3 hours of the time the Federal Aviation Administration flight waiver opens The launch vehicle will be capable of remaining in launch-ready configuration at the pad for a minimum of 1 hour without losing the functionality of any critical on-board components The launch vehicle will be capable of being launched by a standard 12-volt direct current firing system. The firing system will be provided by the NASA-designated Range Services Provider The launch vehicle will require no external circuitry or special ground support Verify in the design stage that the launch vehicle will have a maximum of 4 independent sections. Verify the launch vehicle will be limited to a single stage. Verify the the launch vehicle is capable of being prepared for flight within 3 hours. All vehicle components will be capable of remaining on the launch pad for 1 hour without losing functionality, The launch vehicle is capable of being launched by a standard 12- volt direct current firing system. Verify the launch vehicle will not require any external circuitry or LBR has designed their launch vehicle to be made of three sections independent sections: propulsion, avionics, and payload bay. LBR is a single stage Aerotech motor. LBR will practice preparing the launch vehicle before official launch dates in order to ensure the launch vehicle is capable of being prepared within the time frame. LBR has implemented lipo batteries that have a battery life of over an hour so launch vehicle can maintain functionality for over an hour. LBR has configured the launch vehicle to be capable of being fired by a standard 12 volt firing system. The same system will also be used at every test launch. LBR will not make use of any external group support system to initiate launch 139

141 equipment to initiate launch (other than what is provided by Range Services) The launch vehicle will use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR), Tripoli Rocketry Association (TRA), and/or the Canadian Association of Rocketry (CAR) Final motor choices must be made by the Critical Design Review (CDR) Any motor changes after CDR must be approved by the NASA Range Safety Officer (RSO), and will only be approved if the change is for the sole purpose of increasing the safety margin Pressure vessels on the vehicle will be approved by the RSO and will meet the following criteria: The minimum factor of safety (Burst or Ultimate pressure versus Max Expected Operating Pressure) will be 4:1 with supporting design documentation included in all milestone reviews Each pressure vessel will include a pressure relief valve that sees the full pressure of the valve that is capable of withstanding the maximum pressure and flow rate of the group support system to initiate launch. Verify that the launch vehicle uses a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) that is approved by the National Association of Rocketry (NAR) and the Tripoli Rocketry Association(TRA). Verify all motor changes will be made before CDR and any changes will only be approved by the NASA representative. Verify all pressure vessels will be approved by the RSO. Verify each vessel will have a factor of safety of 4:1, with a relief valve the full pedigree on display. LBR will use an Aerotech motor that has been approved and certified by the NAR and TRA. LBR also acknowledges that the motor decision has to made by the CRD due date. But if changes were made LBR will reach out to their NASA representative. LRR will not be using pressure vessels on the vehicle. But if that were to change LBR will follow all of the following criteria. 140

142 tank Full pedigree of the tank will be described, including the application for which the tank was designed, and the history of the tank, including the number of pressure cycles put on the tank, by whom, and when The total impulse provided by a College and/or University launch vehicle will not exceed 5,120 Newtonseconds (L-class) The launch vehicle will have a minimum static stability margin of 2.0 at the point of rail exit. Rail exit is defined at the point where the forward rail button loses contact with the rail The launch vehicle will accelerate to a minimum velocity of 52 fps at rail exit All teams will successfully launch and recover a subscale model of their rocket prior to CDR. Subscales are not required to be high power rockets The subscale model should resemble and perform as similarly as possible to the Verify that the total impulse of the vehicle will not exceed 5,120 Newton-seconds. Verify the launch vehicle will have a minimum static stability margin of 2.0. Verify that the launch vehicle will accelerate to a minimum velocity of 52 fps at rail exit. Verify a subscale rocket will be launched that should resemble the full rocket prior to CDR. LBR s launch vehicle will have a total impulse of 4280 Newton- seconds. If any changes were to be made the team leads will adhere to this criteria. LBR s will have a static stability margin of 2.6 according to calculations. Team leads have been made aware of this requirements and any changes made to the vehicle must adhere to this criteria. LBR s launch vehicle will have an velocity of 66.9 ft/s at rail exit based on calculations. Team leads have been made aware of this requirements and any changes made to the vehicle must adhere to this criteria. LBR will launch a subscale rocket on November 18th in order to launch before the CDR due date. 141

143 full-scale model, however, the full-scale will not be used as the subscale model The subscale model will carry an altimeter capable of reporting the model s apogee altitude All teams will successfully launch and recover their full-scale rocket prior to FRR in its final flight configuration. The rocket flown at FRR must be the same rocket to be flown on launch day. The purpose of the full-scale demonstration flight is to demonstrate the launch vehicle s stability, structural integrity, recovery systems, and the team s ability to prepare the launch vehicle for flight. A successful flight is defined as a launch in which all hardware is functioning properly (i.e. drogue chute at apogee, main chute at a lower altitude, functioning tracking devices, etc.). The following criteria must be met during the full-scale demonstration flight: The vehicle and recovery system will have functioned as designed The payload does not have to be flown during the full-scale test flight. The following requirements still apply: If the payload is not flown, mass simulators will be used to simulate the payload mass The mass Verify LBR will launch a full scale rocket before FRR in its final flight configurations successfully. Verify that if the payload is not flown mass simulators will be used to simulate mass in the approximate location the payload will be kept in. LBR will work towards having a successful launch two full weeks prior to FRR. To plan in the event that there is a failure in the full scale launch. Also if the payload is not ready by the time of the full scale launch LBR will use mass simulators to simulate the payload. 142

144 simulators will be located in the same approximate location on the rocket as the missing payload mass If the payload changes the external surfaces of the rocket (such as with camera housings or external probes) or manages the total energy of the vehicle, those systems will be active during the fullscale demonstration flight The full-scale motor does not have to be flown during the full-scale test flight. However, it is recommended that the fullscale motor be used to demonstrate full flight readiness and altitude verification. If the full-scale motor is not flown during the full-scale flight, it is desired that the motor simulates, as closely as possible, the predicted maximum velocity and maximum acceleration of the launch day flight The vehicle must be flown in its fully ballasted configuration during the fullscale test flight. Fully ballasted refers to the same amount of ballast that will be flown during the launch day flight. Additional ballast may not be added without a reflight of the full-scale launch vehicle After successfully completing the full-scale Verify if the payload changes the external surfaces of the rocket or manages the total energy those systems will be active during flight of the full scale rocket. Verify that the full scale motor is not required during the full scale test flight. Verify that the launch vehicle will be flown in its fully ballasted configuration during the full scale flight test. Verify after the full scale flight no components will be modified LBR s payload does not changes the external surface of the rocket nor manages the energy but if any changes were LBR will follow these criteria. LBR plans to use the full scale motor during the full scale flight test. However if any changes were necessary and the full scale motor is not used LBR will plan to use a motor as close to the actual as possible. LBR will be launching their launch vehicle in its fully ballasted state during the full scale test flight. After the successful full scale demonstration flight 143

145 demonstration flight, the launch vehicle or any of its components will not be modified without the concurrence of the NASA Range Safety Officer (RSO) Full scale flights must be completed by the start of FRRs (March 6th, 2018). If the Student Launch office determines that a re-flight is necessary, then an extension to March 28th, 2018 will be granted. This extension is only valid for re-flights; not first-time flights Any structural protuberance on the rocket will be located aft of the burnout center of gravity Vehicle Prohibitions The launch vehicle will not utilize forward canards The launch vehicle will not utilize forward firing motors The launch vehicle will not utilize motors that expel titanium sponges (Sparky, Skidmark, MetalStorm, etc.) The launch vehicle will not utilize hybrid motors The launch vehicle will not utilize a cluster of motors The launch vehicle will not utilize friction fitting for motors The launch vehicle will not exceed Mach 1 at any point during flight Vehicle ballast will not without the concurrence of the NASA RSO. Verify all full scale flights must be completed by the start of FRRs but if necessary for a re flight an extension will be granted Verify any structural protuberance of the rocket will be located aft if the burnout center of gravity. Verify that all of the vehicle prohibitions will be followed by LBR. any modifications to any of the launch vehicle's components will not be made without the concurrence of the NASA RSO. LBR plans to have all full scale flights completed by FRRs. If a re-flight is scheduled LBR has made plans to account for this in their scheduling. LBR s airbrake system will protrude aft of the burnout center of gravity. LBR has followed every one the vehicle prohibitions when designing the rocket. If any changes are made with the launch vehicle LBR will be careful to not violate any of these verifications. 144

146 exceed 10% of the total weight of the rocket Table 71: Recovery Requirements Minimum Success Requirement 3.1. The launch vehicle will stage the deployment of its recovery devices, where a drogue parachute is deployed at apogee and a main parachute is deployed at a lower altitude. Tumble or streamer recovery from apogee to main parachute deployment is also permissible, provided that kinetic energy during drogue-stage descent is reasonable, as deemed by the RSO Each team must perform a successful ground ejection test for both the drogue and main parachutes. This must be done prior to the initial subscale and full-scale launches 3.3. At landing, each independent sections of the launch vehicle will have a maximum kinetic energy of 75 ft-lbf Verification Method Verify that the launch vehicle will stage the deployment of its recovery devices. Verify that each team will perform a successful ground ejection test for both the drogue and main parachutes before subscale and full-scale launches. Verify that at landing, each independent sections of the launch vehicle will have a maximum kinetic energy of 75 ft-lbf Verification Plan The launch vehicle will utilize a dual deployment recovery system to execute a successful deployment of the drogue parachute at apogee (5,280-ft) and the main parachute at 500- ft. LBR will utilize black powder ground ejection charge testing (primary and backup) to ensure the drogue and main parachutes successfully eject from the drogue and main airframes. Ground ejection testing will be conducted prior to every subscale and full-scale launch to ensure maximum reliability. Using known equations for descent velocity, drift distance and kinetic energy, the recovery system will utilize a 20 drogue with a 96 main that will create a kinetic energy less than the maximum allowed, ensuring that each 145

147 3.4. The recovery system electrical circuits will be completely independent of any payload electrical circuits All recovery electronics will be powered by commercially available batteries The recovery system will contain redundant, commercially available altimeters. The term altimeters includes both simple altimeters and more sophisticated flight computers Motor ejection is not a permissible form of primary or secondary deployment Removable shear pins will be used for both the main parachute compartment and the drogue parachute compartment Recovery area will be limited to a 2500 ft. radius from the launch pads Verify that the recovery system electrical circuits will be completely independent of any payload electrical circuits. Verify that all recovery electronics will be powered by commercially available batteries. Verify that the recovery system will contain redundant, commercially available altimeters. Verify that the motor ejection is not a permissible form of primary or secondary deployment. Verify that removeable shear pins will be used for both the main parachute and the drogue parachute compartments. Verify that the launch vehicle will descend and land within the 2,500-ft recovery area. section of the launch vehicle does not exceed a kinetic energy of 75-ft-lbf. The launch vehicle utilizes its own Single Avionics Bay that completely separate from the Payload Bay, separating both electronics and circuits from each other entirely. The recovery system electronics will utilize a standard 9V commercial battery as its main power source. The launch vehicle avionics bay will house two PerfectFlite StratoLoggerCF altimeters to maintain redundancy and reduce the risk of recovery failure. The launch vehicle motor will not be used as a form of primary or secondary deployment. The launch vehicle s main and drogue parachute airframes will utilize four ⅛ shear pins on both the front and aft ends of the Avionics Bay as the nylon pins for the separation events. Utilizing known equations on drift distance and descent velocity and with the 20 drogue and 96 main being used, the launch vehicle will recover under the specified 146

148 3.10. An electronic tracking device will be installed in the launch vehicle and will transmit the position of the tethered vehicle or any independent section to a ground receiver Any rocket section, or payload component, which lands untethered to the launch vehicle, will also carry an active electronic tracking device The electronic tracking device will be fully functional during the official flight on launch day The recovery system electronics will not be adversely affected by any other on-board electronic devices during flight (from launch until landing) The recovery system altimeters will be physically located in a separate compartment within the vehicle from any other radio frequency transmitting device and/or magnetic wave producing device The recovery system electronics will be shielded from all onboard transmitting devices, to avoid inadvertent excitation of the recovery system electronics The recovery system electronics will be shielded from all Verify that an electronic tracking device will be installed in the launch vehicle and will transmit the position of the tethered vehicle or any independent section to a ground receiver; any separate section that lands untethered to the launch vehicle will also carry an electronic tracking device and each electronic tracking device will be fully functional during flight. Verify that recovery system electronics will not be adversely affected by any other on-board electronic devices during flight; verify the altimeters will be separate from any frequency transmitting device and that all recovery system electronics are shielded from any transmitting device or magnetic wave generating devices. Verify that the recovery electronics will not be adversely affected by other onboard devices. recovery radius of 2500-ft. through the drogue and main parachute deployments. The launch vehicle s Avionics Bay will utilize the Big Red Bee BRB900 electronic GPS transmitter device and receiver to transmit the position of any tethered vehicle or independent section of the launch vehicle. The current design of the launch vehicle yields no untethered or independent sections during flight. The payload will deploy the rover after landing. The BRB900 electronic GPS device will undergo multiple ground testing with recorded data to ensure reliability during flight. The Avionics Bay will have its own separate section on the launch vehicle with dedicated electronics to ensure that it is not affected by other on-board electronic devices during flight. The StratoLoggerCF altimeters will be located within the avionics bay on the front wooden housing tray on the top side of the 3D printed avionics center casing, ensuring separation from any the BRB900 GPS system, which will be attached on the lower wooden housing tray, separated by the avionics center casing, maintaining separation between the altimeters and frequency transmitting devices. Since the Avionics Bay internal structure is separated by a 3D printed 147

149 onboard devices which may generate magnetic waves (such as generators, solenoid valves, and Tesla coils) to avoid inadvertent excitation of the recovery system The recovery system electronics will be shielded from any other onboard devices which may adversely affect the proper operation of the recovery system electronics. center housing case, the recovery electronics will not be affected by any onboard transmitting devices. The avionics bay will be its own separate section and has its own dedicated recovery system electronics in the airframe, removing any devices that may generate magnetic waves. Since the avionics bay will be its own separate section any other onboard devices which may affect the proper operation of the recovery system electronics will be physically located in a separate section of the launch vehicle.. Table 72: Deployable Rover Requirements Minimum Success Requirement Teams will design and manufacture a custom rover that will deploy from the internal structure of the launch vehicle At the launch vehicle s landing site, the team will remotely activate a trigger to deploy the rover from the rocket. Verification Method A threaded rod will push the nose cone to create an opening on the rocket s structure. The threaded rod will also be used to push the rover out of the rocket for deployment. When the launch vehicle lands, a team member will operate a transmitter to send a signal to the launch vehicle for the rover deployment mechanism to activate. Verification Plan The rover will sit on a threaded rod that runs through the internal structure of the launch vehicle and the center of the rover. When the rod turns, it will unthread the rover along with the nose cone. The rover will then fall out of the airframe. An RC transmitter will trigger the rover deployment mechanism. Initially, the nose cone, coupler, and airframe will separate. Then, the coupler will retract creating an opening between the nose cone and coupler. The rover will deploy from the coupler and then fall into this empty space so that it can drive. 148

150 After deployment, the rover will autonomously move at least 5 ft. (in any direction) from the launch vehicle Once the rover has reached its final destination, it will deploy a set of foldable solar cell panels. An all terrain rover will move at least 5 ft away from the launch vehicle. The rover will have solar panels that will open in half. The team will utilize sensor fusion of GPS and IMU to track the distance traveled by the rover. The solar panels will be built into the sides of the rover. A single servo will open the side flaps to reveal the solar panels. Table 73: Safety Requirements Minimum Success Requirements Verification Method Verification Plan 5.1. Each team will use a launch and safety checklist. The final checklists will be included in the FRR report and used during the Launch Readiness Review (LRR) and any launch day operations Each team must identify a student safety officer who will be responsible for all items in section The role and responsibilities of each safety officer will include, but not limited to: Monitor team activities with an emphasis on Safety during: Design of vehicle and payload Construction of vehicle and payload Assembly of vehicle and payload Ground testing of vehicle and payload Sub-scale launch test(s) Full-scale launch test(s) Launch day LBR will create safety checklist that will be included in FRR report and Launch Readiness Review. Shawn Everts is the official Safety Officer for LBR. LBR Safety Officer will be responsible for all subsections listed. LBR Safety Officer with ensure that the safety checklist are complete. They will also verify that safety checklist is used during the launch day operations. He will assure that the team adheres to all regulations pertaining to the construction, assembly, testing, flight, and recovery phases of the launch vehicle. LBR Safety Officer will be present during the design, construction, assembly, ground testing, sub-scale launch, full-scale launch, launch day, recovery and education engagement activities. The Safety Officer will be monitoring all these activities for safety concerns. 149

151 Recovery activities Educational Engagement Activities Implement procedures developed by the team for construction, assembly, launch, and recovery activities Manage and maintain current revisions of the team s hazard analyses, failure modes analyses, procedures, and MSDS/chemical inventory data Assist in the writing and development of the team s hazard analyses, failure modes analyses, and procedures During test flights, teams will abide by the rules and guidance of the local rocketry club s RSO. The allowance of certain vehicle configurations and/or payloads at the NASA Student Launch Initiative does not give explicit or implicit authority for teams to fly those certain vehicle configurations and/or payloads at other club launches. Teams should communicate their intentions to the local club s President or Prefect and RSO before attending any NAR or TRA launch 5.5. Teams will abide by all rules set forth by the FAA. All members of the Long Beach Rocketry team will be given a safety briefing and are required to sign a safety contract. LBR Safety Officer will be responsible for maintaining the team s hazard analyses, failure modes analyses, procedures, and MSDS/chemical inventory data. LBR Safety Officer will be responsible for writing the team s hazard analyses, failure modes analyses, procedures, and MSDS/chemical inventory data. All members of the Long Beach Rocketry Team will follow the range safety regulation. LBR Safety Officer will be responsible for These briefings will cover proper procedures for construction, assembly, launch and recovery activities. LBR Safety Officer will be briefed on all components of the launch vehicle from each subsystems lead to be able to properly maintain hazard analyses and failure mode analyses. LBR Safety Officer will have a understanding of all components of the launch vehicle to be able to properly write the hazard analyses and failure mode analyses. LBR will follow the range safety regulation as stated in the Long Beach Rocketry Team Safety Agreement that all members have to read and sign to be able to participate in the project. LBR Safety Officer has created LBR Team 150

152 verifying that all rules set forth by the FAA or followed. Safety Agreement which is sign by all members saying that they have read and understood all rules set forth by the FAA. 151

153 9.2 Timeline Figure 55: Gant Chart for Competition 152

154 Figure 56: Gant Chart for Airbrake Table 74: Expected Development Schedule Competition Timeline Task Proposal PDR Subscale Launch CDR Full-Scale Launch FRR Competition Expected Date of Completion 9/20/17 11/3/17 11/4/17 1/12/18 2/3/18 3/5/18 4/4/18 153