NASA Student Launch Initiative Flight Readiness Review

Transcription

1 NASA Student Launch Initiative Flight Readiness Review Ames, IA CySLI: Cyclone Student Launch Initiative March 6 th, 2017 Table of Acronyms

2 Acronym CySLI CFD FEA RCS DOF IMU MAC OD ID Definition Cyclone Student Launch Initiative Computational Fluid Dynamics Finite Element Analysis Roll Control System Degree of Freedom Inertial Measurement Unit Mean Aerodynamic Chord Outer Diameter Inner Diameter NASA SL FRR

3 Table of Contents 1)Summary of FRR Report 1.1)Team Summary 1.2)Launch Vehicle Summary 1.3)Avionics Summary 2)Changes Made Since CDR 2.1)Administrative 2.2)Rocket Team 2.3)Project Plan 2.4)Avionics Team 3)General Information 3.1)Technical Advisors/Educators 3.1.1)NAR/TRA Team Mentor 3.1.2)Technical Advisor 3.2)Program Advisor 3.3)Team Advisor 3.4)Safety Officer 3.5)Project Team Lead 3.6)Team Roster 3.7)TRA Section Affiliation 4)Systems Requirements 4.1)Vehicle Requirements 4.2)Recovery Requirements 5)Mission Performance Prediction NASA SL FRR

4 5.1)Mission Performance Criteria 5.2)Verification Plan 5.3)Confidence and Maturity of Design 5.4)Quality of Workmanship 6)Full Scale Launch Test Results 7)System and Subsystem Overview 7.1)Rocket Overview 7.2)Star CCM Models 7.3)Nose Cone 7.4)Camera Bay 7.5)Parachute Bay 2 7.6)Parachute Bay 1 7.7)Flight Computer Bay 7.7.1)Altimeters 7.7.2)Apogee Control System )Apogee Prediction )Kalman Filter Performance Predictions 7.8)Roll Induction 7.9)Air Brakes 7.10)Servo 7.11)Hinges 7.12)Pulley 7.13)Exterior Fins 7.14)Motor/Motor Mount 7.15)Fiberglass Fins 7.16)Mass Statement NASA SL FRR

5 8)Recovery Systems 8.1)Recovery Overview 8.2)Kinetic Energy Management 8.3)Instron Test 8.4)Black Powder Calculations 8.5)Recovery Electronics 8.5.1)Redundancy Features 8.6)Drift 9)Safety and Environment 9.1)Risk Assessment 9.2)Risk Assessment of the Rocket 9.3)Risk Assessment of Avionics 9.4)Risk Assessment of Lab Work 9.5)Risk Assessment of Environmental Hazards 10)Launch Operations Procedures 10.1)Launch Procedure )Launch Procedure Checklist )Avionics Checklist 10.2)Safety and Quality Assurance )Launch Operation Risks )Launch Operation Risk Assessment )Launch Operations Environmental Concerns 10.3)Environmental Concerns/Mitigations 10.4)Operations Manager 11)Project Plan 11.1)Budget NASA SL FRR

6 11.2)Educational Engagement )First Lego League )Science Bowl 12)Conclusion 1) Summary of FRR Report 1.1 Team Summary Team Name: CySLI (Cyclone Student Launch Initiative) Mailing Address: Department of Aerospace Engineering, 2271 Howe Hall, Ames, IA Mentor: Gary Stroick TRA Number: 5440 Certification Level: Level Launch Vehicle Summary The launch vehicle consists of Blue Tube, aluminum, birch plywood, fiberglass, carbon fiber, and other additional 3D printed materials. The vehicle is designed with efficiency as well as performance in mind, while also remaining lightweight and easy to assemble and disassemble. The Apogee Management System will control the projected apogee of the rocket in order to meet the 5,280 ft. apogee requirement set by NASA. This system consists of a set of four air brakes located aft of the rocket fins at the circumferential midpoint of each fin slot. These air brakes will be controlled by the flight computer, which will deploy the air brakes at calculated intervals in order to asymptotically approach an apogee of 5,280 ft. Launch Vehicle Overview: Diameter: 6 Length: 132 Total Weight: 50 lb. Final Motor Choice: Cesaroni 5015L1115-P Classic Recovery System: Three sections under a 120 main parachute Rail Size: 12 ft 1515 Standard Rail NASA SL FRR

7 1.3 Avionics Summary In order to achieve the required apogee, we will have to have reliable data input for the prediction algorithm. To do this, data filtering is required. We want the most accurate estimations possible to minimize cumulative error effects. A Kalman filter can achieve the accuracy we need. From the prediction, we will determine the duration of airbrake deployment. For our recovery system we will use two redundant altimeters to ensure correct apogee detection and successful rocket separation. 2) Changes Made Since CDR 2.1 Administrative No administrative changes made since the CDR. 2.2 Rocket Team CDR 2-56 Shear pins FRR 4-40 Shear pins subject to second black powder test Reason 2-56 shear pins led to main chute deployment at apogee during full scale test launch. 2.3 Project Plan CDR Complete two full scale test launches prior to FRR. FRR Complete one full scale test launch prior to FRR. Reason Weather 2.4 Avionics Team CDR Flight computer: Nucleo STM32 FRR Flight computer: Arduino Due Reason Ease of use and programming with other electronics used NASA SL FRR

8 3) General Information 3.1 Technical Advisers/Educators NAR/TRA Team Mentor Gary Stroick TRA Number: 5440 Certification Level: Level 3 Phone: (952) president@offwegorocketry.com Website: Gary Stroick holds Bachelors of Sciences in Mathematics and Computer Science, Masters of Science in Computer Science, and a Masters of Business Administration. Involved in highpowered rocketry since 1997, he holds a Tripoli Rocketry Association (TRA) Level 3 certification, in addition to a Low Explosives Import/User Permit. He was Vice President of Tripoli Minnesota from 2007 to 2011 and has been President since He owns Off We Go Rocketry, a business in which he provides rocket hobbyist, commercial, and military organizations with high powered rocket motors and supplies, as well as advises fliers on various aspects of rocketry and certifications Technical Adviser Clayton C. Anderson NASA Astronaut (Former) Distinguished ISU Faculty Fellow - Aerospace Engineering Phone: (515) isuastro@iastate.edu NASA SL FRR

9 Clayton C. Anderson is the first alumnus to become an astronaut. A Nebraska native, he earned his BS in physics at Hastings College. After a summer internship at NASA s Johnson Space Center, he began his MS work at Iowa State in aerospace engineering. Anderson joined NASA full time in the Mission Planning and Analysis Division following his graduation in Over the next 15 years, his responsibilities included designing rendezvous and proximity operations trajectory designs for early space shuttle and space station missions and leading the trajectory design team for the Galileo planetary mission. Anderson was accepted into the astronaut-training program in As an astronaut, he led the development of the Enhanced Caution and Warning System to aid astronauts in diagnosing and correcting problems that occur during space flight. In June 2007, Anderson began a 152-day mission aboard the International Space Station. His duties included three spacewalks totaling 18 hours to prepare the station for additional construction. His mission ended successfully on November 7, He is now a distinguished faculty fellow at, helping students learn and grow in the aerospace engineering department. 3.2 Program Adviser Matthew Nelson Make to Innovate (M:2:I) Program Coordinator - Aerospace Engineering Phone: (515) mnelson@iastate.edu 3.3 Team Adviser Dr. Jonathan Regele Assistant Professor - Aerospace Engineering Phone: (515) jregele@iastate.edu 3.4 Safety Officer(s) Bradlee Fair Phone: (309) bjfair@iastate.edu 3.5 Project Team Leader Austin Kaiser Phone: (312) NASA SL FRR

10 3.6 Team Roster Membership Count: 15 Team members are responsible for the design and construction of the competition rocket. This includes, but is not limited to, rocket recovery, power, stability, flight computer programming, electronics, altimeter bay electrical hardware, material selection, and manufacture of the rocket. A full roster is depicted in Figure 1 below. NASA SL FRR

11 Figure 1: Team Structure 3.7 TRA Section Affiliation For assistance, mentorship, and review, CySLI will be associating with the Tripoli Minnesota High Power Rocketry Club. They are prefecture #45 of the Tripoli Rocketry Association and located in North Branch, Minnesota. NASA SL FRR

12 4) System Requirements 4.1 Vehicle Requirements The vehicle will launch with a target apogee of 5,280 ft. and a small tolerance level. Total weight will be less than or equal to one-third of the certified average thrust produced by the rocket motors. We will achieve apogee based on a precisely pre-measured mass value combined with live flight data fed to the microcontroller. This controller will send commands to the air brake system that will correct trajectory. Upon descent, the rocket body will split into three segments, releasing the drogue and main parachute. The drogue will deploy at apogee, slowing vehicle descent until arrival at 800 ft. Above Ground Level (AGL), where the main parachute will deploy. 4.2 Recovery Requirements Two commercially available and certified altimeters, through redundancy, will ensure black powder charge detonation and deployment of all parachutes distributed between the descending sections. Impact energy for each section shall remain below 75 ft-lb. The recovery system uses shock cord lengths that will mitigate the risks of tangling or accidental interaction between each independent section, which could result in damage to the vehicle as well as a safety risk to those on the ground. 5) Mission Performance Predictions 5.1 Mission Performance Criteria The mission performance criteria for the rocket and payload are as follows: Coast: The rocket shall maintain stability throughout its entire flight. The Apogee Targeting System will use air brakes if deemed necessary to allow rocket to reach 5280 ft +/- 50 ft Onboard video shall be taken of the airbrakes as a verification method of their duty cycle Recovery: The altimeter bay shall deploy its drogue parachute at apogee and slow its descent to allow impact energies below supplied limit The altimeter bay shall deploy both main parachutes successfully Descent under main parachute shall be such that the energies are below 75 ft-lb Upon landing, the rocket shall be fully recoverable and able to be launched again within 1 hour In order to have a successful mission the following criteria are necessary: NASA SL FRR

13 Achieve an apogee of 5280 feet by braking when necessary Deploy drogue Deploy main parachute Achieve video of airbrakes Land safely 5.2 Verification Plan Requirement The rocket shall not exceed an apogee altitude over 5,600 feet AGL. Solution The motor is rated to an impulse of 5,015 N*s and according to OpenRocket cannot possibly exceed an altitude of 5,554 ft. with the current mass budget. The rocket is built to fly and descend safely and upon landing be reusable with only a new motor reload and an electronic systems reset. Verification Openrocket simulations, subscale test launch and full scale test launch data will provide verification that the rocket will not surpass 5,600 ft. altitude. Static tests will be done for body reconstruction, electronics simulations, and ejection charges. A full scale test launch will be done to proof the rocket s reusability as a whole. The launch vehicle shall be capable of being prepared for flight at the launch site within 2 hours, from the time the Federal Aviation Administration flight waiver opens. The design is focused on simple, modular integration to allow for orderly, proper assembly within a short period of time Keep in mind assembly and integration during design phase. Test runs will be done to ensure the system can be correctly assembled within allotted timeframe. The launch vehicle shall be capable of being launched by a standard 12-volt direct current firing system. The firing system will be provided by the NASAdesignated Range Services Provider We will be designing around a standard igniter for the motor, which will require a 12v power source. Tests with voltmeter will be done to check output voltage around igniter location once fabrication is complete. Wiring schematic is built around a 12v power source. The launch vehicle shall use a commercially available solid motor propulsion system The rocket will be using a Cesaroni 5015-L1115-CS-P Motor which has been Documentation can be found on Cesaroni s website. The launch vehicle shall be designed to be recoverable and reusable. Reusable is defined as being able to launch again on the same day without repairs or modifications. NASA SL FRR

14 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). The total impulse provided by a launch vehicle shall not exceed 5,120 Newtonseconds (L-class). The subscale model should resemble and perform as similarly as possible to the full-scale model, however, the full-scale shall not be used as the subscale model. All teams shall successfully launch and recover a subscale model of their fullscale rocket prior to CDR. At landing, each independent section of the launch vehicle shall have a maximum kinetic energy of 75 ft-lbf. Air Brakes do not alter rocket s CP out of line with rocket s longitudinal axis designed to meet NAR/TRA rules and regulations. The total impulse of the selected rocket motor, a Cesaroni 5015-L1115-CS-P, is 5,015 N*s which is below the designated limit. The subscale model is designed to be dimensionally 1:3 scale to the full size model. Additionally, the impulse of the selected motor is approx. 1:3 of the full scale. A 1:3 subscale rocket will be built with respect to the dimensions but not to the impulse, apogee, and weight to ensure accurate stability. Upon landing, the rocket will be under a 120 main, 24 pilot, and 18 drogue parachute. This arrangement ensures steady descent velocities that will result in kinetic energies less than 75 ft-lbf for each section. All of the airbrakes will be attached to a single servo so if there is a failure none of the airbrakes will be deployed, avoiding the change in CP and drifting of the rocket. If one line Documentation can be found on the manufacturer s website, which is assumed to be accurate. Sub-scale test launch data provided by onboard altimeter(s) will verify accuracy of OpenRocket drag coefficient predictions, as well as performance scaling based on rocket dimensions. Simulations will be run to ensure the goals are met to replicate the stability and aerodynamics as closely as possible to the full scale vehicle. Kinetic energy is equal to 1/2mv^2. As shown in the calculations in the recovery section, the velocities will be at reasonable speeds so that the kinetic energies are below 75 ft-lbs. Testing will consist of applying uneven loads to the airbrakes to determine if they still function under these circumstances. NASA SL FRR

15 breaks, opposite airbrakes will not actuate and not affect the rocket s stability. 5.3 Confidence and Maturity of Design The design has gone through multiple iterations of redesign and has evolved into a reliable, arguably simple system that should theoretically satisfy all of the competition requirements. For the rocket and electronics systems, technology/methods have been selected that have been proven and tested by the predecessors and by other sources. This approach ensured the building of a reliable system with a low probability of failure. The less variables introduced to the system, the lower the likelihood of one of them causing a problem. We are confident that our previous designs have set us up for a greater degree of success. For the airbrakes, the flaps are deployed by utilizing cables routed to the outside of the rocket body from a servo near the flight computer. While the method is not traditional, the calculations provide a degree of confidence that the system will function as intended. The large amount of research put into the design grants it relatively high expectations and trust in its projected performance. 5.4 Quality of Workmanship We worked primarily in the Iowa State Boyd and Make to Innovate labs with quality, inspected equipment. These labs always have multiple laboratory technicians on duty to ensure that proper safety procedures are followed. Every team member working on the construction of the Rocket completed shop, fire, and protective equipment training courses prior to working in the labs, complying with Iowa State s Environmental Health and Safety standards. In addition, team members who wished to work with powered-equipment had to complete an equipment-specific training for each type of equipment. To more efficiently manage training time, training for certain pieces of equipment was performed only by certain individuals who would then perform all necessary tasks with that particular piece of equipment. This helped cut down on excessive time spent learning to use equipment and also helped foster increased levels of proficiency in the equipment experts. To ensure the team members with the proper skills were available when needed, s were always sent out ahead of time detailing the planned construction for the day. NASA SL FRR

16 Figure 2: M:2:I lab Throughout the construction of all of the components, we maintained a philosophy of marginally overbuilding to provide room for error. Whenever possible, we d always make cuts with a little extra material left attached to the desired product piece, and we would always drill holes slightly smaller than desired. We would then sand the piece down to the desired size and shape; sanding is a powerful way to control error as it removes material very slowly. Some specific examples of this philosophy at work includes, but isn t limited to: the bay bulkheads; the motor-housing attachment rings, the fin slots, and the payload bay foam. We also took care to select proper sand paper grits. The initial removal of materials using coarse sandpaper was always followed up with a finish with fine sandpaper. We recognized that the thickness of pencil and marker markings inherently lends itself to inaccuracy. To remedy this, we took meticulous care to ensure that the actual location of every cut is represented by the left edge of the marking. We d then cut along the right edge of said marking to allow for the margin of error mentioned in the previous paragraph. With measurements, we also recognized the issue of ensuring straight measurements as diagonal measurements will result in inaccurate markings. Whenever possible, we would measure along known straight edges as a guideline. Otherwise, we would extend out guidelines from the nearest available straight edge to use as references. Prior to any and all build procedures, team members were expected to check with one of the team leads to verify the plan. Most build tasks were planned and delegated by the team leads with team members providing frequent feedback to ensure the best course of action is followed. 6) Full-Scale Launch Test Results Date: February 25, 2017 Location: North Branch, MN NASA SL FRR

17 Weather: Partly Cloudy, 34F, mph winds We set up in North Branch, MN on the day of our launch to get in at least one flight. We only got in this one flight due to time constraints and weather so we only had this one attempt. We had to launch soon so that we could launch in AL. As for the test launch, the launch performed flawlessly on ascent despite high level winds. Our first ejection charge for the drogue chute at apogee deployed as expected but the main chute also deployed. This was because the shear pins were not strong enough. A second round of ejection testing will be performed to ensure that we can strengthen these shear pins while maintaining parachute deployment in the current stages. The main section landed safely but was carried away by strong winds and was dragged through a corn field by the parachutes. The winds were very strong and ended up causing cosmetic damage on the airframe When listening in to the Aim USB s, we recorded a height of 4,357 ft. Come competition, we plan on having all of these issues fixed to launch a second launch safely and successfully. Since we won t be using air brakes, we plan on adding and subtracting weight to compensate for a mile high apogee. Figure 3: Velocity vs. Time Above in figure 3, we have our velocity vs. time graph for our test launch. The velocity is in ft/s. The max velocity of 567 ft/s was expected based off of our simulation results in RockSim as well as in OpenRocket. NASA SL FRR

18 Figure 4: Smoothed Altitude and Velocity vs. Time In figure 4, we have the smoothed altitude measured from the Aim USB s. Because there were high winds in excess of 20 mph, our apogee was cut short and measured 4,357 ft. Come competition, we will remove excess weight from the flight computer bay based on the flight conditions on launch day. We will use our RockSim software to compensate the weight since we will not be using air brakes. Figure 5: Altitude AGL vs. Time 7) System and Subsystem Review NASA SL FRR

19 7.1 Rocket Overview Figure 6: Rocket Rendering Figure 7: Rocket Sectional View The rocket will consist of a nose cone and 4 main body sections (see Figure 6 and 7). The diameter of the rocket was chosen to be 6 inches to account for the camera system in the top. The total length of the rocket will be inches and was chosen to provide space for the three couplers, a motor mount, and two parachute bays. Figure 8: Rocket Component Naming Convention The dimensions of the rocket body were determined by the space needed to comfortably house all of the necessary components. We chose the 6 inch diameter blue tube to accommodate the electronics (cameras, avionics, and flight computers) which will be mounted on sleds within couplers. The length of the rocket is determined by the space needed for the parachutes, air brakes, engine block, and other components. The nose cone adds on an additional 30 inches to the front of the rocket. NASA SL FRR

20 Figure 9: Side View of Constructed Rocket The main body sections will be joined together with three couplers, housing the camera electronics, avionics and flight computers respectively. Figure 8 above demonstrates the locations of the main body sections and couplers referenced throughout this document. The forward main body section will be joined to the nose cone shoulder by the payload coupler. The payload coupler will house the camera system. Figure 10: Rocket Open 3D Model The two main body sections aft of the camera bay will house the rocket s recovery parachutes and will be located on either side of the avionics bay. The parachute bay aft of the avionics bay will house the drogue chute, while the parachute bay forward of the avionics bay will house the main parachute. The avionics bay will house the altimeters used in the recovery deployment sequence. Directly aft of parachute bay two will be the flight computer bay, which will house the electronics for the apogee control system and roll induction system, as well as a portion of the air brakes hardware. The aft body tube section will contain the motor mount, fins, and air brake assembly. The motor mount airframe will house the motor tube and centering rings designed to transfer the motor thrust to the outside airframe of the rocket. NASA SL FRR

21 Figure 11: Rocket Performance Specs Diagram 7.2 Star CCM models Figure 12: Figure 13: Pressure NASA SL FRR

22 20.23 lbs (~90 N) Worst case scenario For four air brakes lbs (360 N) Maximum torque of the servo Velocity Maximum velocity for testing 580 ft/s (176.8 m/s) 7.3 Nose Cone We will be using a filament wound fiberglass nose cone with a threaded aluminum tip, as seen in Figure 15. Figure 14: Fiberglass nose cone modeled in SolidWorks Figure 15: Fiberglass nose cone It will have an approximate 6-inch diameter, a 6-inch shoulder and will be 30 inches long, which in turn will produce a 5:1 ogive planform. The 5:1 ogive will shift the center of gravity of the rocket forwards and ahead of the roll control fins. We selected filament wound fiberglass as the material primarily because of negative prior experience with ordinary fiberglass nose cones. In previous years, fiberglass nose cones have cracked or fractured near the tip, likely due to the manufacturing process. The choice of a filament wound nose cone provides a more robust construction; having an aluminum tip adds extra protection during transport and recovery. 7.4 Camera Bay NASA SL FRR

23 At the forward end of the rocket, we plan to employ a 360-degree field of view camera mount system that looks out radially in all directions. The design of this camera system will consist of 5 Mobius HD cameras that look out at equal parts of the horizon. By stitching each camera s video in post-processing software, we can create a 360-degree field of view video. Figure 16: 3D Printed Camera Bay components Along with the 5 cameras for the 360-degree video, we also will have 2 other Mobius HD cameras that look out radially on opposite sides, but will be looking down the airframe of the rocket due to a mirror housing epoxied on the outside of the rocket that directs the light by 45 degrees. These videos will be used for confirming air brake actuation and roll induction rotation for our rocket. 7.5 Parachute Bay 2 The drogue parachute in the recovery configuration will be packed in parachute bay two between the midsection and tail end section of the rocket. This bay will have a length of 20 and a diameter of about 6 inches, creating enough space to house the drogue as well as the 3 quick-links, 8.6 yards of nylon shock cords, and the 2 U-Bolts required to transfer parachute deployment forces to the avionics coupler and flight computer coupler. In addition, this will minimize the rocket weight. Figure 17 below shows the location of Parachute Bay 2. NASA SL FRR

24 Figure 17: Parachute Bay 2 Parachute Bay 2 will be made out of Blue Tube as will all other airframe sections. Compared to fiberglass or phenolic, Blue Tube is stronger and more durable and will not crack as easily. The drawback of adding this extra weight is minimal when compared to the strength and rigidity gained. Finally, Blue tube is commonly used in high-powered rocketry, which allows us to draw on the experience of others when constructing our rocket. 7.6 Parachute Bay 1 In parachute bay one, the main parachute will be packed together between the payload coupler and the avionics bay. This bay will be 27 long and 6 wide. This will give enough room for the 96 main parachute as well as the Nomex pad, two U-Bolts, ~ 8.9 yards of shock cords, and two quick-links. All of the recovery equipment can be fitted into this space without adding extra weight. Figure 18 shows the location of Parachute Bay One. Figure 18: Parachute Bay 1 The parachutes will be deployed by black powder ejection charges which will be wired to the altimeters. Redundancy in both the altimeters and the charges will ensure that the parachutes are deployed during descent. The parachutes and shroud lines will be protected from the ejection charges by Nomex pads of sufficient size to shield the respective parachutes. NASA SL FRR

25 Figure 19: Recovery System Components 7.7 Flight Computer Bay Figure 20: Flight Computer Bay Figure 20 above shows the SolidWorks model for the flight computer bay. The flight computer bay will be manufactured in one of the house labs and will not only contain the different electrical components of the flight computer, but also a servo and the pulleys with mounts that will allow control of the airbrake system. All structural components of the bay will be made of birch plywood excluding the threaded rods. The two threaded rods will be made of stainless steel. The density of the birch plywood will be lb/in^3 and the density of the stainless steel is lb/in^3. NASA SL FRR

26 The individual components consist of the following items: two end cap plates to contain the bay within the coupler, three inner rings for structural stability, one rectangular sled to hold the electronics, and two threaded rods to secure the bay inside the coupler. The two end cap plates have a diameter of in and thickness of 0.25 in. The three inner rings have a diameter of in and a thickness of 0.25 in. The rectangular sled has a width of 5.25 in, height of 8.75 in, and thickness of 0.25 in. The sled has a small cut out measuring 1.5 in wide and 0.5 in tall to allow for the accelerometer to be centered on the upper inner ring to allow for increased precision in measurements. The inner rings should experience less vibrations than the sled due to them having a more direct connection to the airframe. The two threaded rods are 13.5 in long and have a diameter of in. Figure 21 below shows the flight computer sled from another angle. Figure 21: Flight Computer Sled Birch plywood was chosen for the caps and rings because it is a standard material for such an application and is very reliable and strong. Stainless steel was chosen for the threaded rods for the same reasons. For the sled, both G10 fiberglass and birch plywood were considered. The plywood is lighter and cheaper than fiberglass, but fiberglass is stronger. Despite plywood being lighter than the fiberglass, the final weights of both would be relatively similar since less fiberglass may be used to achieve the same result. Birch plywood was eventually selected because it is easier to work with and less expensive than fiberglass. Lastly, it is possible for the fiberglass to carry a static charge, which could interfere with the electronic systems. The microcontroller at the heart of the flight computer is the Arduino Due. This model was selected as our microcontroller due to its compatibility with a large number of electronics equipment, and has NASA SL FRR

27 support for Arduino shields. The Arduino is a popular choice among amateur In the event that there are errors in servo functionality the assembly code will aid in debugging the software. The assembly code will also allow us to have develop a keen understanding of exactly how our software works at a low level throughout the development process. The figure below is a block diagram of the flight computer. The MPU-6050 (IMU), and the BMP-180 are used for data collection, the SD card is used for storing data to process post flight analysis Altimeters Figure 22 - Flight Computer Block Diagram The altimeter bay consists of two 9V batteries which are connected to and AIM USB and a Perfectflite Strattologger. Two separate brands were used to be sure that we were not getting a brand specific error. A backup altimeter is used incase the first altimeter fails and is also used to obtain an average data between both altimeters to get more accurate results. The altimeter and the 9V batteries are connected through A20 copper cables and a lock and key mechanism to ensure that no unintended voltage may escape the battery and switch on the altimeter before preparation for the flight such that any discrepancies in the altimeter results may be avoided. When the key is in the free position (able to be removed from rocket), a complete series circuit will allow the current to flow to the altimeters, which would then beep to ensure that it is functioning properly. The setup of the altimeter bay can be seen as shown in Figure 23. NASA SL FRR

28 Figure 23: Altimeter setup Apogee Control System Apogee Prediction A high level overview of our Apogee prediction algorithm is depicted below in our Software Flow diagram. After launch the software runs in a loop checking to see if the rocket is still in powered flight. After motor burnout begins collecting and saving Acceleration data from the MPU-6050 and barometric pressure data from the BMP-180. Using Newtonian physics formulas this data is used to calculate a projected apogee. This projected apogee is a couple hundred feet higher than the target 5280 ft. Then the software determines brake actuation time to step down the altitude once. In order to reduce the effects of wind on our rocket and to account for perturbations on the path of our rocket the software waits until a certain altitude is reached before actuating the air brakes again. This forces the airbrakes to separate its cycle into pieces in order to more accurately respond to changes in the system. The feedback control loop repeats three times for each actuation cycle. NASA SL FRR

29 Figure 24: Software Flow Diagram Using the position, velocity, and acceleration of the rocket we estimate a predicted apogee. These data points, in combination with the data, which the barometer delivers, provides a check for our on-board computer. The accelerometer data given by the IMU and filtered by our Kalman filter provides a reliable method of determining the relative acceleration of the rocket. A description of the Kalman Filter is shown in a later section. From these points we integrated the data to produce the velocity and altitude for the rocket. As we integrate the velocity and position, the error associated with both will compound, throwing off our current readings. Another problem is the lag experienced by computing previous values as current ones. This causes the computer to react to events later than what would in an ideal situation. The software determines braking time using the C D from CFD analysis of the final rocket design with the brakes open and brakes closed. We use the pressure from the barometer to measure the density of the air. From this point, we calculate the total force of drag on the rocket as it is coasting. With the acceleration of Earth, and the drag force known, we use the values of acceleration from our IMU to calculate the net forces on our rocket. From here, we predict the altitude of apogee without braking. With this newly acquired data, we deploy for a certain duration in order to reduce the velocity of the rocket and achieve the specified target apogee. We repeat this process several times to ensure the NASA SL FRR

30 greatest possible accuracy. Each time we predict the apogee, we lower the target altitude of our apogee, by increasingly smaller magnitudes, until the final iteration with the target of 5280 feet. The software repeats in a feedback control loop until the predicted apogee converges on the target apogee of 5280 feet. This control mechanism uses information from our sensing elements to actuate our controlling device to adjust the Cd variable to achieve the desired result of having an apogee of exactly one mile. The data is then integrated using the formula below. We use the trapezoid method of integration to compute velocity and displacement. The time interval of each trapezoid is of a second. Then we calculate the sum of the first two acceleration values. f(a) is the y-value of the acceleration at point a, and f(b) is the y-value of the acceleration at point b. While the formula uses a time interval of 1/50th of a second, we have not yet selected a suitable time interval to be used in our implementation of the formula. The same process of trapezoidal integration is repeated to achieve the rocket s position. Due to the nature of integration, information will necessarily be lost during this process. The use of the Kalman filter acts to counter some of this inaccuracy, but the fact remains that the velocity and displacement data will become less precise as the errors will compound as time progresses Kalman Filter Performance Predictions The accelerometer data given by the IMU is filtered by our Kalman filter and provides a reliable method of determining the relative acceleration of the rocket. From these points we integrate the data to produce the velocity and altitude for the rocket. Unfortunately, this poses the problem of how accurately the model will track the true motion of the rocket. As we integrate the velocity and position, the error associated with both compound, throwing off our current readings. This causes the computer to react to events later than what would in an ideal situation. We will be doing further research into correcting these errors in computing and reducing lag. We made the decision to use the Kalman filter due to the accuracy provided by the Kalman filter s statistical analysis of the data. By comparison, the complementary filter uses a simple analysis without any computation based on the statistical description for the interrupting noise. The graphs below show a simulated flight of a rocket with random measurement errors of the accelerometer. NASA SL FRR

31 Figure 25: Accelerometer - Time vs. Acceleration The graph above shows simulated measured and filtered acceleration data. The actual acceleration that was simulated is in blue labeled as the ideal acceleration. NASA SL FRR

32 Figure 26: Accelerometer - Time vs. Velocity The graph above shows the filtered velocity data that was integrated using the acceleration data. The gap between the filtered and the ideal velocity is the lag caused by filtering the data. Unfortunately, the airbrakes did not deploy during our test launch. No data was recovered Verification of Apogee Control System Requirement Solution Verification The vehicle shall deliver the payload to an apogee of exactly 5,280 ft AGL Simulations from OpenRocket determine apogee of the rocket if the ACS never actuates. Then using data from onboard sensors the flight computer will determine required actuation to step down the apogee to the Data collected from the onboard instruments and the Visual from onboard camera determine the fidelity of air brake actuation. NASA SL FRR

33 The launch vehicle shall be capable of remaining in a launch-ready configuration at the launch pad for a minimum of 1 hour without losing the functionality of any critical on-board component. target. High capacity batteries power low draw components to ensure the battery life extends that of 1 hour. The batteries have been tested with similar drain rates to that of our components. After more than one hour of testing we checked to see if the battery is capable of actuating the servo, the highest drain component. Air Brakes do not alter rocket s CP out of line with rocket s longitudinal axis The airbrakes are attached to a single servo so if there is a failure none of the airbrakes will be deployed. Avoiding the change in CP and drifting of the rocket. The airbrakes deploy symmetrically. Testing consisted of applying uneven loads to the airbrakes to determine if they still function under these circumstances. 7.8 Roll Induction System The on-board experiment that our team has chosen to design and construct a roll control system (RCS) that will mechanically induce roll during ascent, and then, after two complete revolutions, halt the roll for the remainder of the flight. To do this, we will be using two smaller, secondary fins controlled by servos in order to induce and control roll during ascent. A single flight computer will be located in the avionics bay in order to read and process incoming atmospheric and gyroscopic data, as well as actuate the servos controlling the RCS fins. This flight computer will be simultaneously controlling our airbrakes, however, safeguards will be put in place to make sure that the RCS fins will not be deflected while the airbrakes are deployed. NASA SL FRR

34 Figure 27: Roll Induction System Logic The above figure walks through the RCS logic of the flight computer. A barometric pressure sensor and a 6-DOF Inertial Measurement Unit (IMU) will be located inside the flight computer bay with the flight computer and servos controlling the RCS fins. These fins will start in the neutral position during launch (0o deflection). Once the accelerometer detects ignition, (by the spike in acceleration) the clock will start in the flight computer. After motor burnout, which will be determined from the manufacturer s data, the RCS fins will actuate to a set deflection angle - inducing a roll in the rocket. The gyroscope built into the IMU will read and process the rotational velocity and will integrate the data to determine the number of rotations the rocket has turned through. After two full rotations, the servos will actuate the fins in reverse and stop the roll that had been induced by the actuation of the fins. Once the flight computer registers that the rocket has completed two full rotations, it will begin its active roll control phase. During this time, the flight computer will be taking angular velocity measurements from the IMU and comparing the value against a setpoint of 0 rad/s. Then, using a feedback control loop (PID controller), the computer will iteratively calculate the error against this point and adjust the servo position, and in turn, the fin deflection angle. A detailed breakdown of all of the electronics used in this system can be found in section (Flight Computer Bay) above. NASA SL FRR

35 Figure 28: PID control loop diagram The fins used for roll induction will be much smaller than our main fins. They will have a 3 root chord and a 1 tip chord with a 2 span (see Figure 29 below). These fins will be fabricated out of carbon fiber and they will be made in-house in ISU s composites lab. Figure 29: RCS fin dimensions Knowing where the forces will act on the fin is crucial to the system design. This can be fairly challenging to compute, given that the forces will be distributed across the fin during flight. However, in order to simplify this calculation, it was assumed that the aerodynamic forces would be acting on the mean aerodynamic chord (MAC) and would be located at the quarter chord. Because of this fin s simple geometry, the MAC is located halfway between the root and tip chords. Once this location is known, it is then possible to determine how far apart the aerodynamic forces are, and also the couple moment they create. This is important because using this value and the rotational moment of inertia of the rocket, (obtained from RockSim) we can calculate the angular acceleration induced by each degree of fin deflection. This will be used to accurately adjust our roll angle without overshooting or undershooting. NASA SL FRR

36 The actual mechanism that controls the fins is shown below (Figure 30). Both fins will be controlled by a single servo instead of two so that our rocket will remain stable in the event of a servo failure. If the servo were to fail for whatever reason, the aerodynamic forces would likely cause the fins to reset back to their original orientation. And if, for whatever reason, the fins stayed locked in their positions when the servo failed, the rocket would experience an increased roll rate which would slow the rocket slightly and decrease its desired apogee. Figure 30: Roll induction system (no deflection) Figure 31: Assembled Roll Control Configuration The system will be driven by a HS-7950TH continuous servo which generates 403 oz-in of torque. The servo will be attached to two 1 diameter cylinders through 2 sets of ball and socket joints. These cylinders will either be fabricated from aluminum rods or 3D printed out of ABS plastic. Aluminum would be preferred for its structural integrity and low cost, however it will be more difficult to machine. Both materials will likely be prototyped to see which one is more effective and reliable. NASA SL FRR

37 Figure 32: Roll induction system (with fin deflection) These cylinders will be supported by two independent ¼ diameter aluminum dowel rods. When the servo turns, one of the cylinders will rotate clockwise and the other will rotate counterclockwise. The shorter of the two cylinders (shown on left in Figure 32 above) will be 1 long and have a very simple geometry with a single through hole through its center and a small counterbore hole on its exterior where the ball joint will be fixed. This shorter cylinder will be fixed to the aluminum rod so that the aluminum rod will rotate when the servo rotates the cylinder. The longer cylinder is slightly more complex, however (see Figure 33). It will feature the same counterbore hole for the ball joint, but it will be 2 long and - instead of a through hole for the aluminum rod - the rod will be fixed halfway into the cylinder. There will also be a ¾ deep, ⅝ diameter hole bored into the other end. This hole will house 2-3 ball bearings which will allow the other aluminum rod to rotate with minimal friction. NASA SL FRR

38 Figure 33: Dimensioned drawing of long cylinder for RCS In between each cylinder and its respective fin, there is a L-shaped platform that will hold up the aluminum rods and keep them level and stationary. These will be screwed into the bulkhead they rest on to ensure that they do not move during flight. There will be a ⅝ hole in the vertical half of each mount where another pair of ball bearings will be placed. As shown in the figure below, the aluminum rods will run through these bearings (shown in black) and through into/through the cylinders Figure 34: Exploded view of RCS main shaft 7.9 Air Brakes NASA SL FRR

39 Figure 35: Air Brakes Assembly Figure 36: Air Brake Design Figures 35 and 36 above show the general layout of our air braking system. As can be seen in the figures, the air brakes are operated by cables that run the length of the motor mount tube and are controlled by the actuation of the servo which in turn is managed by the on board flight computer. The high torque servo is located on the aft flight computer bulk plate. A command from the flight computer will operate the servo which will in turn rotate a pulley winch attached to the servo shaft. The cables attached to the NASA SL FRR

40 winch will wrap around the winch as it rotates. The cables will be routed around the pulleys located radially out from the servo winch and mounted on the same bulk plate. This will convert the cables horizontal displacement to vertical displacement along the motor mount centering rings. The figure below shows the pulley-winch configuration. Figure 36: Servo and Pulley Assembly The cables, now orientated vertically along the motor mount, will travel through tubes drilled through the motor mount centering rings for integration purposes which will be explained later. Once the cables reach the aft fin centering ring, they will travel through an aluminum tube which will be structurally attached to an interior mounting bracket to be explained later. The aluminum tube will transfer the cable tension around a radius of 3 inches where it will travel to the exterior of the rocket. The cable, once exterior to the rocket, will travel along the leading edge of exterior fins which will be explained in detail later. The fins will house an exterior pulley which will curve the cable back to a vertical position where it will be attached close to the base of the air brakes. Figure 37 below shows the exterior fins containing the pulley. NASA SL FRR

41 Figure 37: Exterior Fin Pulley Assembly The air brakes are hinged to an interior hinge bracket mounted to the underside of the aft fin centering ring. A spring attachment will keep the air brakes closed during off duty cycle and will also prevent the cables from losing contact with the pulleys during air brake retraction Servo The selection of the servo was heavily dependent on the torque that we need to produce. We wanted to ensure a single servo could extend all the air brakes with a relatively fast actuation time in order to reduce the computational loads on the flight computer. The servo we ultimately selected was the HS7950TH Servo that produces a maximum torque of 486 oz/in at an operating voltage of 7.2 volts. A 7.2 Volt 2100 mah LiPo Battery Pack will power both the high torque servo as well as the arduino microcontroller. A NiCd battery was chosen as it is capable of handling the high current needs of the high torque servo during air brake actuation. In addition, NiCd batteries are easily rechargeable and are lighter than its alkaline counterpart. NASA SL FRR

42 Figure 38: Flight Computer Servo Assembly Figure 39: Servo and pulley assembly A half inch radius spool will be mounted to the shaft of the high torque servo. The spool s function will be to reel up the cables as the servo actuates, thus putting tension on the cables which will in turn actuate the air brakes. The spool will be constructed from two servo hub horns with 8 steel bolts running down to form a spool. Each of the cables (one for each air brake) will be mounted to their respective bolts on the spool. Mounting the cables will be simple and there will be very little force variance during rotation. Figure 39 above demonstrates this configuration. The maximum torque of the servo will produce 272 N at half an inch. Between the cables there will be 68 N for each individual cable. The pulleys will be mounted inside the flight computer bay which will transfer the force parallel to the rocket body. Each pulley will be mounted by a fabricated aluminum bracket that will run atop of the pulley to ensure the Bowden cable stays in its track. The fabrication will consist of using a water jet to slice a piece of aluminum and then use a bending machine to make the NASA SL FRR

43 correct bends. The pulleys are made of brass and contain bearings so there will be relatively zero force loss through the transfer. The 4 pulleys will be mounted to brackets which will be attached to the motor centering ring. The bracket will be manufactured with aluminum to withstand the high temperatures associated with being in close proximity to the motor and not deform. Aluminum was also chosen due to the material s light weight. The bracket will be fabricated using three 1 mm thick aluminum sheets. The bottom plate, which will be attached to the centering ring, will have one edge up against the motor mount with a 90-degree arc. The outside edge will be up against the body wall with a 45-degree arc. The plate to which the pulley will be attached to, through the wall of the rocket, will be attached perpendicular to the outside edge of the bottom plate and extend up 30 mm. This plate will match the curve of the rocket wall. The top plate will sit perpendicularly on top of the second plate and will provide a structure for the cable controlling the airbrakes to pass through. The cables will then run down alongside the motor mount into a 3 inch arched tube traveling through 35 degrees. This will direct the cables to the outside of the rocket where there will be the same model pulley, as on the flight computer bay. Here the cable wraps over the pulley and attached down to the airbrake 1 inch from the hinge. The cables being used are Bowden cables and the expected force loss is 15% through the system Hinges Each air brake hinge will be manufactured with aluminum. Aluminum was chosen primarily due to its ability to withstand the high temperatures and its low density. Plus, they would be strong to withstand any shear loads due to the air brake. Aluminum will be less susceptible than other materials such as 3D printed materials to weakening and deformation when exposed to high temperatures. Figure 40: Air Brake Assembly NASA SL FRR

44 The hinges will be fabricated by using thick aluminum sheets. Two sheets both approximately 1 in width and 2.5 in length will be needed for each hinge. One piece will have a square section of approximately 1/3 x0.5 removed and the other will have two sections of the same dimensions removed. The side where the material was removed on each sheet will then be bent in a circle with an inner radius of approximately These two pieces can then be held together by a pin that will allow for rotation. Holes will be cut in a line on each plate of each hinge with a diameter of to allow screws to hold hinges in place and to attach airbrake to the hinge. The holes will be placed with centers at no less than the times the diameter apart and a distance from center to the edge of no less than 2 two times the diameter. Figure 41: Interior view of the airbrake connection The screw holes are created in line with one another to allow for easier attachment of the hinge to the airbrake. Finally, one end of the hinge will be rounded to match the curvature of the inside of the airbrake. One end of each hinge will be mounted onto a bulk plate on the inside of the rocket and the other end will be attached to airbrake. A possible alternative to this method of fabrication is to mold each hinge. Molding each hinge will allow the material to be attached at both ends of the bend, and will address concerns with respect to unwanted bending in the hinge during use of the air brakes. In order to test the aforementioned design, ANSYS software will be employed to analyze component performance under loading. Any yielding or deformation will be noted and addressed in a design revision Pulley The exterior fin pulley will transfer the cable tension to the airbrakes. The pulley will mount to the interior mounting bracket through the airframe, minimizing the air brake stress applied to the airframe. NASA SL FRR

45 Figure 42: External View of Airbrake Assembly The specific pulley that will be used will be a 1 diameter pulley with a.1250 bore diameter from zore.com. The pulley wheel is fabricated from brass and includes a steel bearing which will be connected to a steel shaft. The bearings will eliminate any torsional friction resulting in a more efficient air brake system. The pulley is shown in Figure 43. The mounting bracket will be laser cut from.125 aluminum sheet metal and bent into the required 3D shape using metal bending tools in an accessible mechanical engineering lab. Once the design is further developed, a stress analysis will be performed using FEA to ensure the design will adequately carry the air brake forces. Figure 43: Airbrake Pulley 7.13 Exterior Fin NASA SL FRR

46 The new design for the small exterior fins is a right trapezoid shape. The fins will be able to be screwed together to allow easy disassembly into the two halves. The pulley will be housed within the exterior fin and the steel cable will exit out through the bottom of the exterior fin to connect to the airbrakes. This design choice allowed us to route the cables outside of the body of the rocket, simplifying the connection from the servo to the flaps themselves. Routing a cable over a pulley to the exterior surface of the flaps minimizes mechanical losses and maximizes the force we can apply to the brakes with the servo. Both halves of each fin will be 3D printed to ensure an exact fit, and will be further prepped for flight. Rough surfaces will be sanded down and the fins will be painted over, thus having the minimal effect on the drag coefficient of the vehicle. The design for the interior of the smaller fins was more of a challenge. The first proposition was to add cross-etching structures in order to increase the structural support. The decision was that this might not be the most efficient way to manufacture the fin. The final design for the interior was to add solid sections located around the edges that will take the most stress. The final design for the exterior fin is shown below in Figure 44. Figure 44: Exterior Fin/Pulley Assembly This design is more efficient to manufacture and the mass of the fin will be less. The fin will be manufactured using a three-dimensional printer and the material of choice will be ABS plastic. The loads that will be placed on the fins are of no concern because the fins will be plenty strong to withstand pressure loads and any loads from the air brakes themselves. The surface area of each fin is small enough where there will be no substantial force to try to collapse them Motor/Motor Mount The motor mount will house the chosen motor and transfer the thrust from the motor to the rocket airframe. Figure 46 shows the overall assembly of the motor mount tube. The motor housing tube will consist of standard 75 mm Blue Tube airframe. An aeropack flanged retainer (shown in figure 45) will screw into the rear centering ring and will secure the motor casing inside the motor tube. This retainer cap is easily screwed into place requiring no tools and reducing assembly time. This aft retainer ensures NASA SL FRR

47 motor retention during. A motor change will require the simple removal of the rear closure cap from the retainer base. Figure 45: Motor Mount Retainer The motor tube will be centered along the rocket airframe using a series of the centering ring configurations. The two middle centering rings will be constructed out of ½ birch plywood and will serve the dual purpose of increasing the fin adhesion surface as well as centering the motor tube. The bulkhead at the forward end of the motor mount will be constructed out of ½ birch plywood and will be attached to a ¼ centering ring for motor mount stability. The aft centering rings will consist of one ¼ thick birch plywood centering ring with OD of the airframe OD as well as a similar centering ring with OD of the airframe ID. This will allow some of the motor thrust to transfer directly into compression of the airframe itself. Figure 46: motor mount with centering rings Since the motor mount will also be directly transferring loads from the airbrakes, the team plans to perform numerous analysis and tests on the motor mount to ensure structural stability. Finite Element Analysis (FEA) analysis is planned to be performed on all critical parts of this motor mount as well as verified with actual in lab results using an Instron machine located in the M2I labs. NASA SL FRR

48 Our selected motor, a Cesaroni 5015-L1115-CS-P, with our current mass budget brings us to an apogee of 5,554 ft. To calculate this apogee, we needed a realistic approximation of both the weight of our rocket and the amount of space needed for each subsystem. Using anticipated hardware required for each subsystem, we created a spreadsheet of weights and dimensions for each component of the rocket and determined a length and mass estimate from the spreadsheet. This apogee was calculated using Openrocket, an open sourced model rocketry simulator. While the selected motor gives us an apogee over 5,280 ft, we feel that the excess will both give us room in case our initial mass estimate was low as well as provide an excess of apogee for the airbrakes to control Fiberglass Fins Figure 47: The fiberglass fins The fins for the rocket were be cut to dimensions from.125 thick G-10 fiberglass sheets. G-10 fiberglass is a strong and durable composite material used very often in high powered rocketry and widely recommended. Since no additional layers of support are required, the fabrication of the fins will be at minimal which will make it easier for fabrication of the rocket to continue on schedule. The fins will be cut in-house, as the specified dimensions are very specific and aren t available for fabrication by a rocketry outlet. The split fin design was created not only for aesthetics, but also for strength during the in-flight performance. Fin flutter analysis was done for each fin using NACA s Fin Flutter Boundary equation for a launch in AL. The forward fin calculated a fin flutter speed at 1,022.2 ft/s and the aft fin calculated a speed at 1,182.9 ft/s. These speeds well exceed the expected 700 ft/s velocity and provide a sufficient margin of safety. To further confirm this flutter velocity, we will be using a software called FinSim to calculate the fin flutter using more advanced and accurate results. Using a vice to perfectly align the fins, epoxy will be applied to ensure max structural rigidity Mass Statement The calculated mass total is a cumulative sum of tabulated data specific to every part added to the rocket as the model was developed. As much information as possible was gathered from vendors, manufacturers, and calculations; this kept estimates to a minimum. The total approximate weight of the launch vehicle is 40.1 lb., while individual subsystem weights can be seen in the breakdown table in Table 1 below. The current projected apogee needs to remain about 300 ft above the goal apogee of NASA SL FRR

49 5,280 ft. to allow a window for the airbrakes to control. In order to meet this goal, the maximum margin of change in total mass that the rocket can undergo is around 1.38 lb. located at the CG. This small reserve is roughly 3.91% of the weight of the rocket. Despite the low number, there are several options by which weight can be cut. Regardless, the mass budget is thorough enough to give a feasible prediction of the final weight and small adjustments should not provide much of an issue. To be more specific on calculations and measurements done, the masses of the airframe and motor were calculated using OpenRocket, while all other masses were calculated via Solidworks using built in densities and overridden masses for known parameters. Some approximations still had to be made, including that of paint, specific electronic components, and other body materials. Table 1: Weight Breakdown Subsystem Motor Mount/Flight Computer Bay Assembly (excluding motor) Motor (3757L800 - PreFlight) Parachute Bay 1 (including parachutes) Avionics Bay Parachute Bay 2 (including parachutes) Specimen Bay (including specimen weight) Nose Cone Paint (Estimate) Total (Full Scale Launch) Total (Including Paint) Weight (lb) ~ The mass is estimated using an excel parts list that also doubles as a mass and cost calculator. The parts list is maintained by a team member and updated whenever changes are made to components within respective teams. The masses/prices of each part are estimated based on supplier specifications and assumed to be correct. The estimated mass of the rocket will increase as more information is received from each team regarding parts. Obviously as the parts list grows, the mass will increase, but the rocket mass should not exceed the critical point at which a critical design change would have to be made. 8) Recovery Systems 8.1 Recovery Overview NASA SL FRR

50 Figure 48: Recovery Configuration The recovery scheme will consist of two events. The first occurring at apogee and the other at 800 ft. AGL. At apogee, the first event will occur aft of the avionics bay, ejecting an 18 elliptical drogue chute connected to both sections of the rocket. This first ejection charge will be 3.7 grams of black powder needed to separate both sections successfully. This attachment consists of the shock cords each connected to a quick-link hooked on to a U-Bolt. This U-Bolt is secured to the bulkhead of the rocket section. This event is shown above in configuration 1 to descend at a steady velocity of ~97 feet per second. As to minimize the potential interference between the two sections of the rocket, the shock cords connecting each section of the rocket will be different. The top section of the rocket will have a cord length of 200 and the bottom section a shock cord length of 110 of which both are connected to the drogue chute. At 800 ft. during descent, altimeters will fire the second event s black powder charges in the main parachute bay, jettisoning the rocket s specimen bay section away from the lower half of the rocket. The ejection charges will be 4.1 grams of black powder to pull out the main parachute. During separation, a 96 main elliptical parachute attached to the altimeter bay bulkhead will be ejected from the main parachute bay. The main parachute will have a shock cord length of 150 and the camera bay a length of 120. These shock cords attach to a quick-link which is connected to the U-Bolt on each section. NASA SL FRR

51 Figure 49: Main, pilot, and drogue parachutes Currently, the main and drogue parachute will fall at 13.9 feet per second. The 15 feet per second descent rate for the individual sections is a standard in high powered rocketry and will be sufficient to induce impact energies below 75 ft-lbf in each section. The tables below under the subtitle Descent, show the calculated impact energies to be well below 75 ft-lbf. The table below also gives the current masses and descent rates for the recovery configurations as shown above. Table 2: Recovery Descent Rates Configuration 2 (Main/Drogue Combination) Section Mass: 27.9 lb Parachute Sizes: 120 (Elliptical Main) and 18 (Elliptical Drogue) Descent Rate: 13.7 ft/s NASA SL FRR

52 Table 3: Recovery Descent Rates Configuration 1 (Entire assembly on drogue) Section Mass: 31.9 lb Parachute Size: 18 (Elliptical) Descent Rate: ft/s 8.2 Kinetic Energy Management The kinetic energy management of our rocket in final stage of ascent will be control by the use of our carbon fiber air brakes, which will aid in reducing the rocket s kinetic energy to zero at the specified 5,280 ft apogee. From the point directly after apogee until the second parachute event at 800 ft, during descent, the kinetic energy of our rocket will be decreased and then maintained through the use of an 18 elliptical drogue chute. Kinetic energy of the airframe sections in the rocket will be reduced substantially after the second parachute event until landing through the use of a 24 pilot elliptical parachute and a 120 main elliptical parachute. The kinetic energy of these sections will be well within the safety value limits, due to the size of the parachutes selected and the parachute packing techniques that will accompany them. 8.3 Instron test For each recovery event, the force on the rocket is pretty high so making sure the attachments are strong enough is critical. The quick-links and U-Bolts have a pound strength test of 1,500 lb. and the shock cords a test of 1,800 lbs. Although this recovery hardware is more than robust enough to handle the recovery forces, a tensile test was performed to confirm this. Using our Instron machine, we tested the structural tensile strength of our avionics bay to simulate the shock load force of our recovery system. We calculated 514 lbs of force as our shock load on our main parachute so we used this value as our maximum force. We tested up to 560 lbs with no breaks or fractures at all. This test confirmed the possibility of a shock load of 560 lbs that the avionics bay section NASA SL FRR

53 could experience. The U-Bolt on the avionics bay is rated to 1,500 lbs so this rating will give us a safety rating of just under 3. Figure 50: Avionics Bay Tensile Test in Instron Machine Figure 51: Avionics Bay Tensile Test Results NASA SL FRR

54 In past years, our rocket was required to separate into two sections jettisoning the rocket s specimen bay away from the lower half of the rocket. Because this requirement is no longer in place we are keeping our recovery system in one piece to eliminate some potential risks. Last year at competition the two independent sections were too light and dragged after landing which made recovering them more difficult and led to a higher chance of damaging components. Therefore our new recovery system should provide for an easier and safer recovery. 8.4 Black Powder Calculations The mass (in grams) of black powder required or adequate separation due to the forward changes was calculated using the gas properties of FFFFg (or 4F) black powder and the ideal gas equation: PV = NRT; P is the pressure of the compartment in psi; V is the volume of the compartment being pressurized in cubic inches; N is the mass of black powder required in lb.; R = 266 in-lbf/lbm for 4F black powder; T = 3307 R (combustion temperature). When solving for the mass of black powder required with the values above, the ideal gas equation reduces to: N = 0.006*D^2*L (grams); D is the diameter of the pressurized compartment in inches; L is the length of the pressurized compartment in inches. Force is defined as the product between pressure and area, a relationship that can be substituted into the reduced ideal gas equation for 4F black powder, resulting in the following equation: N = *FL (grams); F is the force in lbf; L is the length of the pressurized compartment in inches. A good rule-of-thumb is to design for 15 psi within the pressurized compartment and it was assumed that the entire mass of the ejection charge was being burned and converted into a gas. However, the amount of force produced by a charge needs to be a reasonable limit of around lbf. 15 psi with a 6-inch diameter rocket will result in a force of 424 lbf, which is outside the recommended limit. 12 psi will result in a force of around 339 lbf and requires a black powder mass of around 3.28 grams. Adding an extra 25% to account for friction and pressure loss results in 4.1 grams of 4F black powder required for each of the redundant forward charges. 8.5 Recovery Electronics We selected an AIM USB and Perfectflite Stratologger altimeter for use as recovery electronics because it is known to give accurate readings during operation, and provide accurate deployment of our ejection charges. The altimeters are lightweight and small. Just 2.76 long, and wide for the AIM USB altimeter. The altimeters are well suited to calculate the altitude and apogee of the rocket. When the rocket descends to 800 feet, the altimeters ignite the igniters, which ignite the black powder charges. To ensure that the charges are blown, there is a redundant charge ignited. This ensures the separation of the rocket in the case that the initial charge is compromised. The result force splits the rocket and the parachute deploys. In order to keep the altimeters on we use rotary switches, which are resistant to the NASA SL FRR

55 vibrational forces which are exerted on the switches, to arm our altimeters. Each altimeter has its own switch and power supply to decrease the chances of a failure. Figure 52: Avionics Electronics Schematic In order to meet the safety requirements, we followed a preflight checklist to ensure that all devices operated correctly. The battery voltage was checked via a voltmeter and ammeter and the connections between devices inspected to avoid short circuit and power failures. The rotary switches were locked into place prior to launch by the pre-flight inspector. These switches are accessible from the exterior of the rocket s fuselage and cannot move from their locked position. The black powder charges were not be able to fire unless the altimeters are armed Redundancy Features There are many redundant features included in the avionics bay. The bay contains two completely independent different branded altimeters each connected to their own separate drogue and main charges and each powered by their own 9 volt battery. Each altimeter is connected to a separate rotary switch. In the event of a switch failure, ejection charge failure (black powder wet, igniter failure), power failure or altimeter failure, the avionics bay will still successfully deploy the parachutes. The recovery NASA SL FRR