Illinois Space Society University of Illinois Urbana Champaign Student Launch Maxi-MAV Preliminary Design Review November 5, 2014

Transcription

1 Illinois Space Society University of Illinois Urbana Champaign Student Launch Maxi-MAV Preliminary Design Review November 5, 2014 Illinois Space Society 104 S. Wright Street Room 321D Urbana, Illinois 61801

2 Contents Table of Figures...5 Glossary of Terms...7 General Information...8 Managers... 8 Major Sub Team 1: Structures and Recovery... 8 Major Sub Team 2: AGSE... 8 Minor Sub Teams:... 9 NAR Section... 9 I) Summary of PDR report Launch Vehicle Summary AGSE Summary II) Changes made since Proposal Changes to Project Plan Changes to Vehicle Criteria Changes to Sample Container System Changes to AGSE System III) Vehicle Criteria Selection, Design and Verification of Launch Vehicle Mission Statement Requirements Mission Success Criteria System Design Review Subsystem Descriptions Performance Characteristics Vehicle & Recovery Requirements Risk Management Necessary Components and Risk/Delay Impact Planning of Manufacturing, Verification, Integration and Operations Confidence and Maturity of Design Dimensional Drawings Mass Statement Recovery Subsystem Page 2 of 139

3 Recovery System Analysis Robustness of Recovery Components Mission Performance Predictions Mission Performance Criteria Flight Profile Simulations Drift Interfaces and Integration Payload Integration Internal Interfaces Interfaces between Vehicle and Ground Interfaces between Vehicle and Launch System Safety Preliminary Checklists Comprehensive Checklist Safety Officer Preliminary Hazard Analysis Failure Mode Analysis Environmental Concerns IV) AGSE/Payload Criteria Selection, Design, and Verification of AGSE System Review Subsystem Overview Performance of the subsystems AGSE Verification Plan and Component Requirements Preliminary integration plan Precision of instruments and repeatability Electrical Schematics Key Components of the AGSE System AGSE/Payload Concept Features and Definition Creativity, significance, and level of challenge Science Value AGSE Objectives AGSE Success Criteria Page 3 of 139

4 AGSE Approach AGSE Testing AGSE Progress V) Project Plan Status of activities and schedule Budget Plan Funding Plan Timeline Educational Engagement Plan VI) Conclusion Appendix A Illinois Space Society Educational Feedback Form Appendix B: ISS Tech Team Safety Policy Appendix C: Flight Simulation Code Flight Simulator File flightsim.m Equations of Motion eomfun.m Sample Container Equations of Motion eomfunsamp.m Landing event for main vehicle crashevent.m Landing event for sample container crashevent2.m Appendix D: AGSE Launch Rail Torque Calculations Appendix E: Robotic Arm calculations Page 4 of 139

5 Table of Figures Figure 1 Model of Full System... 9 Figure 2 Booster System Figure 3 Cutaway View of Booster System Figure 4 Exterior View of Sample Container System Figure 5 Cutaway View of Sample Container Figure 6 Zoomed In View of Sample Container Cutaway Figure 7 Exterior View of Avionics Bay Figure 8 Cutaway view of Avionics Bay showing Altimeters Figure 9 Cutaway View of Avionics Bay showing Batteries Figure 10 Close up view of Motor Retainer Figure 11 View of Sample Container without Airframe Tubing Figure 12 Side view of Sample Container without Airframe Tube Figure 13 Exterior view of Sample Container showing the Hole in the Airframe Figure 14 Cutaway view of the Sample Container Showing Flex Sensors in Green Figure 15 Cutaway view of Sample Container showing Altimeters Figure 16 Cutaway view of Sample Container showing Gear System Figure 17 Cutaway view of Sample Container showing Batteries Figure 18 Sample Container Electronics Schematic Figure 19 Image of Rail Button Placement Figure 20 Zoomed in Image of Rail Button Figure 21 Full Vehicle Dimensions Figure 22 Avionics Bay Dimensions Figure 23 Sample Container Dimensions Figure 24 Fin Dimensions Figure 25 Main and Drogue Parachute Avionics Figure 26 Sample Parachute Avionics Figure 27 Skyangle Parachute Figure 28 Iris Ultra Parachute Figure 29 Bulkhead Assembly Figure 30 First Stage of Recovery Deployment Figure 31 Second Stage of Deployment Process Figure 32 Third Stage of Deployment Process Figure 33 OpenRocket Altitude Simulation Figure 34 OpenRocket Velocity Simulation Figure 35 Custom Altitude Simulation Figure 36 Custom Velocity Simulation Figure 37 Custom Velocity Simulation Results for Sample Container Figure 38 Altitude Simulation Comparison between OpenRocket and Custom Simulation Figure 39 Velocity Simulation Comparison between OpenRocket and Custom Simulation Figure 40 Aerotech L1150 Thrust Curve Figure 41 Vehicle Stability Margin Figure 42 Rocket on the Lowered Rail Page 5 of 139

6 Figure 43 Rocket on the Raised Rail Figure 44 Robotic Arm System Figure 45 Rail System Linear Actuator Figure 46 Motor Igniter Insertion System Figure 47 AGSE Electrical Schematic Figure 48 Relationship between required Actuator force and Rail Angle Figure 49 Gantt Chart of Project Plan Page 6 of 139

7 Glossary of Terms Booster: The lower section of the vehicle which includes the lower airframe tube, fin, rail button and motor assemblies. Serves as the storage compartment for the drogue parachute. Avionics Bay: All hardware composing and contained within the vehicle s center coupler. This includes bulkheads, avionics, rotary switches, the switch band and the coupler tube itself. Sample: The PVC pipe component placed and sealed within the vehicle. Simulates a Mars Soil Sample. Upper Airframe: The upper body tube of the vehicle. Contains the main parachute and sample canister, as well as the sample parachute. Sample Canister: The portion of the upper airframe dedicated to the processes of sealing and jettisoning the sample. This section is capped by bulkheads at either end and contains the hatch system, altimeters, and the housing for the sample itself. Hatch System: Components used to seal the sample within the vehicle. Includes motors, power supplies, flex sensors, gears, and the hatch door itself. Robotic Arm: The equipment used to pick up the sample and place it within the vehicle. Igniter System: The equipment used to insert the igniter into the vehicle. Rail System: The system which acts as the vehicle s launch platform and is capable of being autonomously erected to the launch position. AGSE: The combination of the Robotic Arm, Igniter System, and Rail System, as well as the methods of controlling these systems. Drogue Parachute: A small parachute stored in the booster airframe to slightly slow the entire vehicle during descent from apogee to 1000 feet during descent. Main Parachute: A large parachute stored in the upper airframe used slow the booster and avionics bay to a safe landing speed. Sample Parachute: A moderately sized parachute used to slow the upper airframe and nose cone to a safe landing speed. Page 7 of 139

8 General Information Safety Officer Derek Awtry Team Leader David Knourek, Project Manager Phone: (708) Managers Project Manager: David Safety Officer: Derek Structures and Recovery Manager: Jacqueline AGSE Manager: Ian Webmaster: Derek Educational Outreach Director: David The ISS Tech Team participating in this competition consists of about 30 students, essentially split evenly into two major sub teams. Major Sub Team 1: Structures and Recovery The first main sub team of about 15 students is the Structures and Recovery team. This team will be responsible for design and construction of the vehicle, as well as the recovery avionics and parachute systems. The Structures and Recovery team will also be responsible for the system of sealing and jettisoning the sample. The Structures and Recovery manager is Jacqueline. David, Derek, Mike and Kamil are key technical members for the Structures and Recovery team. Specifically, David is responsible for the design of the vehicle, and Derek is responsible for construction procedures. Mike is charged with management of the recovery systems and Kamil is in charge of the sample canister and hatch systems Major Sub Team 2: AGSE The second major sub team is the Autonomous Ground Support Equipment team. This team will be responsible for design and construction of a robotic system to contain the sample within the vehicle, as well as systems to erect the rocket from the horizontal position and install the motor igniter. Ian is the AGSE manager. Alex, Chris and Rick are key technical personnel for the AGSE systems. Alex is tasked with leading the design and construction of the robotic arm, and Chris will manage the motor igniter installation system. Rick is responsible for the system which raises the rocket from the horizontal loading position to the launch configuration. Page 8 of 139

9 All sub team managers are mainly charged with organizing their respective teams and overseeing design and work meetings, however they are also integral to the technical design of their systems. Although key technical members are listed for the major sub teams, technical work will be equally split between all team members. In this way, the team may draw on the experience of past members while building the knowledge of new members. Minor Sub Teams: Minor sub teams of 5 to 10 students will be responsible for web design, safety planning, and educational outreach. Each student on these sub teams is also a member of either the AGSE or Structures and Recovery sub teams. Derek will manage the web design and safety sub teams, and David will manage the educational outreach activities.. NAR Section The ISS Tech Team will be working with members of Central Illinois Aerospace (CIA) to facilitate launch test launches, mentor the team, and review system designs. Specifically, Mark Joseph will be the NAR mentor for the ISS Tech Team. CIA is section 527 of the National Association of Rocketry. The CIA organizes bi-weekly launches at several locations close to the university, depending on the time of year and launch field conditions. Figure 1 Model of Full System Page 9 of 139

10 I) Summary of PDR report Team Name: Illinois Space Society Student Launch Team Mailing Address: 104 S. Wright Street Room 321D Urbana, Illinois Team Leader: David Knourek, Project Manager Phone: (708) Safety Officer: Derek Awtry Mentor: Mark Joseph, NAR Number: 76446, Certified Level 2 Launch Vehicle Summary For the purpose of obtaining hands on engineering work experience, it was the team s decision to design, build and implement a rocket from custom selected materials and components. The vehicle will be a single stage rocket with one motor and a triple deploy recovery system. The vehicle for this project will measures inches from the aft end of the motor retainer to the tip of the nose cone. The vehicle body will be constructed of 5.5 inch Blue Tube, which has an inner diameter of 5.38 inches. The current design mass of the vehicle is lbs, however the motor and parachutes chosen allow for a flight mass of 45 lbs. The motor chosen for this vehicle is the Aerotech L1150 reloadable rocket motor. The recovery system for this vehicle is based off of the standard dual deploy system, however it will include a third parachute for the purposes of safely returning the sample to the ground. At apogee, a Skyangle 36 drogue parachute will deploy from the booster airframe. At 1100 feet above ground level, an Iris Ultra 60 sample parachute will eject from the upper airframe between the sample container and the nose cone. At 1000 feet above ground level, an Iris Ultra 96 compact main parachute will deploy from the upper airframe much as it would in a dual deploy vehicle. However, this main parachute will only be attached to the lower portion of the rocket, and not to the upper airframe or nose cone. This allows the ejection event to jettison the sample canister, which is contained within the upper airframe. The Milestone Review Flysheet is presented as a separate document on the team s website, and may be found at AGSE Summary The team will build a robotic arm to capture the payload, an actuator powered system to raise the rocket to a vertical position, and another actuator powered system to insert the igniter into the motor of the rocket. The systems will be powered by a 12V lead acid battery. A Raspberry Pi computer will be used for the system processing. The robotic arm components will be 3D printed. An actuator used to raise the launch rail and rocket will be connected to the main computer. The ignition system will utilize a z-plate and actuator to insert the motor igniter into the motor. Page 10 of 139

11 II) Changes made since Proposal Changes to Project Plan No major changes have been made to the team s project plan other than additional details in all aspects of the plan. The budget has been broken down into individual components, however this was solely to update the relevant financial information, and represents no major impact on the success of the project. Details were added to the project timeline to include planned completion dates for design and construction activities. Changes to Vehicle Criteria The motor to be used for this vehicle has been changed from an Aerotech K780 to an Aerotech L1150. The team expects the mass of the vehicle to increase throughout the design and construction process and increasing the motor size allows for a heavier vehicle to meet the altitude target. This is a relatively large increase in motor size, so the team expects to include ballast mass in the vehicle in an attempt to reach the target altitude as precisely as possible. The tether system previously used in the jettison process has been removed from the project design. The initially proposed vehicle included a pyrotechnically detachable tether connecting the jettisoned sample canister to the remainder of the rocket. This was due to a misinterpretation of the competition requirements, as the team believed RSO approval was needed before jettisoning the sample. Removing this system allows for a much simpler deployment system and will highly increase the chances of recovery success. Most importantly, removing this system highly reduces the probability of vehicle components becoming tangled with parachutes during descent. Changes to Sample Container System After further reviewing the sample container system (also known as the hatch system) design modifications implemented into the current design. It was earlier stated that motors would continue to drive for the duration of flight to ensure the door for the payload bay remains shut. It has been decided that the continuously running the motors could jeopardize the reusability of the hatch system. To ensure that the rocket remains shut, the team decided to use continuous servo motors that will lock the door into place and make certain that the door does not go in the reverse direction mid-flight. In addition, the decision was made to have the door rest on the payload container to ensure that it closes effectively and stays level throughout the process. Lastly, instead of a pressure pad, flex sensors were chosen as a better alternative. Due to their dimensions and practicality, the sensors prove to be easy to work with and are predicted to effectively give an output for the door to begin shutting. Lastly, changes of several component dimensions and positions were made. In particular, it was decided that there will only be one avionics sled in the bottom of the sample canister instead of the two avionics sleds on either side. There is now another set of rails on the upper side of the canister that supports the sliding door system. This is explained in further detail in the Hatch Subsystem section beginning on page 23. Page 11 of 139

12 Changes to AGSE System After further review of the AGSE system described in PDR, design changes have been made to each of the subsystems. For the robotic arm, the structural design and location of the motors have changed while the lengths of the segments and degrees of freedom have remained the same. In the original draft, the segments consisted of two parallel plates with rods connecting them for strength and rigidity. This design was discarded because of the difficulties in creating the joints. The new structure will consist of a 3D printed cylinder with a hollow tube running throughout the length of the component. The tube running the length of the arm segments will be used to enclose cables that will run to each of the servo motors. The segments will have a low mass, so the belt system will no longer be necessary. The servos will now be located at each of the joints of the robotic arm. Using this new system has the benefit of reducing the chance of failure in the system because it is simpler and has fewer moving parts than the previous design. The ignition subsystem was originally going to use a ball and screw drive system powered by a stepper motor. The current design exchanges the ball and screw system for a linear actuator that will be set off to the side of the blast plate. The system will still include the igniter being raised up through a hole in the blast shield and into the motor; however, this will now be done at an angle five degrees off of vertical to match the angle of the rocket. The last subsystem, the rail system, has also had a major design change since proposal. The system described in the proposal used a high torque, low rpm motor to rotate an axle that was attached to the hinge. One of the flaws with that system is that the motor had to impart a large force on the axle to fulfill the large moment needed to rotate the rail and rocket. To account for these risks, another linear actuator system was devised to raise the rail and rocket. This involves attaching both ends of the actuator to a pivot which would then be mounted to the base and to the rail. This system reduces the force needed to rotate the rail. III) Vehicle Criteria Selection, Design and Verification of Launch Vehicle Mission Statement The mission of the Illinois Space Society Student Launch Team is to safely launch and recover a reusable higher power rocket meeting all project requirements. The mission includes autonomously inserting a sample into the vehicle and preparing the rocket for launch, reaching an altitude of 3,000 feet, and jettisoning the sample at 1,000 feet above ground level. Requirements Official project requirements and their respective design features and verification methods are given in Table 2 on page 31 of this report. However the team has determined several unofficial requirements to serve as project goals. Many of these are closely related to official requirements. Page 12 of 139

13 1.) The vehicle must conform to the highest safety standards at all times. 2.) The vehicle shall attain a maximum altitude between 2,900 and 3,000 feet. 3.) The vehicle shall be highly reusable, such that the ISS may recreationally launch the vehicle with minimum effort upon competition completion. 4.) The vehicle shall be able to function both with the custom AGSE system, as well as a standard high power rocketry launch rail configuration. 5.) The vehicle shall have a visually appealing design, reflecting the months of extreme effort dedicated to its design and construction 6.) The vehicle design and construction shall serve as a high level learning experience for team members, providing all team members with significant crucial experience in the real world design and engineering process. 7.) The vehicle design must be well defined and reports shall be given with the highest amount of detail possible. Mission Success Criteria The mission shall be deemed successful upon completion of the competition if, while meeting the highest possible safety standards, all mission requirements are satisfied and all competition goals are met. Additionally, the team s personal requirements listed above should be satisfied to deem the mission a complete success. Specifically, the launch vehicle should seal the sample within the sample canister, reach an apogee of 3,000 feet, and then jettison the sample at 1,000 feet above ground level during descent. Additionally, the recovery avionics should record and report the vehicle s flight profile and properly initiate all ejection events. Safety standards will be considered satisfied if all strict requirements are met and at no point during launch operations are any team members or observers placed in a situation of inordinate risk. System Design Review Booster System The vehicle s Booster System includes all vehicle components contained within or attached to the lower body tube, excluding recovery components. This includes the Motor, Fin, and Rail button subsystems. As the booster must contain many crucial flight subsystems, the design of the booster system is relatively constrained. The team only considered options that include the motor, rail Page 13 of 139

14 button, and fin subsystems at the far aft end of the rocket. The only variation in considered options was the location of the hatch and sample canister. The two available empty volumes in the vehicle are the upper portion of the booster and the upper airframe. These were the two locations considered for the location of the sample canister. Due to the requirement that the sample must be jettisoned during descent, the team determined that the best location for the sample is within the upper airframe. This allows the entire upper body tube to be jettisoned, fulfilling the requirement. In past events, the ISS team has attempted to jettison payloads from the booster airframe of high power rockets. However, this was invariably found to add unnecessary complexities to the vehicle and recovery construction and procedures. Thus, it was determined that the booster airframe was not a sufficient location for the sample to be sealed within. The Booster System is inches long and 5.5 inches in outer diameter. The booster airframe is composed of Blue Tube, a lightweight but high strength material specifically designed for use in high power rocketry. The functional requirements of the booster system mainly pertain to the ascent stage of the flight. The booster section houses the motor assembly, the critical components providing the vehicle s propulsion. The booster also serves as the mounting point for the rail buttons. These are critical to the liftoff stage of the flight. Additionally, the fins provide an aerodynamic restoring force to allow the vehicle to maintain a straight flight path under thrust. The booster is responsible for proving the vehicle with a safe and stable ascent and thus is crucial to the success of the mission. In designing the rocket airframe, team members researched various materials for construction of the main body and fins. Initially, aircraft plywood and balsa wood were considered as possible materials for the fins while Blue Tube, carbon fiber, and fiberglass were evaluated for possible use in the main body. Each material was later assessed in regards to its respective advantages and disadvantages as seen in Table 1 below. A score of 5 represents the best possible score in a category, while 1 represents the poorest possible score in a category. Table 1 Vehicle Construction Material Qualities Material Strength Cost Ease of Use Safety Aircraft plywood Balsa wood Blue Tube Carbon fiber Fiberglass Team members first decided on a material for the main body of the rocket. Research into fiberglass revealed that there are many safety hazards when working with this material. Namely, the high weight of fiberglass and the potential to inhale dangerous strands of glass when the fiberglass is cut is a major health hazard and was thus fiberglass ruled out first. The team then debated between carbon fiber and Blue Tube. It was ultimately decided that the added strength of carbon fiber was unnecessary and did not justify its much higher cost. In addition, Blue Tube is Page 14 of 139

15 easier to work with than carbon fiber. Its heat capacity is sufficient to protect against the heat output of the motor, and its reinforced cardboard makeup poses fewer safety concerns when it is being cut. These benefits, combined with its relatively high strength at an affordable price, led Blue Tube to emerge as the team s chosen material for the main body. Focus then shifted to deciding between balsa wood and aircraft plywood for the fins. Team members decided that the material would have to be moderately strong and relatively easy to work with, especially because fins require extensive shaping and sanding before being attached to the rocket. Although balsa wood is extremely easy to cut and shape, it was almost immediately ruled out due to its low strength. Aircraft plywood, on the other hand, was found to be an excellent material for fins that fit both of the team s main requirements: moderately strong and relatively easy to shape. As a bonus, aircraft plywood is not high in price and does not present any unacceptable safety hazards. Due to structural concerns and the reusability requirement, however, the fins will also be lightly reinforced with carbon fiber. A thin skin of carbon fiber will add an immense amount of strength to each fin without adding much weight. A full image of the booster design, as well as a cutaway version, are given below in Figure 2 and Figure 3. The specific components are described in depth in the Subsystem section of this report beginning on page 20. Figure 2 Booster System Page 15 of 139

16 Figure 3 Cutaway View of Booster System Sample Container System The vehicle s Sample Container System includes all components contained within or attached to the upper body tube, excluding recovery components. This includes the Hatch Door mechanism and the nose cone. As previously noted, locations considered for the sample container were the booster airframe and the upper airframe. As described in the Booster System section of this report on page 13, it was determined that the booster airframe was not an appropriate location for the Sample Container. By process of elimination, the only available volume within the vehicle, the upper airframe, was chosen as the location for the sample container. The sample container system is critical to the success of the mission in the context of completing the prescribed task. The main requirement of this system is that it must seal the sample within the vehicle, and maintain this seal throughout the flight. The sample container also houses the equipment used to jettison and recover the sample itself. A full image of the Sample Container system is given below in Figure 4, along with cutaway images of the system in Figure 5 and Figure 6. The components of this system are discussed in further depth in the Hatch Subsystem section of this report beginning on page 23, and a much more detailed set of images is provided. Page 16 of 139

17 Figure 4 Exterior View of Sample Container System Figure 5 Cutaway View of Sample Container Page 17 of 139

18 Figure 6 Zoomed In View of Sample Container Cutaway Recovery System The Recovery System comprises all vehicle components used to ensure safe recovery of the vehicle. This includes the main, drogue, and sample parachutes and their attachment hardware; the avionics bay; and all altimeters. This system also includes the equipment used to attach and ignite the parachute ejection charges. This system is of the greatest importance to mission success as it contains many subsystems critical to the safety of the flight. The Recovery System must properly deploy all three parachutes to provide the vehicle with a safe descent to the ground. Additionally, the vehicle must jettison the sample container system to allow the vehicle to meet project requirements. The Recovery System must be armed through switches on the exterior of the rocket. The system shall also provide an isolated environment for all recovery electronics to function and provide power to these components. A full image of the avionics bay, as well as cutaway images, are given below in Figure 7, Figure 8 and Figure 9. Page 18 of 139

19 Figure 7 Exterior View of Avionics Bay Figure 8 Cutaway view of Avionics Bay showing Altimeters Page 19 of 139

20 Figure 9 Cutaway View of Avionics Bay showing Batteries Subsystem Descriptions Motor Subsystem The first and most critical subsystem of the vehicle is the Motor Subsystem. The motor serves as the vehicle s propulsion system for the flight. The main components of this subsystem are shown above in Figure 3. Shown in orange is the vehicle s RMS 75/3840 motor casing. This component is designed to contain a 75mm reloadable motor and is just under 21 inches in length. In terms of safety, this is possibly the most important flight component. Due to this, the motor casing is professionally made of precisely machined aluminum. This ensures that the propellant has a proper environment in which to burn without adversely affecting the remainder of the vehicle. This component also serves as the lower attachment point for the drogue parachute. The eye bolt for this parachute is shown in Figure 3 in gray and screws directly into the motor case. Shown in green is the motor mount tube. This is a 98mm diameter tube 24 inches long which is composed of Blue Tube. This component is designed to house the motor case separately from the rest of the vehicle. As this motor mount tube is not the same size as the motor case, an adapter is required to fit the motor case within the motor mount tube. This design feature allows significant flexibility in the design, as a 98mm motor can also be flown in the vehicle without any modification. Given in white in Figure 3 above are the vehicle s centering rings. These are used to ensure that the motor mount tube, and thus the motor casing and motor itself, are seated directly in the center of the vehicle. These rings are composed of high quality plywood and are designed for the specific purpose of centering the motor. The vehicle contains three centering rings. One at the Page 20 of 139

21 extreme aft end of the booster tube, one at the top surface of the fins, and one at the top of the motor mount tube. The motor itself used for the flight is the Aerotech L1150 reloadable 75mm motor. An Aerotech motor was chosen as this is a highly reputable company that the team has had significant dealings with in the past. Additionally, Aerotech is one of the most well-known motor manufacturers, and a large number of motor hardware products compatible with Aerotech products are available. The final component of the motor subsystem is the motor retainer, shown below in Figure 10 in grey. This is a high strength aluminum component used to prevent the motor from moving forward or aft during flight. The retainer consists of two pieces: a body and a screw on cap. The body of the retainer is permanently affixed the lowest centering ring. After the motor case is slid into the rocket, the retainer cap simply but securely threads on to the body of the retainer. This prevents the motor from inadvertently moving during flight, but provides a quick method of loading and removing the motor casing. Figure 10 Close up view of Motor Retainer Page 21 of 139

22 Fin Subsystem The vehicle s Fin Subsystem is designed to provide the vehicle with an aerodynamic restoring force that will allow a straight and stable flight. The design includes three trapezoidal fins spaced 120 degrees apart. After researching the construction of rocket fins, team members brainstormed geometric and sizing constraints along with materials. Trapezoidal fins were chosen to allow a larger amount of surface area to be farther away from the fuselage helping to stabilize the rocket. It was decided that the root of the fins will span inches, will have a height of inches and will have a tip span of inches. These dimensions were determined through the OpenRocket simulation software. This allowed the team to change the dimensions of the fins and instantaneously achieve an estimate of vehicle stability. The fins will be level with the aft of the rocket to ensure that the fins do not break upon ground impact. The fins will extend through the body of the vehicle and connect directly to the motor mount tube, as well as the outer airframe tubing. To ensure structural integrity, the fins are attached between the lower and middle centering rings, providing for additional contact surfaces on which epoxy may be applied. Additionally, epoxy will be injected into the airframe via plastic syringes to allow for an even stronger bond. The fins are shown in red in the above images of the booster airframe. Avionics Bay Subsystem The avionics bay is located in the vehicle s center coupler and contains the majority of the components necessary for deploying the rocket s parachutes. Shown above in Figure 4 in blue is the vehicle s coupler, composed of 12 inches of Blue Tube designed to function as a coupler for 5.5 inch Blue Tube airframes. The brown disks shown above in Figure 4, Figure 5, and Figure 6 are the coupler bulkheads, composed of high quality plywood. These provide a physical barrier between the recovery electronics and the remainder of the vehicle. Shown in black are barrier blocks designed to accommodate the E-matches that will ignite the ejection charges. Wires connect to one side of these blocks to the altimeters, and the E-matches are attached to the other side. Also mounted on the bulkhead and shown in red in the aforementioned images are charge cups designed to hold the recovery system s ejection charges. These are small PVC cups that will be filled with black powder and an E-match is inserted. The caps are then covered with foil tape to contain the powder. The final components mounted to the bulkheads are the eye bolts and quick links, shown above in gray. The eye bolts run through the bulkheads and are attached with a nut and washer on each side, as well as a small amount of epoxy. These provide a secure attachment point for the parachute shock cords. The quick links, shown below as gray ovals, are used to attach the parachute shock cords to the eye bolt. These provide for a strong attachment point that may be easily assembled before flight and removed afterwards. Shown in the above mentioned images in green are the rails for the avionics bay. These are composed of threaded aluminum rod and span the length of the coupler. These rods are attached to each bulkhead via a nut and washer on each side. These both hold the bulkheads on the coupler and provide a rail system for which to slide the payload sled into the bay. This avionics sled is shown as a brown rectangle in the cut-away images above. This is a thin sled composed of aircraft plywood with small copper tubes attached to one side. These tubes serve as guides, allowing the sled to smoothly slide on the rails and remain fixed within the system. Page 22 of 139

23 On one side of the sled and shown in yellow are the primary and secondary Stratologger altimeters used to record the flight profile and deploy the main and drogue parachutes. On the other side of the sled, shown as yellow rectangles, are the 9 volt batteries which serve as the power supplies for the altimeters. These batteries are intentionally located on opposite sides of the sled. In the case that a battery breaks free of the payload sled, this will prevent the battery from physically impacting the altimeters and destroying flight critical hardware. The final component of the avionics bay is the switch band, shown above in orange. This is a 3 inch long piece of 5.5 inch Blue Tube mounted to the exterior of the coupler. This serves as a mounting point for the rotary switches which arm the vehicle s avionics. Parachute Subsystem The vehicle will utilize three parachutes: a drogue, main, and sample parachute. These systems are discussed at length in the Recovery Section of this report beginning on page 53. Hatch Subsystem The team researched several design concepts for storing and ejecting the recovered sample. It was determined that having the cargo bay be an existing portion of the rocket body was the best option. Having a separate container hold the payload sample within the body and eject was also considered, however, the team ran into several issues designing a system to close and seal the hatch and payload container door simultaneously. Separately ejecting the payload without damaging any of the electronics controlling the closing system was also a considered factor. The two design considerations that were not chosen for the hatch system include an articulating arm and a rotating hinge described as follows: Articulating Arm: The hatch door would be left open during initial setup, with an articulating arm attached to the door and to the inside of the airframe to close when the sample was loaded on board. This system exhibited the benefit of securely sealing the payload with a simple building technique. However it was determine that the rocket does not contain the interior volume required to store a robotic arm. Rotating Hinge: The hatch door would be left open during initial setup, with the door attached via a rotating barrel-style hinge. Once the sample was loaded, the hinge would be rotated using a DC motor or Servo motor to shut the door. This system also exhibited the benefit of a simple building technique. However this method was found to be highly restrictive in terms of mounting locations for the actuating motors. Page 23 of 139

24 A sliding door mechanism was selected as the best option by the team. The hatch door will be left open during initial setup. Rather than a traditional hatch, the vehicle will utilize the sliding door-style mechanism for the opening and closing of the hatch housed within the airframe of the rocket body. Images of this system are given below in Figure 11 through Figure 17. The hole that will be utilized to receive the payload will be larger than the specified sample to account for any error from the AGSE system. The hatch door will be cut out of a coupler of Blue Tube, since couplers are designed to fit perfectly inside of airframe tubes and to stay consistent with the body material of our flight vehicle. This will ensure that the inner surface of the airframe tube and the outer surface of the door remain flush. A thin metal plate will be attached via epoxy to the bottom end of the door to ensure a flat surface. This thin metal plate is shown in gray in the following images. This subsystem will be controlled by utilizing two 0.3 ounce continuous rotation servo motors displayed in dark gray on the images below. These motors will drive a shaft shown in gray. This shaft will run through two 12-tooth,.5 pitch diameter gears shown in white that are feeding a gear rack also shown in white. The gear racks will be attached to the back of the hatch door plate with epoxy. These motors will be fixed via brackets shown in a copper color upon rails that are shown in gray. These rails run through the upper portion of the sample canister from the bulkhead of the sample parachute to the bulkhead of the main parachute. Both of these bulkheads are shown in tan below. The rail-motor interface vertically supports the motors and rear section of the door, while adding structural support to the sample container as a whole. The front side of the door will be supported by the hollowed, rectangular sample container shown in a dark brown. The gear racks on the door itself will be positioned to slide across the payload container, making certain that the door is guided and will not fall into the payload area when the door is closing. The amount of torque required by the servo motors will be rather low since the door is not exerting its full weight onto the motor. The motors do not work to lift the door against gravity, but rather just to slide it horizontally and overcome frictional forces. The door will be programmed to close by the bending of three 2.2 Sparkfun Flex Sensors fed through slits in the box shown in green. As these sensors bend, a different resistance output is measured. The team will be utilizing the bending of these sensors by having the door begin to shut once a certain resistance is met. This resistance will be determined at a later time after empirical testing. All system electronics will be mounted to a plywood sled shown in tan. This sled will be mounted to two additional rails via copper brackets near the aft end of the sample container. The entire door subsystem will be powered by a 9 volt battery shown in gray, with an Arduino Pro Uno Rev3 board controlling the motors and receiving the input from the flex sensors. The Arduino board is shown in green and gray on the inner surface of the sled adjacent to the 9 volt batteries. Since the Arduino offers multiple pin outputs and inputs, the motors and a resistance flex sensor will be running off of the Arduino. The two altimeters used in this section will be placed on the back side of the sled. They are also modeled in green and gray as well as powered by a 9 volt battery each mounted on the inner surface of the sled. An electrical schematic is given below in Figure 18 Page 24 of 139

25 Once the payload is dropped into the canister, the flex sensors will bend giving feedback to the Arduino which will wait 10 seconds before sending a signal to the motors to begin spinning. The motors will be programmed to spin for a certain distance until the door is safely shut. Once the door is shut, the motors will stay locked, securing the door. Figure 11 View of Sample Container without Airframe Tubing Figure 12 Side view of Sample Container without Airframe Tube Page 25 of 139

26 Figure 13 Exterior view of Sample Container showing the Hole in the Airframe Figure 14 Cutaway view of the Sample Container Showing Flex Sensors in Green Page 26 of 139

27 Figure 15 Cutaway view of Sample Container showing Altimeters Figure 16 Cutaway view of Sample Container showing Gear System Page 27 of 139

28 Figure 17 Cutaway view of Sample Container showing Batteries Figure 18 Sample Container Electronics Schematic Page 28 of 139

29 Rail Button Subsystem The rail button subsystem is responsible for holding the vehicle to the rail during the initial stage of the flight. These components are shown in yellow below in Figure 19 and Figure 20. These will be standard 1515 rail buttons designed to work on a 1.5 inch slotted rail. Each rail button will be attached to a mounting point secured to one of the vehicles centering rings. This mounting point will consists of a plywood block with a T-nut. This allows the rail buttons to easily screw in and out in the case that one needs to be replaced but also provides for a secure mounting configuration. Three rail buttons were chosen for the sake of redundancy. In the case that a rail button fails, two will remain and will be capable of holding the vehicle to the launch rail. No more than three rail buttons were chosen, as additional buttons increase the drag on the rail as the vehicle launches. Figure 19 Image of Rail Button Placement Page 29 of 139

30 Figure 20 Zoomed in Image of Rail Button Performance Characteristics There exists a multitude of performance characteristics relevant to the success of the mission. A large number of these relate to the flight of the vehicle, such as flight apogee, maximum Mach number, and rail exit velocity. In order to meet competition requirements, the vehicle is required to reach an altitude of 3,000 feet. A motor was selected that will allow the vehicle to reach this target altitude with a mass of 45 pounds. In the case that the final vehicle weighs less than 45 pounds, ballast mass will be added to achieve this launch mass. Although modeling has provided evidence that the vehicle will reach this target altitude, this characteristic will not be empirically verified until the completion of the full scale test launch. The maximum Mach number is also a characteristic of note. As the vehicle approaches the transonic flow regime (Mach numbers roughly between 0.8 and 1.2), compression waves will form, vastly affecting vehicle performance. However the maximum Mach number of the vehicle is only predicted to reach 0.41, well short of the transonic region. In fact, this Mach number is much closer to a value at which air density is essentially constant, and compressibility effects can be ignored completely. Although this characteristic may not be empirically verified until the full scale test flight, it is extremely unlikely that the vehicle would approach this transonic region. The rail exit velocity of the vehicle is highly important to early flight stability. The recommended minimum speed of the rocket exiting the rail is 45 ft/s. This will allow the fins to experience sufficient airflow to provide an appropriate restoring force in the case that the rocket deviates from the desired flight path. The simulated rail exit velocity with the anticipated 45 pound vehicle is 65 ft/s, well above the recommended minimum. Although this value has been predicted Page 30 of 139

31 with what is believed to be relatively high accuracy, this characteristic again may only be verified through the full scale test flight. Another performance characteristic of note is the ability of the recovery system to properly deploy the vehicle s parachutes. The ejection procedure will be simulated on the ground through a series of charge tests upon the completion of vehicle construction. This will verify that the chosen ejection systems are robust enough to sufficiently deploy the parachutes. It is also important that the parachutes are deployed at the correct altitudes. The altimeter settings may be verified through simple ground tests and inspections; however, the functional verification of this characteristic may again only be verified during the full scale launch. One final performance characteristic of the vehicle is the ability of the hatch system to seal the payload within the vehicle, and remain sealed throughout the flight. Sufficient ground testing will be undertaken to ensure that the hatch door will seal upon insertion of the sample payload. Testing will also be done to determine the capabilities of the hatch door to remain closed, however the full scale test flight will provide the final verification that this characteristic has been met. Vehicle & Recovery Requirements Table 2 given below lists all project requirements placed on the vehicle, as well as the design features that satisfy these requirements and the methods of verification. Table 2 Vehicle Requirements and Verification Requirement Design Feature Verification Method (Inspection, analysis, and/or testing) Deliver payload to altitude of 3,000 feet AGL. The selection of the Aerotech L1150 motor shall allow the rocket to easily exceed this altitude. Ballast will then be added as required to decrease altitude to the target 3,000 feet. Modeling, simulation, and flight test. Barometric altimeter will be used to record official altitude. Page 31 of 139

32 Designed to be recoverable and reusable. Have a maximum of 4 independent sections. Rocket limited to a single stage Capable of being prepared for flight within 2 hours. Capable of remaining in launch ready position for 1 hour. Capable of being launched by a 12 volt direct current firing system. Use a solid motor propulsion system using APCP that is approved and certified. Total impulse provided by launch vehicle should not exceed 5,120 Newton-seconds. All materials used in construction have been evaluated to ensure durability. The rocket will employ a series of three parachutes to ensure that the each section of the rocket lands. with less than 75 ft-lbf of kinetic energy. During apogee and descent, the rocket has been designed to break apart into 4 independent sections: booster, coupler, upper airframe tube, and nosecone. The rocket has been designed to only carry one single-stage motor, the Aerotech L1150. Vehicle components such as the motor retention system and payload sleds have been chosen to allow for quick assembly. All power supplies are designed to function for well in excess of this time limit. The vehicle employs a standard motor igniter compatible with the standard 12 volt system. An Aerotech L1150 reloadable rocket motor will be used and has been certified by the TRA. The total impulse of the Aerotech L1150 is 3517 Ns. Hand calculations of the kinetic energy of each rocket section upon landing; adequate construction techniques; visual inspection. Modeling and visual inspection. Modeling and selection of motor. Assembly testing and practice prior to launch events. Testing all electronic components to ensure their have sufficient power lifetimes. Design and inspection. Design and inspection. Design and inspection. Page 32 of 139

33 Team must provide an inert or replicated version of the motor, with matching weight and size. Burst/Ultimate pressure Vs. Max Expected Operating Pressure shall be 4:1, with supporting design documentation. Low cycle fatigue life must be at least 4:1. Pressure vessels must contain solenoid pressure relief valves that sees complete pressure of tank. Complete pedigree of tank must be provided, including the application of the tank, its history, number of pressure cycles put on the tank, by whom and when. Launch and recover a subscale model of the full-scale rocket prior to CDR that should perform similarly to the full-scale model. Prior to FRR, the full-scale rocket shall be launched and recovered, in order to ensure the vehicle and recovery system function properly, in fully ballasted configuration, with no additional modifications being made after successful completion of test flight. Maximum budget of $10,000 A hollow motor shell will be produced and filled with ballast to match the weight of the functional motor. The vehicle does not contain any pressure vessels. The vehicle does not contain any pressure vessels. The vehicle does not contain any pressure vessels. The vehicle does not contain any pressure vessels. A subscale rocket will be constructed upon completion of PDR. The rocket will match the aerodynamics and performance of the full scale vehicle as closely as possible. The team plans to construct and launch the fully assembled final rocket prior to FRR. The current budget of the final system is $ and future spending will be tracked to ensure the limit is not exceeded. Design and inspection. Design and inspection. Design and inspection. Design and inspection. Design and inspection. To be confirmed through subscale flight test and analysis. To be confirmed through full scale flight test and construction Budget planning and inspection. Page 33 of 139

34 Deploy drogue chute at apogee. Deploy main chute at a lower altitude. Each independent section of the launch vehicle shall have a maximum kinetic energy of 75 ft-lbf. Electrical circuits of the recovery system shall be completely independent of electrical circuits of the payload. The recovery system must contain redundant, commercially available altimeters. Dedicated arming switches shall arm each altimeter from the exterior of the rocket. Altimeters must have dedicated power supplies. The primary altimeter will detect when the rocket has reached apogee. It will then trigger an E-match to ignite the black powder charge on the lower bulkhead of the coupler, separating the booster and coupler and releasing the drogue parachute. The primary altimeter will detect when the rocket has descended to 1,000 feet. It will then trigger an E-match to ignite the black powder charge on the upper bulkhead of the coupler, separating the coupler and upper airframe tube and releasing the main parachute. The main and sample parachutes were chosen to provide enough drag to slow each rocket section to a terminal velocity that allowed for an acceptable maximum kinetic energy. See Kinetic Energy calculations in Mission Performance Criteria section. The only function of the recovery electronics are to record the flight profile and ignite ejection charges. Each ejection event is controlled by fully redundant and independent avionics components. Each of the four altimeters has a dedicated rotary switch on the exterior of the rocket. Each altimeter has its own battery power supply. Modeling and testing of parachute deployment systems. Modeling and testing of parachute deployment systems. Hand calculations of the kinetic energy of each rocket section upon landing. To be confirmed based on empirical test data. Design and inspection. Design and inspection. Design and inspection. Design and inspection. Page 34 of 139

35 Arming switches must be capable of being locked in the ON position. Removable shear pins shall be used for both the main parachute and drogue parachute compartments. Electronic tracking devices shall be installed in the launch vehicle. Any untethered section or payload component shall have its own electronic tracking devices. Recovery systems electronics cannot interfere with any other on-board electronic devices during flight. Recovery system altimeters must be physically located in a separate compartment from other radio frequency transmitting and/or magnetic wave producing devices. The rotary switches chosen are capable of being locked in the ON position. Shear pins connect the booster and coupler, which separate to deploy the drogue parachute. Shear pins also connect the coupler and upper airframe tube, which separate to deploy the main parachute. A final set of shear pins connects the upper airframe to the nose cone. The vehicle will utilize the GPS capabilities of the Telemetrum altimeter, as well as a radio frequency tracking device. The separately jettisoned sample container contains the GPS tracking Telemetrum altimeter. No onboard components are expected to interfere with the recovery electronics. The only radio transmitting component will be exterior to any avionics bays, and no components produce magnetic waves. Design and inspection. Design and inspection. Design and inspection. Design and inspection. Interference and functional testing of recovery components upon avionics bay construction. Design and inspection. Page 35 of 139

36 Recovery system electronics shall be shielded from onboard transmitting devices. Recovery system electronics shall be shielded from devices that may generate magnetic waves. Recovery system electronics shall be shielded from any other devices that may interfere with proper operation of recovery system electronics The recovery electronics are located in a physically separate compartment from the radio frequency transmitter. No onboard components generate magnetic waves. No onboard components are expected to interfere with recovery electronics. The electronics are physically isolated and shielded regardless. Design and inspection. Design and inspection. Testing to ensure recovery systems will function when all electronics are running. Risk Management Given below in Table 3 is a summary off project risks, as well as their likelihoods, impacts and solutions. Table 3 Summary of Risks, Likelihood, Impact and Solutions Risk Likelihood Impact Solution Necessary materials are out of stock or come late. Low to Moderate. Construction delay. Order necessary materials far ahead of time and order from store that currently has stock. Hatch fails and materials/payload are prematurely ejected from rocket. Low to Moderate. Bodily injury to any bystanders in the launch area. Carefully manufacture and test hatch and locking mechanism to ensure it can repeatedly withstand the aeronautical load. Communication problems. Low to Moderate. Members unaware of responsibilities, tasks are either duplicated or not done. Frequent meetings and s from managers clarify distribution and progress of tasks. Page 36 of 139

37 Unable to launch on desired day due to weather or other unrelated reasons such as driver doesn t arrive. Low to Moderate. Unable to obtain desired data from launch, resulting in not being able to complete competition requirements. Launch well ahead of deadlines and have multiple launch days set up as backup. Team members unable to perform assigned task. Moderate. Additional responsibilities are now thrust upon remaining members, which may lead to rushed and subpar results. Consistently monitor progress at frequent meetings to minimize impact and plan around any potential failings of a team member. Misuse of potentially dangerous power tools and machinery. Moderate. Possible grievous bodily injury. All members using such tools will be properly taught on how to safely use the said equipment. Additionally, members unfamiliar with the machinery will be put under experienced supervision when using the tools. Electric Matches. Low. Minor Burns. Electric matches will be stored in a safe, grounded condition inside a certified explosives storage container. When stored, ends of matches are grounded to prevent a net charge buildup. Devices turned off before installation to prevent misfire. Black Powder. Moderate. Serious burns to skin. Overexposure can result in irritation to skin and eyes. Fire. Black powder will be stored in a safe, inert, certified explosive proof container. The powder will only be handled and used by the team mentor. Gloves and eye protection will be worn at all times. Page 37 of 139

38 Ground Testing Dangers. Moderate. Minor or serious bodily injury. Potential damage to the rocket, electronics, and payload. Fire. All members will wear appropriate safety gear when testing the rocket. Appropriate distance from rocket will be maintained at all times. All handling of explosives and electric matches will be under the supervision of the safety officer. Safety and readiness checklists will be created and followed. Launch Dangers. Moderate. Damage to rocket and payload. Potential serious bodily harm. Fire. All launches will be conducted in compliance with NAR High Power Rocket Safety Code, Federal Aviation Regulations, and all other laws, regulations, or safety codes that pertain. The launches will take place at locations that have standing FAA waivers. All team members will be familiarized with the NAR safety code and will have signed safety agreements. The team mentor will be present to ensure safety and proper handling. Safety and readiness checklists will be created and followed. Rocket Motor (Ammonium Perchlorate). Moderate. Potential unplanned ignition could result in bodily harm. Overexposure could cause irritation to skin, eyes, digestive tract, and mucous membranes. Motors will be stored by the team mentor and all handling and usage will be done under his supervision as well as that of the safety officer. Gloves and eye protection will be worn when any handling or work is being done on the motor. Page 38 of 139

39 Environmental Safety. Moderate. Damage to the environment. Members will be informed of the environmental regulations regarding materials. Particularly the proper disposal of hazardous or environmentally harmful materials. Battery Failure. Low to Moderate. Unable to collect necessary data from instruments or hatch is unable to function properly. Test batteries before placing them in the body of the rocket to ensure they re working properly. Project falls behind schedule. Moderate to High. Work becomes rushed or does not get finished. Frequent meetings and s from managers clarify distribution and progress of tasks to ensure tasks are completed on time. Project goes over budget. Moderate. Unable to buy necessary components or spare parts, forcing design changes. Develop thorough design that specifies all needed components, account for spares for critical components, and reuse components from previous projects whenever possible. Unable to afford necessary materials. Low. Unable to obtain desired components, forcing design changes. Become aware of possible budget constraints well ahead of deadlines. Contact potential sponsors to secure funding/materials for project. Key team members leave project. Low. No one capable of taking over exmember s responsibility. Make sure at least two team members know how to perform every step necessary to complete project. Necessary equipment unavailable. Low to Moderate. Unable to work on project until required machinery is available, may force design changes. Plan far ahead with managers/owners of equipment to ensure that equipment is available when they are needed. Page 39 of 139

40 Launch rail/pad structurally fails. Low to Moderate. Rocket launches horizontally if at all. Repeatedly test launch system/ ensure that system is reinforced beyond that strictly necessary for the weight of the rocket. Hatch does not shut. Low. Payload left not secure. Run door closing mechanism many times through different environment conditions. Failure of electrical equipment. Low. No parachutes/hatch system. Catastrophic failure. Ensure components work smoothly through stress tests. Component comes loose mid-flight. Moderate to High. Either everything else stays secure or the loose components knocks the altimeters/batteries off. Secure all components tightly. Motors too weak to close door. Low. Door for hatch will not effectively shut. Ensure servos motors have enough torque to close door and stay in place during flight. Flex Sensors not triggered. Low. Hatch will not be activated. Test sensor to react using custom made PVC payload. Gap between outer wall of rocket and door. High. Weakened aerodynamics. Shave down outer wall close to the door to create a smaller gap. Sample is not placed in rocket. Low. Hatch door will not shut. Make several practice runs to ensure reliability. Igniter misses motor. Low to Moderate. Motor will not ignite. A guide will be used to direct the igniter into the motor. Robot doesn t pick up capsule. Moderate. Robot won t know it s not there, hatch will not shut. Extensive testing will be done to optimize predictability. Robot drops capsule. Low. Hatch door will not shut. Strong gripper motors and adequate gripper friction will be implemented. Linear actuator will not raise rail. Low. Ineligible for Maxi- MAV competition, will need to be raised by hand. The actuator will be selected to be able to lift more than the estimated weight. Page 40 of 139

41 Launch stand tips over. Low Possible damage to rocket Moment calculations have been done, and extra weight will be added to the stand to secure it to the ground. Pin locking mechanism fails Moderate Loss of redundant locking The linear actuator will lock and prevent the rail from moving out of position Limit switch for igniter fails Low Actuator may break, and igniter could fall out of the motor The switch will be tested to ensure reliability Limit switch for rail actuator fails Low Actuator may break, rail could fall back down if it is not locked first The switch will be tested to ensure reliability. Necessary Components and Risk/Delay Impact All components necessary to the construction of the vehicle are listed in both the Mass Statement and Budget sections of this report beginning on pages 50 and 107 respectively. However several characteristic components will be discussed here. The vehicle contains many critical components that are able to be removed from the rocket. This includes such components as the motor casing, altimeters, parachutes and payload sleds. Although these components are of course critical to flight operations, they are not necessarily critical to the construction process. Thus, delays in obtaining these parts will have little to no impact on the construction process. As long as these components are obtained and tested before launch, they should pose no risk to mission success. Construction of the avionics bay is not likely to cause a large delay impact on the project. Avionics bays are in many ways independent from the rest of the vehicle, due to their required isolation. Thus, issues with coupler components such as the coupler tube itself or the threaded rods is not likely to have a significant impact on the project timeline. The Arduino microcontroller utilized by the hatch system is a component with moderate delay impact. Although the vehicle is capable of flying without the hatch system being operational, the team intends to run through the full competition procedures during the full scale test launch. In the case that the microcontroller and its software are not complete before the full scale test flight, the risk of the hatch system not functioning during the competition activities is somewhat elevated. Many other components are highly critical to the construction process, and represent a very high delay impact. The motor mount components represent the largest possible risk to the construction timeline. The booster airframe is the first and most critical system that must be constructed. A delay in obtaining a motor mount tube or airframe tube will cause a massive delay in construction. The vast majority of flight critical elements are in some way integrated within the Page 41 of 139

42 booster airframe during construction. Thus, issues with the booster tube or motor mount tube will exhibit a harsh impact on the project timeline. Due to the complexity of the system, issues relating to the vehicle Hatch Subsystem are likely highly disadvantageous to the project timeline. Although not necessarily structurally integral to the vehicle design, the hatch system is absolutely necessary for the completion of mission requirements. Due to the vast amount of manufacturing and testing that must be undertaken to ensure the hatch system will function properly, it is highly important that the hatch construction proceed relatively smoothly and without delay. Planning of Manufacturing, Verification, Integration and Operations The manufacturing and assembly of the flight vehicle will be broken down into several individual sections and will take place periodically throughout the academic year after all designs have been finalized. Safety has been the primary factor while determining these construction techniques. Safety equipment such as gloves, safety glasses, and earplugs will be worn when necessary throughout the build process. Members of the team building the flight vehicle will rotate in turns to insure a small group of students working at any given time. Work instructions will be written before all build meetings and all work will be documented at the end of build meetings to eliminate any progress confusion for following build meetings. The projected construction techniques are subject to change as the team approaches obstacles in the manufacturing process. An assortment of tools will be used from basic office supplies to power tools. Basic supplies will include: pencils and pens, masking tape, mixing sticks, sandpaper of assorted grit, a ruler, drafting squares, a level, an X-Acto knife, a C-clamp, razor saw, thread locker and rubbing alcohol. Power tools will include: a drill and bits, a Dremel tool and a palm sander. Epoxy will be used for bonding major areas of the flight vehicle. The amount applied will be determined by the structural integrity and consequential drag effects while in flight. Since there are many hazards associated with exposure to epoxy fumes, great caution will be used when handling this resin system. Before construction, all parts will be inventoried, weighed, cleaned and labeled. All parts will then be checked for proper fitting. The vehicle will be assembled as fully as possible without permanently attaching any components to ensure all components were manufactured to correct dimensions. General construction practices will include marking all hole locations, confirming all hole and insert sizes, and double checking locations before drilling. Surfaces that will have epoxy applied will be sanded with 60 grit or coarser sandpaper and later payload cleaned with rubbing alcohol. The projected plan is to start with the construction of the motor mount. While constructing the motor mount tube and centering rings several things will be accounted for. Motor retention will be ensured by a screw-on motor at the base of the rocket. The Aeropack retainer will be mounted to the vehicle s lowest centering ring, which must be mounted at the extreme aft end of the rocket. Three center rings will be used for additional support and ease of alignment. The location of the center rings will be marked on the motor mount and body tube in three different locations: the top ring slightly below the motor mount tube, the middle ring to be aligned with the top of the fins, and the bottom ring to align the retainer with the bottom of the rocket. Page 42 of 139

43 Rail button positions will be marked on the airframe. The rail buttons will be attached before the motor mount is fixed inside of the rocket. T-nut interfaces will be created on the inside of the rocket. These T-nut interfaces will consist of a T-nut inserted into a block of plywood. These interfaces will be mounted to each of the three centering rings. The motor mount will be inserted into its marked location in the booster airframe at a later time. The inside of the booster airframe and the fin slots will be sanded. Epoxy will be applied to the top of the center rings. For the bottom center rings, epoxy will be applied through a hole for the top ring and through a fin slot for the middle ring. The avionics bay will be assembled through many subparts. The bulkheads will have threaded rod rails, eyebolts for parachutes, charge cups and terminal blocks. A switch band will be created and attached next. Finally, a sled will be created by marking out electronics, attachment placement and attaching tubing to the bottom for the rail guides. There are also several methods being considered for the construction of the fins, which are composed of aircraft grade plywood wrapped in a carbon fiber skin. The flight vehicle will be constructed with through-the-wall fins. These must be able to fit between the middle and bottom center rings. A fair amount of epoxy will be applied between the fins and the body tube for support. Internal fillets for the fins will be used since the fins must be fit tight to the motor mount tube and the center rings must fit snug to the top and bottom of the fins. Fin alignment will be insured through the use of a simple jig consisting of slots placed 120 degrees apart. The CAD model of the hatch system will be refined before construction of the hatch occurs to ensure proper dimensions and verify predicted alignment. All parts will be further researched to verify commercial readiness of all components. The hatch door mechanism will be constructed with the same techniques of measuring and marking all parts as described above to ensure proper fitting. More careful and precise measurements will be made since this system is very small. In addition, any amount of Blue Tube removed on the airframe for the purpose of receiving the sample payload will be mimicked on the opposite side to maintain a balance of weight and ensure stability. The sample payload receiving end will be sealed using the same techniques mentioned above in the sample canister section. The opposing end will be sealed with epoxy. Finishing the build process will include priming and painting with the possible application of decals. Pressure relief holes in the airframe sections will determined and drilled to allow pressure to equalize in flight. Verification of the vehicle construction will be completed on individual components on a case by case basis. As components are completed, experienced team members and the team mentor will inspect all connections and other construction features to ensure the manufacturing process proceeded as planned and provided for sufficient structural integrity. Additionally, each time a component is completed, its integration with other completed parts will be tested. This includes interfaces such as the coupler to airframe connections and other sizing constraints such as parachute bay lengths. Inspecting components as they are completed will allow the team to catch any errors or defects as early as possible, and rectify these errors without delay. Page 43 of 139

44 As many methods of structural testing are destructive in nature, the team will not be able to directly test and measure the failure strengths of the constructed vehicle components. Components will be inspected and manually loaded to ensure all connections are secure. The team will be able to simulate the loadings on the vehicle components expected during flight, however the full scale test flight will provide the true empirical evidence that the vehicle is completely structurally sufficient. The full scale test flight is by far the most important test undertaken during the project, as it will prove that all vehicle components and subsystems are able to function when fully integrated. One critical system of the vehicle, the hatch system, will be fully tested on the ground upon completion of its construction. As the hatch is actuated before flight, the full functionality of the door mechanism will be rigorously tested to ensure the door is capable of closing securely every time. The integration with the AGSE systems may also be tested on the ground. This includes loading the vehicle on the rail, inserting the sample payload, erecting the launch pad and inserting the motor igniter. As these are the most complicated systems that the team has the least amount of experience with, it is highly important that these systems be verified through significant ground testing. As the loadings and physical requirements of these tests are relatively fixed and well known, the testing will essentially consist of running the system many times in succession to prove that the components have been properly constructed and integrated. Confidence and Maturity of Design Given that this is relatively early in the design process, the team is highly confident in the design of the vehicle. It is expected that minor changes will occur before CDR, however all design issues explored by the team have been matched with solutions. Although many new team members completed much of the design and analysis work for the vehicle, all designs were overseen and reviewed by team members with significant rocketry experience. Additionally, the team has been able to draw on the past experiences of ISS Student Launch Teams. This competition marks the fifth time that ISS has competed in this event, and the current team has access to many old design documents and reports. This allows the team to learn from the past failures and successes of other teams. All of the flight critical components used in this vehicle have seen significant usage in previous high power rocketry applications. This includes use by both past ISS teams, as well as the general rocketry community. The team believes the current design defines an innovative system for completing the mission requirements, without sacrificing confidence in flight safety and reliability. In terms of maturity, the team believes the current design is relatively well defined. Team members met for an average of six hours per week since the initial request for proposals, and the design was allowed to evolve freely throughout this time period. After defining a system that meets all mission requirements for the project proposal, the vehicle details were continuously analyzed and redefined to achieve the best possible preliminary design. Although changes are always expected in the final design and construction processes, the team believes the vehicle defined within this support is more than sufficient to meet or exceed all project requirements. Page 44 of 139

45 Dimensional Drawings Dimensioned drawings of the vehicle are given below with all dimensions given in inches. The first image is a drawing of the vehicle as a whole. Critical vehicle dimensions are given, such as the length of the body components and parachute compartments. The lengths of the drogue, main, and sample parachute compartments are 18.5, 14, and inches, respectively. Both body tubes are 48 inches in length, and the nose cone is 21 inches long. All outer diameters of airframe tubing components are inches. Figure 21 Full Vehicle Dimensions Figure 22 below gives the dimensions of the avionics bay, with all dimensions given in inches. Of note are the overall length of the coupler, 12.5 inches, and the switch band length of 3 inches. The majority of the additional components are determined by the sizes of purchased components and are not variable in terms of vehicle design. Page 45 of 139

46 Figure 22 Avionics Bay Dimensions The following dimensional diagrams are of the payload canister. The hole to receive the sample payload will be 1.25 in width and 5.5 in length. The door will be slightly bigger than the hole to ensure that the hole has been properly sealed. The door is currently 1.75 in width and 6 in length. This allows for 2.5 inches of coverage on either end along the longer side of the hole. There is a little bit more coverage on the width of the hole with inches on either side due to less space constraints. Further placement specifications are displayed in the diagrams below. Page 46 of 139

47 Figure 23 Sample Container Dimensions The dimensioned drawing given below provides the geometric characteristics of the vehicle s fins, with the trailing edge of the fin pointing upwards. The fin geometry was defined in an effort to provide the vehicle with a sufficient stability margin to provide for a stable flight. The fin contains a 0.75 inch straight tab, representing the portion of the fin which fits through the booster tube and extends to the motor mount tube. Page 47 of 139

48 Figure 24 Fin Dimensions Given below are electrical schematics for the recovery systems. The first image pertains to the avionics bay of the vehicle. This system is used to deploy the main and drogue parachutes. Of note is the fact that all recovery electronics are completely redundant. The batteries, switches, altimeters, and charges are all completely independent, providing a full additional system in the case that any single component fails. Page 48 of 139

49 Figure 25 Main and Drogue Parachute Avionics The schematic given below pertains to the recovery electronics within the sample container. These electronics are used to deploy the sample parachute after the sample container has been jettisoned. This system again includes redundant altimeters, however there is only a single deployment charge activated by each altimeter. Additionally, one of the altimeters in this system is a Telemetrum rather than a Stratologger. This Telemetrum provides the GPS tracking capabilities for the sample container, as required to meet mission criteria. Figure 26 Sample Parachute Avionics Page 49 of 139

50 Mass Statement Table 4 given below provides masses for all components designed to be on the rocket during flight. Components are broken down by subsystem, and materials are listed when applicable. For components that include multiple units on the vehicle, such as the fins (3) or hatch gears (4), the masses given are for the combined net mass of these components, not a single unit. The team believes the current mass statement is relatively accurate for this early stage of the project. The vast majority of masses have been retrieved from manufacturer specifications or data ISS has collected throughout previous projects. An epoxy mass of 4 pounds has been estimated to account for the weight due to construction. Although the team believes the statement accurately reflects the current design mass of the vehicle, it is understood through experience that the vehicle will almost certainly grow in mass by the end of the project. Given this reality, the team has allowed for the final launch mass to be roughly 45 pounds. This represents a mass increase of almost 50%. The team does not anticipate a mass growth this large, however this will allow for flexibility in attempts to reach the target altitude. The current motor and parachute choices have been designed to optimally meet contest requirements for the 45 pound rocket. Upon measurement of the vehicle s final constructed mass, the team will be able to add ballast mast to the vehicle in order to meet the designed launch conditions as closely as possible. It will be much easier for the team to add mass to the vehicle if necessary, compared to the difficulty of removing hardware mass from a rocket. As most rocketry components somehow affect the safety or structure of the vehicle, it is typically unreasonable to remove any significant amounts of mass from the vehicle. Table 4 Mass Statement Subsystem Component Material Mass (lb) Booster Body Tube (4 ft, 5.5 inch) Blue Tube 2.38 Three Fins Plywood & Carbon Fiber 1.32 Aerotech L1150 Motor APCP/Aluminum 8.10 Motor Mount Tube Blue Tube 0.80 Motor Retainer Aluminum 0.63 Three Centering Rings Plywood 0.53 Drogue Parachute Nylon 0.69 Eye bolt Steel 0.10 Shock Cord (20 ft) 1/4 inch tubular Kevlar 1.00 Page 50 of 139

51 Epoxy 2.00 Booster Total Coupler Body Tube (12 inches) Blue Tube 0.66 Bulkheads (2) Plywood 0.28 Rails (2) Aluminum 0.13 Sled Plywood 0.30 Avionics 2 Batteries and 2 Stratologgers 0.10 Key switches and wiring Eye bolts (2) Forged Steel 0.20 Quick links (2) Steel 0.25 Charge Cups PVC 0.14 Nuts (6) Steel 0.17 Washers (8) Terminal Blocks (4) Plastic/metal 0.10 Eyebolt Nuts (4) Plated steel 0.20 Eyebolt Washers (4) Plated steel 0.04 Epoxy.50 Coupler Total: 2.92 Upper Airframe Tube Body Tube (4 ft, 5.5 inch) Blue Tube 2.38 Includes Hatch Components Bulkheads Plywood 0.28 Sample 0.25 Key switches and wiring Eye bolt Forged Steel 0.10 Quick links (2) Steel 0.25 Charge Cups PVC Nuts (6) Steel 0.17 Washers (8) Sample Parachute Nylon 0.68 Page 51 of 139

52 Shock Cord (30 ft) for main and sample parachutes 1/4 inch tubular Kevlar 1.50 Servo Motor x Gears x Gear Racks Battery/Power Source (don't include altimeter batteries) x D-shafts Flex sensors 0.10 Microcontroller Main Parachute Spectra and Nylon 1.56 Avionics 2 Batteries and 2 Altimeters 0.10 Terminal Blocks (2) Plastic/metal Eyebolt Nuts (2) Plated steel Eyebolt Washers (2) Plated steel Epoxy 1.50 Upper Airframe Total 9.46 Nosecone Nosecone polypropylene 1.30 Total Mass Total Mass with Ballast Page 52 of 139

53 Recovery Subsystem Recovery System Analysis The recovery system design of this vehicle is highly dependent on the total vehicle mass during the descent stage of the rocket. The mass statement given above in the Launch Vehicle Criteria represents the current mass inventory of all parts included in the current design. It is expected that this mass will grow throughout the design and construction process. To limit the effects of this mass growth, the recovery systems have been designed to function for a rocket weighing 45 lbs, about 50% heavier than the current design. Parachute sizing analysis has been completed to determine which recovery devices will most effectively allow the vehicle to safely return to the ground and ensure reusability. Due to previous projects with similar vehicles, the ISS team has a small inventory of appropriately sized parachutes. Analysis was initiated by investigating whether any parachutes currently in the ISS team s possession would be sufficient for use on this project. Simulations in OpenRocket and the team s custom simulator suggest that the Iris Ultra 96 parachute is an appropriate main parachute for this mission. The simulated descent speed of the vehicle under this parachute is feet per second. The SkyAngle 36 parachute was also shown by the simulations to be a sufficient drogue parachute for this vehicle. The simulated drogue descent speed is about 60 ft/s. Due to the high quality of the Iris Ultra 96 parachute already owned by the team, another Iris Ultra parachute, this one with a 60 diameter, was investigated for the parachute on the sample container. Simulations again showed this parachute to be a sufficient option, allowing for a descent speed of ft/s. Images of the Iris Ultra and Skyangle parachutes are given below, retrieved from the FruityChutes and B2Rocketry websites, respectively. Figure 27 Skyangle Parachute Page 53 of 139

54 Figure 28 Iris Ultra Parachute The simulated parachute descent speeds were then confirmed by hand by calculating the terminal velocities of the vehicle. The equation for terminal velocity is given below: VT = 2mg ρac d Where VT is the terminal velocity, m is the mass of the components descending under the parachute, g is the local acceleration due to gravity, ρ is the density of air, A is the surface area of the parachute, and Cd is the drag coefficient of the parachute. Page 54 of 139

55 Drogue Parachute (SkyAngle 36 ) Descent Velocity m=45 lbm (total) lbm (propellant)= lbm g= ft s 2 ρ = lbm ft 3 A=π(1.5 ft) 2 =2.25π ft 2 Cd= lbm ft s VT= lbm ft π 1.34 = ft/s Main Parachute (Iris Ultra 96 ) Descent Velocity m=45 lbm (total) 15 lbm (jettisoned upper section) lbm (propellant)= lbm g= ft s 2 ρ = lbm ft 3 A=π(4 ft) 2 =16π ft 2 Cd= lbm ft s VT= 2 = ft/s lbm ft 3 16π 2.2 Sample Container Parachute (Iris Ultra 60 ) Descent Velocity m= 15 lbm g= ft s 2 ρ = lbm ft 3 A=π(2.5 ft) 2 =6.25π ft 2 Cd=2.2 VT= 2 15 lbm ft s lbm ft π 2.2 = ft/s Page 55 of 139

56 Safe drogue descent speeds are typically considered as those below 80 ft/s, and landing speeds below 25 ft/s are typically considered safe. As shown the calculations above, the chosen parachutes exhibit terminal velocities well below these maximum limits. The relevance of these values is further discussed in the Flight Profile Simulation section of this report beginning on page 61. Attachment scheme As the recovery systems are one of the most safety critical portions of the vehicle, all attachment schemes have been designed to use proven high strength methods to achieve maximum safety and reliability. The drogue parachute is stored in the booster airframe of the rocket, above the motor mount tube and below the avionics bay. Tubular Kevlar shock cord is used to attach this parachute to the vehicle. The shroud lines of the parachute are passed through a loop in the shock cord, and the parachute itself is then passed through the looped shroud lines. This allows for a simple but secure method of attaching the parachute to the shock cord. The shock cord is then attached to the motor casing via a steel quick link and a forged steel eye bolt. The eye bolt screws into a threaded slot at the top of the motor casing specifically designed for this purpose. A steel quick link is then used to attach the Kevlar shock cord to the eye bolt. These quick links allow for secure attachment throughout flight and descent, while still providing team members a quick method of assembling the vehicle on launch day. The drogue parachute is attached to the bottom of the avionics bay in much the same manner. Another forged steel eye bolt will be permanently affixed to the bottom bulkhead of the avionics bay. This attachment point will feature a nut and washer on each side of the bulkhead. Additionally the nuts and washers will be epoxied to the bulkhead in order to further ensure structural integrity. An image of the bulkhead parachute attachment point is shown below in Figure 29. Figure 29 Bulkhead Assembly Page 56 of 139

57 The vehicle s main parachute is stored in the lower portion of the upper airframe, above the avionics bay and below the hatch system. This main parachute is attached to a Kevlar shock cord in the same manner as the drogue parachute. However, this main parachute only has one hard attachment point to the vehicle. An eye bolt is attached to the upper bulkhead of the avionics bay in the exact configuration used for the drogue parachute s bulkhead attachment point. This parachute does not have an upper attachment point to allow for the sample container to be jettisoned from the main portion of the vehicle. This is discussed further in the following Deployment Process section of this report, beginning on page 57. The vehicle s third and final parachute is the sample parachute, stored in the upper portion of the upper airframe above the hatch system and below the nose cone. Again, this parachute will attach to the shock cord in the same manner described for the drogue and main parachutes. The shock cord s lower attachment will also use the same eye bolt/bulkhead configuration utilized for the attachment points on each end of the avionics bay. The upper attachment point will consist of a steel quick link attached to an attachment hook which comes standard on the nose cone. Deployment Process The deployment process for this vehicle is based on the standard dual deploy system with modifications to allow the vehicle to jettison the sample as required. At apogee, which is targeted for 3,000 feet, the primary altimeter in the avionics bay will send a signal to an E-match, igniting a black powder charge on the lower bulkhead of the avionics bay. This ejection charge will break the shear pins attaching the avionics bay and booster airframe and allow the drogue parachute to deploy. All vehicle components will remain tethered together during this portion of the descent. In the event that this ejection charge does not sufficiently eject the drogue parachute, the secondary altimeter will send a signal to a second E-match, igniting a second black powder charge on the lower bulkhead of the avionics bay. This will occur with a delay of one second after apogee. In the likely event that the first charge does deploy the drogue parachute, the secondary charge will harmlessly ignite in the open air. An image depicting the vehicle upon completion of the drogue event is given below. Page 57 of 139

58 Figure 30 First Stage of Recovery Deployment After descending under the drogue parachute to 1,100 feet above ground level, the primary altimeter in the sample container will send a signal to an E-match, igniting a black powder charge on the upper bulkhead of the sample container. At this point, the primary altimeter within the sample container will send a signal to an E-match on the top bulkhead of the sample container. In the same manner described above, this system will push apart the nose cone and upper airframe portions of the vehicle and deploy the sample parachute. This system again uses a redundant deployment method where all components are fully independent. The secondary ejection event will occur at 1,050 feet above ground level. An image depicting the vehicle configuration upon completion of the sample event is given below. Figure 31 Second Stage of Deployment Process After descending to 1,000 feet above ground level, the primary altimeter in the avionics bay will fire the primary main parachute ejection charge. This will break the shear pins holding the upper airframe to the avionics bay, pushing the components apart and deploying the main parachute. At this point, the booster and avionics bay form one tethered section of the vehicle, while the upper airframe and nose cone form a second tethered section. At this point the two tethered sections are independent, fulfilling the requirement that the sample must be jettisoned at 1,000 feet above ground level. The booster and avionics bay will then descend to the ground under Page 58 of 139

59 this main parachute. The main parachute deploy utilizes the same fully redundant scheme used for the drogue deployment. This includes a completely independent battery, rotary switch, altimeter, E-match, and ejection charge. The secondary ejection event will take place at 900 feet above ground level. An image depicting the vehicle configuration upon completion of the main/jettison event is given below. Figure 32 Third Stage of Deployment Process The drogue and main parachutes to be used for this vehicle have been used for previous flights of Illinois Space Society Student Launch projects. Due to this fact, the team has significant experience with packing these parachutes such that they will properly deploy. Upon vehicle construction, the team will work to determine the best method of packing the parachutes, and will include the detailed process within the preflight checklist. Recovery System Testing Test plans for ejection charges at this point are tentative and highly dependent on the vehicle s construction timeline. However, the same testing process will be carried out regardless of the timing. Ejection charge testing will be completed with the vehicle loaded as closely as possible to the launch configuration. The major exception will be the removal of fragile components to be replaced by equivalent mass. Ejection charges and parachutes will be loaded into the vehicle in the same manner as during the actual launches. However, instead of connecting the E-matches to the altimeters, a wire will run from the E-match to a remote firing system operated by team members. This will allow the team to prop the vehicle up on a test stand and remotely ignite the ejection charge to deploy the parachutes one at a time. A sufficient number of shear pins will first be determined based on the weight of the individual rocket sections and the expected accelerations during flight. It is important that shear pins break due to the force of the ejection charges, but also that they remain intact before the charges have ignited. Based on the size and number of shear pins used to hold the vehicle together, black powder charges will be sized to sufficiently break these shear pins. Testing will then either Page 59 of 139

60 confirm the adequacy of the chosen charge sizes or provide a starting point for further refinement of charge size. Electronic testing may be done on a much more flexible schedule than the charge testing. Power lifetime testing can be done to prove that the vehicle can remain operational on the launch pad in excess of one hour. This will simply be completed by leaving electrical components powered on and operational for over an hour on their respective power supplies. Interference testing will also be carried out to prove that none of the onboard electrical components interfere with other critical components. This may simply be done by powering on all electronics in the same area and testing the functionality of each individual component. However, in the case that the operational modes of multiple components do conflict, it would be necessary to load the full vehicle and determine if the physical separation inherent in the design provides enough shielding for the components to function. These functional tests will be relatively simple, and include changing the position and elevation of the rocket to ensure location transmitters and altitude recording devices are capable of performing their desired functions at the same time. Robustness of Recovery Components The major components of the recovery system are the eye bolt attachments, quick links, shock cords, and the parachutes themselves. Both the eye bolts and quick links are composed of steel. Although relatively dense, steel provides one of the strongest readily available structural materials with a tensile yield strength of about 366,000 psi. The relatively high weights of these components are accepted as an acceptable trade off given the structural integrity of the components. Additionally, the eye bolt is composed of forged steel rather than steel bent into the shape of an eye bolt. This causes the eye bolt to be one solid piece that will not bend open under the loadings of deployment. The closing mechanism of the quick links consists of a threaded cap which screws onto a threaded bolt shaped attachment point on the quick link. This cap is capable of being quickly but robustly sealed to ensure the quick link remains closed during flight. All three of the vehicle s shock cords will be composed of high strength quarter inch tubular Kevlar. As strong as steel is, Kevlar is in fact even stronger, with a tensile strength of about 520,000 psi. The final and most important components of the recovery system are the parachutes themselves. The parachute canopies are composed of high strength rip stop nylon. The shroud lines and bridles are constructed out of high strength spectra, nylon, and Kevlar. All of these materials exhibit sufficient strength to declare the components sufficiently robust. It is important to note that all of these components have been heavily utilized for high power rocketry applications in the past. In particular, the ISS team has utilized the same eye bolts, quick links, and Kevlar components for many rockets spanning several years. These are high quality components designed to be used in applications where strength is the highest priority. Additionally, the parachutes used for this vehicle are professionally manufactured and designed to be used on rocket applications such as these. Nothing about the vehicle s recovery system falls outside the realm of standard recovery procedures or loadings. The parachutes will be used for their exact design purposes and are thus determined to be robust enough to survive deployment, deceleration and descent. Page 60 of 139

61 Mission Performance Predictions Mission Performance Criteria The Mission Performance will be evaluated through a set of characteristic criteria. These include the vehicle apogee, landing speed, and kinetic energy. The vehicle should reach an altitude of 3,000 feet, within a reasonable margin due to local atmospheric conditions on the day of launch. Although no hard limits are directly placed on landing speed, a common safety guideline is that vehicle components should land with speeds between 10 and 25 ft/s. The kinetic energy is strictly limited by competition rules and must be below 75 ft-lbf for any independent or tethered section of the vehicle. This requirement places a firm restriction on landing velocity in relation to individual section masses. Additionally, several qualitative criteria factor into the performance of the mission. These include the ability of the hatch system to accept and seal the sample payload, as well as the ability of the recovery system to return the vehicle to the ground in a recoverable manner. This is not restricted to the landing speed of the rocket but includes the drift distance. The vehicle shall not drift too far away or into an obstructed area, as this may cause the rocket to be unrecoverable. Flight Profile Simulations The flight profile for this mission has been simulated through two separate methods. The simulation was completed by modeling the vehicle in OpenRocket software, and the second simulation was completed via a custom program written by the team. This custom program was implemented in MATLAB and is attached to this report as Appendix C. Both simulations utilize the simulated component weights, vehicle data, and thrust curve. The team designed simulator uses thrust curve data and vehicle characteristics to model the forces on the vehicle. The simulator then numerically integrates the vehicle s equations of motion to obtain a flight profile. The integrator utilized a modified version of the Runge-Kutta fourth order scheme for a high level of accuracy. This simulation is completed in two dimensions and directly solves for horizontal and vertical position and velocity at each time step. The data output by this program also allows the prediction of vehicle drift, descent rates, launch rail exit velocity, maximum velocity and apogee. This simulator includes the effects of drag on the vehicle, as well as the actual design characteristics of the vehicle and parachutes. The custom simulator has the benefit of modeling the flight characteristics of the jettisoned sample container separate from the profile of the main portion of the vehicle. Flight simulation results are given below, with discussion following the plots. All simulations have been completed using a rocket weight of 45 lbs. In the case that the final vehicle does not weigh this much, ballast mass will be added to match the simulated conditions as well as possible. Page 61 of 139

62 Figure 33 OpenRocket Altitude Simulation Figure 34 OpenRocket Velocity Simulation Page 62 of 139

63 Figure 35 Custom Altitude Simulation Figure 36 Custom Velocity Simulation Page 63 of 139

64 Figure 37 Custom Velocity Simulation Results for Sample Container Figure 38 Altitude Simulation Comparison between OpenRocket and Custom Simulation Page 64 of 139

65 Figure 39 Velocity Simulation Comparison between OpenRocket and Custom Simulation For the most part, the two simulations agree relatively well. The team believes both simulations have their beneficial qualities. The benefit of the custom simulation over OpenRocket is that the team is highly aware of the assumptions made in the custom simulation, and how forces on the vehicle are treated. The calculations evaluated by OpenRocket are all completed outside of the user interface, and the team does not know any assumptions or decisions made by OpenRocket. However, OpenRocket is a widely known and often used piece of simulation software. The mathematics behind the OpenRocket simulation are likely more robust than the custom program. The apogee reported by the OpenRocket simulation is 3316 feet, and the altitude reported by the custom simulation is 3046 feet. Through past endeavors, the Illinois Space Society has found that OpenRocket consistently overestimates the apogee of ISS s rockets. Due to this, the team has a higher confidence in the predictions given by the custom simulation. However until about 2500 feet, the altitude profiles for both simulations agree very well. The velocity profiles agree almost exactly until apogee. During the drogue descent phase, there is also a relatively high amount of disagreement between the two simulations. This is due to the different descent velocities calculated by the two programs. OpenRocket shows a drogue descent speed of almost 100 ft/s, but the custom simulation shows a drogue descent speed of about 50 ft/s. This difference is likely due to the difficulty of modeling a Skyangle parachute for simulation purposes. In previous flights with this parachute, Page 65 of 139

66 the team recorded a descent velocity of 70 ft/s on a 53 lb rocket. Several online descent calculators for Skyangle parachutes yield descent speeds between 55 and 65 ft/s. Due to these factors, the team believes that the custom simulation s descent speed under drogue is much more accurate than the OpenRocket Simulation. In the case that the rocket needs to descend more quickly to eliminate excessive drift, the drogue parachute shroud lines can be reefed a small amount to increase descent velocity. Despite the disagreement in drogue descent speed, the descent speeds under the main parachute agree very well. Due to the close correlation between these two values, the average main descent speed will be used for calculations. This main parachute descent speed is ft/s. Due to the method used by the custom program to terminate the integration, the custom simulation shows the altitude instantaneously decreasing to zero at about 300 ft above ground level. This does not represent the actual predicted motion of the vehicle. As the vehicle has achieved steady descent before this sharp decrease occurs, this is not expected to have any effect on the accuracy of the simulations. Regardless, the team intends to correct this issue moving forward. Another major benefit of the custom simulation is that the sample container may be simulated separately from the remainder of the rockets. The results of this simulation yield a sample container descent and landing velocity of ft/s. In summary, the two simulations agree to a relatively high degree. Both predict a stable flight with a safe main descent speed. The altitude of the custom simulation is believed to be more accurate, as is the case for the drogue descent speed. As shown above, the simulated descent speeds correlate relatively well to theoretical terminal velocity calculations. A drogue descent speed of under 80 ft/s is typically considered safe for a high power rocket, and all both the team s simulation and the theoretical value fall well below this value. Although the separate simulations gave somewhat different drogue descent rates, the team s past data and experience with OpenRocket lead to the conclusion of a 60 ft/s descent speed, agreeing very well with the ft/s calculated value. Additionally the team has flown a heavier vehicle with this drogue parachute before, further increasing confidence in the sufficient sizing of this drogue parachute. Landing speeds for high power rockets are typically considered safe if the vehicle lands under 25 ft/s. Again, both the simulated and calculated values predict the vehicle to land well short of this upper limit. The simulated and calculated values only differ by about 0.4 ft/s, a margin the team considers well within the acceptable range for this stage in the design. Additionally, it is important that the vehicle not descend too slowly, as this will lead to excessive drift distances and make recovery more difficult. A common rule of thumb is that the vehicle should descend at no less than 10 ft/s. All predicted values provide evidence that the vehicle will descend greater than this minimum recommended speed. It is also important to note that the kinetic energy of the vehicle is also highly relevant to launch safety. Both the kinetic energy and drift distance of the vehicle upon landing are analyzed in the following sections of this report on pages 72 and 73. Page 66 of 139

67 The apogee of the vehicle was also analytically calculated using the following equations. ya = yb + yc yb = ( M ) ln (T Mg kv2 ) 2k T Mg yc = ( M ) ln (Mg+kv2 ) 2k Mg v=q (1 e xt ) (1+e xt ) t= I T x = 2kq M q= T Mg k k= 1 2 ρ C d A Where ya is the apogee altitude, yb is the altitude achieved during boost, and yc is the change in altitude during the coast to apogee. M is the full mass of the rocket, T is the average thrust, and g is the local acceleration due to gravity. V is the velocity of the rocket at burnout. X and k are simplifying parameters to ease calculations, and k is a drag coefficient, characterizing the drag in the absence of the knowledge of velocity. Cd is the vehicle drag coefficient and A is the cross sectional area of the vehicle. For ease of calculation these equations are completed using metric units, and the final answer is converted to imperial units for comparison to simulation data. ρ = kg/m3 Cd = 0.75 A = m2 T = 1150 N I = 3517 N-s M = kg g = 9.81 m/s2 k=0.5*1.225*0.75* = q= 1150 ( ) x= = =.253 t= = v= (1 e ( ) ) (1+e ( ) ) = yb= ( ) ln (1150 ( ).00704( )2) T Mg yc= ( ) ln ( (.00704)( )2) Page 67 of 139

68 ya = meters ya=ya*3.28 ft/m= feet Apogee = feet This hand-calculated apogee value is only about 20 feet off from the apogee of 3046 feet reported by the custom simulation. This relatively small margin of error is acceptable, and in fact expected at this stage of the design. Thrust Curve The motor thrust curve is given below, as retrieved from Thrustcurve.org. This data is believed to be highly accurate, as was utilized in both the OpenRocket and custom simulations. Figure 40 Aerotech L1150 Thrust Curve Most importantly, this motor achieves maximum thrust very shortly after ignition. This will allow the vehicle to achieve a sufficient velocity off of the launch rail to provide for a safe flight. A common rule of thumb is that a rocket should achieve at least 45 ft/s of velocity upon exiting the rail. Using this Aerotech L1150, both simulation programs yield a rail exit velocity of about 65 ft/s. This velocity is more than sufficient to allow the fins to create an aerodynamic restoring force and stabilize the flight of the vehicle. Page 68 of 139

69 The vehicle s Center of Pressure location was determined analytically through the use of the Barrowman equations. These represent an analytical method of finding the center of pressure based on vehicle geometry. These equations are given below. Nose Cone Terms: (CN)N =2 XN=0.466LN (CN)F=[1 + Fin Terms: R R+S ] [ S 4 N ( d ) ( 2 L f C R +C T ) 2 XF=X B + X R 3 (C R+2C T C R +C T ) [(C R + C T C RC T C R +C T )] Center of Pressure: Xbar= ((CN) N X n + (CN) F X F ) 1 (CN) R (CN)R is the sum of the coefficients (CN)N+(CN)T+(CN)F. Xbar is the final answer found through this analysis, and is a measure of the distance between the center of pressure and the tip of the vehicle s nose cone. The meanings of the quantities in the above equations are given below. Page 69 of 139

70 The characteristics of the vehicle are given below: Quantity Value (inches) Ln d Cr Ct S Lf R Xr Xb N 3 Note that the all transition terms are zero, since the vehicle airframe has a constant diameter. XN=0.466* = inches (CN)F=[ ] [ 4 3 ( ) ( )2] =9.215 XF= ( ) [( )] = Xbar= ( ) = inches Thus, the center of pressure is calculated to lie inches from the tip of the nose cone. This is relatively close to the location given by the OpenRocket simulation, which yields a center of pressure 87 inches from the nose cone. For the purposes of center of pressure location, it is believed that OpenRocket is likely more accurate than the manual calculation. However the value calculated by hand produces a more stable vehicle configuration than the OpenRocket simulation. Thus, regardless of which estimate is more accurate, the vehicle is still predicted to be stable. The center of gravity used for stability analysis is the simulated value retrieved via OpenRocket simulations. As shown below, this location is 199 cm, or 78.3 inches from the nose of the vehicle. Although not strictly impossible, analytically determining the center of mass of Page 70 of 139

71 each individual vehicle component is relatively difficult in for certain components. The assumptions made to analytically complete this calculation would cause a large reduction in the accuracy of the calculated value. Thus, the center of gravity simulated via OpenRocket will be used to determine the stability margin. Upon construction of the vehicle, the center of mass will be found empirically by locating the balance point of the rocket. The vehicle stability margin is defined as: SM = (Cp-Cg)/D Where Cp is the center of pressure, Cg is the center of gravity and D is the rocket diameter. The standard recommended stability margin is between 1 and 2 calibers, where a caliber is defined as the rocket diameter. An under stable vehicle has a stability margin under 1 caliber, and will experience aerodynamic moments detrimental to the flight safety and stability. An over stable vehicle has a stability margin over 2 calibers, and may turn into the wind causing an elevated amount of horizontal motion to occur. The OpenRocket simulation gives a simulated stability margin of calibers. Using the analytically derived Center of Pressure location inches, the stability margin is calculated to be calibers. This value falls just on the lower edge of an over stable configuration. The vehicle s stability will need to be monitored upon construction to determine the final stability margin. However both of these predicted values imply that the vehicle configuration is stable. The OpenRocket stability simulation is given in the figure below. Figure 41 Vehicle Stability Margin Below are the kinetic energies of each section of the rocket upon landing. These figures were calculated by hand using the equation for kinetic energy given below: Page 71 of 139

72 E k= 1 2 mv T 2 Where Ek is the kinetic energy upon landing, m is the mass of a given rocket section, and VT is the terminal velocity of that rocket section. Again, note that to account for mass growth during design and construction, the recovery systems have been designed for rocket sections weighing roughly 50% more than the current design. Previously hand-calculated values of terminal velocity were used as well. Kinetic Energy of Booster upon Landing m=23 lbm (total) lbm (drogue parachute and shock cord) lbm (propellant) = lbm=0.532 slugs VT= ft/s Ek=0.5*0.532 slugs*( ft/s) 2 = ftlbf Kinetic Energy of Coupler upon Landing m=4 lbm=0.124 slugs VT= ft/s Ek=0.5*0.124 slugs*( ft/s) 2 = ftlbf Kinetic Energy of Upper Airframe Tube upon Landing m=12 lbm (total)-3.74 lbm (main parachute, sample parachute, and shock cords) =8.26 lbm=0.257 slugs VT= ft/s Ek=0.5*0.257 slugs*( ft/s) 2 = ftlbf Kinetic Energy of Nosecone upon Landing m=2 lbm= slugs VT= ft/s Ek=0.5* slugs*( ft/s) 2 = ftlbf Page 72 of 139

73 As shown by these calculations, no vehicle section is expected to approach the design limitations of 75 ft-lbf of kinetic energy. It is important to note that these calculations were based on a vehicle using the current recovery system with 50% mass growth. Even with this large safety margin, the vehicle is expected to fall about 33% under the maximum energy limit. Drift To analyze drift, simulations were completed both in the team s custom simulation and in OpenRocket. The results of these simulations are given in the tables below. Note that in the first table, the drift distances at 5 and 10 mph are actually less than the drift distance with no wind. This is because the wind direction acts against the rocket s horizontal flight direction. Since the vehicle will be launched at 5 degrees off of vertical, the drift in no wind is predicted to be higher than the drift with 5 or 10 mph winds. As OpenRocket does not allow the user to change wind direction, it is not able to calculate the vehicle s drift in the case that wind acts to push the vehicle further in its direction of travel. Due to the fact that the wind will almost certainly not act directly parallel to the horizontal flight direction of the rocket, these values should be taken as upper and lower bounds on the drift of the vehicle during the actual flight. Although these predictions are relatively large for high wind speeds, this is to be expected with 20 mph winds and slow descent rates. None of the drift distances are expected to be prohibitively large due to the onboard tracking devices and the relatively large and open recovery area. Table 5 Drift Predictions with Wind Acting against Vehicle s Direction of Travel Wind Speed [mph] OpenRocket Prediction [feet] Custom Prediction [feet] Table 6 Drift Predictions with Wind Acting in Vehicle s Direction of Travel Wind Speed [mph] Custom Prediction [feet] Page 73 of 139

74 Interfaces and Integration Payload Integration The payload integration plan for the vehicle is relatively complex considering the requirement that the sample container must be jettisoned during descent. However the processes of sealing and jettisoning the sample are treated as design drivers for this project, ensuring that the vehicle is highly compatible with all payload requirements. The sample container is integrated within the upper airframe. This component is typically a relatively empty volume in the airframe and serves as an optimal location for the sample canister. Additionally, isolating the payload systems in the upper system prevents them from interfering with any components in the booster airframe or avionics bay. As these two sections contain the vast majority of the flight critical hardware, it is highly beneficial that the sample payload be isolated from these systems. Another critical aspect of the sample container location is that no other components will interfere with payload components. The payload will be located in the completely isolated compartment, safe from the influence of other subsystems. To ease integration of the payload systems during launch preparation, the sample container will feature a rail system similar to the system used in the avionics bay. This allows critical payload components to be easily removed from the vehicle for inspection and preparation and easily inserted back into the vehicle during assembly. The bottom bulkhead of the payload container is a bulkhead that is permanently epoxied to the upper airframe. Four threaded aluminum rods run through this bulkhead and serve as the container s rail system. All removable components will be mounted on sleds with tubular rail guides, allowing the components to directly slide in and out of the container. The top bulkhead may then be inserted onto the rails and secured via nuts and washers to seal the container. All stresses placed on the payload section will be transmitted through the four aluminum rails. This removes all loadings from the critical payload components themselves. This is fairly similar to the avionics bay system, which utilizes two aluminum rails. However, this sample container will feature four rods, causing the sample container to be even more robust than the avionics bay. Thus, loadings will be transmitted through the sample container, but not through the payload equipment itself. Internal Interfaces The vehicle design includes many internal interfaces critical to mission success. The first major interface is the connection between the avionics bay and the booster airframe. The coupler is specifically designed to insert within the booster airframe. However, it is important that these portions be able to easily slide apart during the drogue deployment event. Due to this, the coupler and booster tubes may need to be sanded or otherwise adjusted to provide for a smooth interface. These components must separate easily at the desired flight event, and they must also remain connected prior to this event. To achieve this goal, removable nylon shear pins will be used to attach avionics bay and the booster airframe. These shear pins will insert through holes in both Page 74 of 139

75 the component tubes. These size and number of shear pins will be determined through charge testing upon vehicle construction. The avionics bay and booster airframe are kept physically separate by a plywood bulkhead. This bulkhead will be fixed to the avionics bay via a series of threaded aluminum rails and will provide a barrier between subsystem components. The interface between the avionics bay and the upper airframe is identical to the connection described above. The coupler tube of the avionics bay will slide directly into the upper airframe tube. Shear pins will keep the vehicle together before the main parachute deployment event but will allow the vehicle components to separate upon the ignition of the ejection charges. An identical bulkhead will be utilized to cap the upper end of the avionics bay. The interface between the nose cone and the upper airframe is also very similar to the two aforementioned connection points. Similar to the coupler, the nose cone is also sized to slide directly into the upper airframe. Again this joint may need to be adjusted to ensure a smooth sliding surface. Shear pins will also be used at this interface, as the nose cone must eject to deploy the sample parachute after the sample container has been jettisoned from the vehicle. One interface that has not yet been thoroughly discussed is the connection between the motor casing and the motor mount tube. Since the motor case is designed for a 75mm motor, and the motor mount tube is 98mm, a motor adapter is required for the system to function. The component used will be an Aeropack 75/98 Motor Adapter assembly. This system consists of three aluminum rings. The top ring utilizes a screw clamp to tightly attach to the upper portion of the motor case. The second ring slides and is allowed to push flush against the motor case s aft closure. The third ring acts as a spacer, and is inserted between the motor retainer body and the motor retainer cap. This system provides a secure method for using a 75mm motor in a 98mm motor mount. Other critical interfaces are the connection points between the parachutes and the bulkhead hard points. However, these connections are discussed in detail in the Recovery System section of this report beginning on page 56. Interfaces between Vehicle and Ground Other than Launch system connections discussed in the following section, the vehicle design does not contain any mechanical connections to the ground. The vehicle is also not electrically connected to the ground other than through the motor ignition system, again discussed below. The vehicle will utilize a Telemetrum altimeter in the sample container to meet the GPS tracking requirement. This altimeter transmits a 70cm HAM radio signal that can be received and interpreted by a ground station. This ground station will consists of a Teledongle ground station radio that attached to a compatible Windows, Mac, or Linux computer. Interfaces between Vehicle and Launch System Based on the design of the AGSE and the hatch system, there will be no interface between the two electrically. Instead of having the vehicle communicate with the AGSE s computer, a timer will be implemented to offset the time from when the robotic arm drops the payload to when the rail starts to rise. That delay will give the hatch door time to close before the rail starts to rise. The reason the vehicle will not communicate with the AGSE is because it would add more complexity to the system without any justification. The way the system works is that when the sample is Page 75 of 139

76 detected by the sensors in the vehicle, the hatch would close after waiting a specified period of time for the robotic arm to move any of its parts out of the way if need be. If the two systems were to communicate with each other the same thing were to occur, it would just be waiting for the robot to move out of the way. One reason for having the vehicle talk to the AGSE was so that sensors placed on the robot could be used to tell the vehicle that the robot was out of the way. Doing that was deemed unnecessary because the AGSE will be executing the same procedure every time with the same weights meaning that the time needed should remain the same. A timer can be used instead with extra time added to account for the error. There will be an interface physically, as well, between the vehicle and the AGSE. The vehicle will be connected to the AGSE system via rail buttons that are meant to slide up and down the rail. This connection will allow the AGSE to raise the vehicle to the five degrees off of vertical and also give it guidance when it is launched. The other way the vehicle interfaces physically with the AGSE is through the ignition system. The AGSE will autonomously interface with the vehicle when the igniter system is used to insert the igniter into the vehicle s motor. This interface is necessary to ignite the motor which makes the ignition system a critical portion of the AGSE s design Safety Preliminary Checklists Recovery Preparation Checklist 1. Prepare Recovery Electronics a. Assemble avionics bay payload sleds, check that all connections are secure i. Altimeters should be wired to switches, batteries, and two terminal blocks each ii. Insert and connect fresh batteries b. Lock altimeter switches in off position c. Attach e-matches to altimeters via terminal blocks d. Turn on altimeters and check continuity, then turn off altimeters (3 beeps for continuity) e. Slide avionics sleds into couplers and attach bulkheads i. Insert sleds so that the altimeters face the key switches ii. Thread a nut and washer on each bulkhead on each rail (6 total nuts and 6 washers) f. Check altimeters are off, and attach ejection charges to terminal blocks 2. Pack drogue parachute a. Packing Procedure to be determined through assembly testing and practice 3. Insert Drogue parachute into booster airframe a. Attach quick link to eyebolt on motor case, confirm quick link is closed b. Insert wrapped shock cord into booster, pack down firmly Page 76 of 139

77 c. Push packed drogue parachute into booster, with the protector facing upwards d. If the parachute is too tight, adjust packing to make the package wider or longer as necessary e. Attach the upper quick link to the bottom eyebolt of the avionics bay, ensure the link is closed 4. Pack main parachute and flame retardants a. Packing Procedure to be determined through assembly testing and practice 5. Insert main parachute into airframe a. Attach quick link to avionics bay, confirm quick link is closed b. Insert wrapped shock cord into airframe, pack down firmly c. Push packed main parachute into airframe, with the protector facing downwards d. If the parachute is too tight, readjust packing to make the package wider or longer as necessary Motor Preparation Checklist (Assembly to be completed by Mentor) Adapted from Manufacturer's instructions 1. Assemble Forward Closure a. Apply lubricant to all threads and O-rings b. Insert the smoke charge insulator into the smoke charge cavity c. Lubricate one end of the smoke charge element and insert it into the smoke charge cavity 2. Assemble Case a. Deburr the inside edges of the liner tube b. Insert the larger diameter portion of the nozzle into one end of the liner, with the nozzle liner flange seated against the liner. c. Install the propellant grains into the liner, seated against the nozzle grain flange. d. Place the greased forward seal disk O-ring into the groove in the forward seal disk. e. Insert the smaller end of the seal disk into the open end of the liner tube until the seal disk flange is seated against the end of the liner. f. Push the liner assembly into the motor case until the nozzle protrudes approximately 1-3/4 from the end of the case. g. Place the greased forward (1/8" thick X 2-3/4" O.D.) O-ring into the forward (bulkhead) end of the case until it is seated against the forward seal disk.78 h. Thread the forward closure assembly into the forward end of the motor case by hand until it is seated against the case i. Place the greased aft O-ring into the groove in the nozzle Page 77 of 139

78 j. Thread the aft closure into the aft end of the motor case by hand until it is seated against the case Launcher Setup Checklist 1. Lower the launch rail 2. Slide the rocket on the launch rail a. Ensure team members are supporting the weight of the rocket b. If the rail buttons do not slide smoothly, rotate the vehicle rather than applying more pressure 3. Power on the AGSE a. Master switch will be turned on b. The computer will receive power i. LED flashes showing power is connected to the computer c. The commands will be paused once the computer is booted up d. 4. Pause button will be pressed a. AGSE will start its commands i. Sample payload will be picked up by the robot ii. Sample payload will be placed in the rocket iii. Hatch door will shut and the robot will move out of the way simultaneously iv. Rocket will be erected to 5 degrees off of vertical v. Mechanical locking mechanism will engage vi. Igniter will be inserted into the motor 5. Turn on altimeters and check settings and continuity a. Primary altimeter: i. 3,1,10,10,10 beeps for main deployment altitude ii. Series of beeps for last flight data iii. Series of beeps for battery voltage (Volts, tenths of Volts) iv. Three quick beeps for continuity b. Secondary Altimeter i. 4,9,10,10 beeps for main deployment altitude ii. 5 second siren for apogee delay iii. Series of beeps for last flight data iv. Series of beeps for battery voltage (Volts, tenths of volts) v. Three quick beeps for continuity 6. Attach the launch controller to the motor igniter 7. The all systems go light will be activated after passing safety verifications Page 78 of 139

79 Launch procedure 1. Proceed to the safe area 2. Acquire signal from the vehicle's transmitters 3. Launch the Vehicle Troubleshooting 1. Altimeters a. If incorrect settings are reported, connect altimeters to a computer to reset settings b. If continuity is not confirmed, check that connections between altimeters, terminal blocks, and E-matches are secure c. If the altimeter doesn't power on, check key switch and power supply wiring 2. Motor Doesn't Ignite a. Wait for RSO clearance to approach the pad b. Confirm that launch controller is connected to igniter c. If so, disconnect launch controller and remove/inspect igniter d. If necessary, disarm altimeters and remove the rocket from the rail for further inspection of the motor assembly Post flight inspection 1. Wait for rocket to land 2. Upon range clear: retrieve rocket and sample, check for undetonated charges and remove 3. Return to safe area 4. Remove altimeters from coupler and collect data 6. Turn off all avionics and store for transport 7. When travel back is finished, clean all dirty components, remove power sources from avionics, and store all materials for future flights Comprehensive Checklist Pre-Launch - Day Before: 1. Check that mentor has: a. Correct Aerotech L1150 Motor b. Correct charge size for each separation event Charge sizes to be determined Page 79 of 139

80 2. Check that all flight hardware is stored for transportation to launch site a. Booster Airframe b. Motor casing c. Motor Adapter (three pieces) d. Motor Forward Seal Disk e. Main, Drogue and Sample Parachutes f. Main, Sample and Drogue shock cords g. 8 quick links h. Coupler (assembled with sled, end-cap bulkheads, and altimeters) i. Flat Screwdriver for Rotary Switches j. Upper Airframe k. Screws for upper Airframe attachment l. Payload fairing (assembled with electronics) m. Shear Pins n. Motor retainer ring 3. Check that all backup equipment and tools are prepared to complete any necessary final fixes or alterations a. Phillips screwdriver for screws b. Small screwdriver for altimeter contacts c. Adjustable wrenches 4. Check that all ground support equipment is packed a. Ground Station Antenna b. Laptop with ground station software c. Ethernet Cable d. GPS tracker e. Binoculars 5. Check that all team members have read or heard safety briefing and are informed of their responsibilities Pre-Launch - Day Of: 1. Pack equipment for travel, as listed above 2. Travel to launch location 3. Unpack equipment at launch site Page 80 of 139

81 4. Perform preflight checks of AGSE and Hatch Systems a. AGSE i. Make sure power reaches all components ii. Ensure safety switch pauses commands iii. Check that master switch functions properly iv. Check that limit switches will stop the actuators v. Test that the motors run vi. Complete a dry run of the system 5. Assemble avionics bay and hatch payload sleds, check that all connections are secure a. Altimeters should be wired to switches, batteries, and two terminal blocks each b. Insert and connect fresh batteries 6. Lock altimeter switches in off position 7. Attach e-matches to altimeters via terminal blocks 8. Turn on altimeters and check continuity, then turn off altimeters (3 beeps for continuity) 9. Slide avionics sled into coupler and attach bulkheads a. Insert sled so that the altimeters face the key switches b. Thread a nut and washer on each bulkhead on each rail (4 total nuts and 4 washers) 10. Check altimeters are off, and attach ejection charges to terminal blocks 11. Pack drogue parachute Packing procedure to be determined through assembly testing and practice 12. Insert Drogue parachute into booster airframe a. Attach quick link to eyebolt on motor case, confirm quick link is closed b. Insert wrapped shock cord into booster, pack down firmly c. Push packed drogue parachute into booster, with the protector facing upwards d. If the parachute is too tight, adjust packing to make the package wider or longer as necessary e. Attach the upper quick link to the bottom eyebolt of the avionics bay, ensure the link is closed Page 81 of 139

82 14. Attach coupler and insert shear pins for drogue parachute a. Refer to labeling on coupler (This end down, shear pin alignment marks) b. Push coupler into booster airframe c. Rotate as necessary to line up shear pin holes d. Insert three shear pins 15. Pack main parachute and flame retardants Packing procedure to be determined through assembly testing and practice 16. Attach upper airframe to coupler a. Refer to alignment marks b. Insert three screws 17. Insert main parachute into airframe a. Attach quick link to avionics bay, confirm quick link is closed b. Insert wrapped shock cord into airframe, pack down firmly c. Push packed main parachute into airframe, with the protector facing downwards d. If the parachute is too tight, readjust packing to make the package wider or longer as necessary 18. Insert sample parachute into airframe a. Attach quick link to sample container, confirm quick link is closed b. Insert wrapped shock cord into airframe, pack down firmly c. Push packed sample parachute into airframe, with the protector facing downwards d. If the parachute is too tight, readjust packing to make the package wider or longer as necessary e. Attach quick link to nose cone 19. Attach nosecone to upper airframe a. Refer to alignment marks b. Insert four shear pins 20. Insert motor into booster airframe a. Attach the adapter rings (three pieces) b. Insert into motor mount c. Screw on retainer ring, confirming the motor is secure 21. Bring rocket to RSO for safety inspection 22. Make changes as specified by RSO Page 82 of 139

83 Launch: 1. After RSO approval, wait for range clear 2. When range is clear, move rocket to pad 3. Lower launch rod and mount rocket on the rod a. Ensure team members are supporting the weight of the rocket b. Rail button should slide easily along rail. If not, don't apply pressure, rather rotate the rocket 4. Raise rod and rocket to upright position, be sure to support the rocket while lifting 5. One at a time, turn the key switches; listen for continuity, settings check a. Payload altimeter: i. Verify the altimeter turns on b. Payload computer: i. Verify ground station receiving from transmitter c. Primary altimeter: i. 3,1, 10, 10,10 beeps for main deployment altitude ii. Series of beeps for last flight data iii. Series of beeps for battery voltage (Volts, tenths of Volts) iv. Three quick beeps for continuity d. Secondary Altimeter i. 4,9,0,0 beeps for main deployment altitude ii. 5 second siren for apogee delay iii. Series of beeps for last flight data iv. Series of beeps for battery voltage (Volts, tenths of volts) v. Three quick beeps for continuity 6. Check pad power is off and attach igniter to pad controller 7. Insert igniter into motor and plug 8. Leave range and wait for launch 9. Acquire signals from GPS transmitters and camera system before launch Page 83 of 139

84 10. Launch rocket 11. At apogee, wait for separation 12. Wait for rocket to land 13. Upon range clear: retrieve rocket, check for undetonated charges and remove 14. Return to safe area Post Launch: 1. Remove altimeters from coupler and collect data 2. Collect data from liquid sloshing experiment 3. Turn off all avionics and store for transport 4. When travel is finished, clean all dirty components, remove power sources from avionics, and store all materials for future flights Safety Officer The safety officer this year will be Derek Awtry. He is a student studying Aerospace Engineering at the University of Illinois. He worked with the SLI Structures and Recovery team, and as such he has worked on projects similar to this in the past. Preliminary Hazard Analysis The safety officer will ensure that every single member of the team knows the risks associated with their respective sub teams. Each member in the structures and recovery team and the AGSE team shall complete the necessary lab safety training and will be aware of the risks associated with the handling and disposal of hazardous materials. As such, Material Safety Data Sheets (MSDS) will be provided for those who are working with hazardous materials. These MSDS s will also be provided on the team website and will be updated regularly. When construction begins, operation manuals will be provided for all tools and equipment used to build both the rocket and the AGSE parts. These operation manuals will be posted on the website as soon as it is deemed necessary. Personal Protective Equipment (PPE) will be provided to and Page 84 of 139

85 required by team members who are working with these materials or working in a lab with machinery that poses risks to those team members. The Engineering Student Project Lab (ESPL) will deal with larger machinery that the team members do not have the qualifications for. The usage of this machinery requires completion of training courses provided by ESPL. In the event that the safety officer or the team mentor cannot supervise a potentially dangerous situation, the safety officer will ensure that more experienced team members who have worked in these situations before, like the team leader, are able to supervise. Motor storage and transportation will be in accordance with the National Fire Protection Agency, specifically NFPA code The motor shall be stored in a Type 3 or Type 4 indoor magazine because our rocket motor is under 50 lbs. Transportation of the motor will comply with 49 CFR Subchapter C Hazardous Materials Regulation, which covers the packaging, handling, and transportation of High Powered Rocket Motors. The operations manual for the motor will be posted on the website as soon as the motor arrives. The National Association of Rocketry (NAR) code for high power rocketry will be followed at all times. The minimum distance table will also be consulted every time the team launches and tests the rocket. The NAR safety code and the Minimum Distance Table are both posted on the website. Tables of Hazard and Failure mode analyses are given below. Table 7 Hazard Analysis Hazard Likelihood Severity Mitigation Chemical Burns from the rocket motor Moderate to High Moderate Ensure that the team mentor will be working with all components related to the motor, as per regulations, and that the minimum distance table in consulted before all launches and tests of the motor. Burns from black powder usage Moderate to High Moderate Ensure that the team mentor will be on hand to monitor the tests of the black powder. Skin Irritation Moderate Low Ensure that all team members that work with material dangerous enough to induce chemical burns are wearing nitrile gloves and the rest of their skin is covered. Sensitization to Epoxy and Dermatitis Moderate Low Ensure that all team members working with epoxy know not to breathe in fumes from the epoxy directly, especially if the epoxy is highly concentrated. Page 85 of 139

86 Bodily injury from heavy machinery Moderate Moderate Each team member will be required to take a general lab safety course, and team members using tools they have not used before will be trained under the supervision of the safety officer and/or more experienced members. Electric hazard like electric shock Low to Moderate Moderate Ensure that all team members have worked with sensitive electronics before, and that proper grounding procedures are followed Cuts from rocket assembly Electric hazards like burns from battery acid Low Low Ensure that all team members are handling power tools and other tools properly according to their safety manuals. Low Low Ensure old batteries are thrown out, and that new batteries are not faulty. Sufficient testing of the batteries are required Failure Mode Analysis Table 8 Failure Mode Analysis Failure Mode Impact Mitigation Hatch doesn t close Autonomous robot arm does not pick up the sample payload Autonomous robot arm does not get the sample payload into the rocket Rocket doesn t raise autonomously Sample Payload could fall out while the rocket is lifted up or when rocket launches The hatch never closes and then the rocket does not launch The hatch never closes because the sample payload isn t in it The rocket would not be able to launch Make sure the pressure sensor works and that it is connected to the gearbox. Also ensure that the gearbox is functioning in order to close the hatch Proper testing of picking up the sample via the robot arm as a separate test from the entire system. Test the autonomous system until it is determined that there is a very small chance that the robot arm fails Test the autonomous system until it is determined that there is a very small chance that the rocket doesn t raise Page 86 of 139

87 Launch platform doesn t raise autonomously Igniter is not inserted into the rocket Igniter fails to ignite The rocket would have no launch pad to launch off of and the team would have to cut the launch off The rocket would not be able to launch The rocket would not be able to launch Make sure that sufficient tests of the launch platform happen, and that the launch platform never raises without the rocket on it. Ensure proper testing of the igniter system in separate tests from the system as a whole Ensure proper testing of the igniter system in separate tests from the system as a whole, as well as sufficient testing of the igniter system during test launches Motor case fails The rocket would explode Only purchase motor casing from reputable sources, so as to insure no mechanical failure. Motor fails to fully ignite Autonomous systems lose power Payload never gets ejected Rocket never separates The rocket would become unstable and could pose a threat to the surrounding environment. The rocket would not be able to launch due to unmet requirements on the autonomous control side of operations The rocket falls to the ground with the payload still in it The parachutes would not be able to deploy and thus the entire rocket will crash to the ground and the rocket will most likely be destroyed. rendering the rocket not reusable. Only purchase motor casing from reputable sources, so as to insure no mechanical failure. Ensure that all batteries are replaced before the final launch, and that all batteries are tested for functionality before testing. Ensure that the onboard power systems don t lose power so that the altimeter can correctly signal the release of the payload Sufficient testing of the black powder separation system on board the rocket will be needed. All igniters and electric components (including altimeters) will be checked for functionality before a launch. Page 87 of 139

88 Rocket separates early Rocket separates late Parachutes never deploy Parachutes deploy early The flight could take another path that doesn t coincide with our preliminary flight data The parachutes could deploy too late and the rocket could go crashing into the ground and potentially destroy the rocket, rendering the rocket not reusable. The entire rocket will crash to the ground and the rocket will most likely be destroyed, rendering the rocket not reusable. The flight could take another path that doesn t coincide with our preliminary flight data Sufficient testing of the black powder separation system on board the rocket will be needed. All igniters and electric components (including altimeters) will be checked for functionality before a launch. The team will make sure the altimeters are set to the right altitude Sufficient testing of the black powder separation system on board the rocket will be needed. All igniters and electric components (including altimeters) will be checked for functionality before a launch. The team will make sure the altimeters are set to the right altitude As well as testing of the black powder separation system, the parachute deployment system will be tested until it is determined that there is a small chance that the parachutes never deploy As well as testing of the black powder separation system, the parachute deployment system will be tested until it is determined that there is a small chance that the parachutes never deploy early Parachutes deploy late The rocket could go crashing into the ground and potentially destroy the rocket, rendering the rocket not reusable. As well as testing of the black powder separation system, the parachute deployment system will be tested until it is determined that there is a small chance that the parachutes never deploy late. Page 88 of 139

89 On board electronics could malfunction during flight because of aerodynamic forces on the rocket On board electronics could malfunction due to lack of power on board Our altimeters could get damaged and the rocket would not be able to separate. Our altimeters could get damaged and the rocket would not be able to separate. The on board electronics will be secured well enough on the rocket that the aerodynamic forces will have no effect on the performance of the electronics. All the batteries shall be replaced before all launches to ensure that none of the on board electronics are without power. Environmental Concerns There are many environmental concerns to be considered when launching from a field. For instance, on launch day there could be too much wind, causing the launch to be delayed. It could also have rained in the days before the launch, causing the ground to be muddy and soft. The day of the launch could be cloudy and the team could potentially lose the rocket if our on board GPS fails. Some of the ways that the rocket could affect the environment is that the rocket motor could explode or have too much exhaust and potentially damage its surroundings. If the rocket were to explode, debris would be all over the field and potentially harmful if cleanup is not done properly. The rocket could launch and become unstable, thus posing a threat to nearby spectators and occupied areas. The parachutes of the rocket could not deploy and cause damage to the surrounding environment because the rocket then would not be able to achieve a soft landing. IV) AGSE/Payload Criteria Selection, Design, and Verification of AGSE System Review Procedures The sequence of events followed by the AGSE system is listed below as a method of introducing the system definition. Images of the full system are given in Figure 42 and Figure The robotic arm will reach down to the ground where the sample payload is laid 2. The gripper on the robotic arm will clamp the sample payload Page 89 of 139

90 3. The robotic arm will lift the sample payload and bend 180 degrees to the other side where the rocket is placed 4. Once the sample payload is held vertically above the opening in the rocket, the gripper will release the sample 5. The sample payload will fall a short distance into the compartment of the rocket 6. The robotic arm will move to a position behind its blast shield yet out of the way of the rocket. It is important to make sure that the arm does not get into the way of the rocket and rail being lifted 7. The hatch on the rocket for the compartment will close shut 8. The linear actuator for the rail system will be activated 9. Once the rocket is at a vertical, 5-degree angle, the linear actuator will be stopped 10. A mechanical locking mechanism will be triggered to lock the rail in place 11. After the rocket is in the upright position, the linear actuator of the igniter system will be activated 12. The linear actuator of the igniter system will stop at 24 inches 13. The system will be paused so the vehicle s electronics can be armed 14. The safety interlock will be manually switched 15. The rocket is ready to be launched Subsystem Overview The three subsystems of the AGSE consist of the robotic, rail, and ignition systems. These systems work consecutively to place the sample payload in the rocket, raise the rail and then ignite the motor. All of these systems will run off the same computer so that one system cannot start until the previous system has finished. This computer will have a pause switch and at least two different colored LEDs. The three subsystems will also share a 12V battery that will have enough amp hours to run the whole system. Due to the precision required for each subsystem, as well as the precision required of each subsystem interacting with the others, a rigid structure will connect all three subsystems together. Although this structure is rigid, it will not be permanently fixed due to shipping constraints and general movability of the system. This structure will also prevent the system from tipping over because the rail on the ground extends past the combined center of mass of the rail and rocket. Page 90 of 139

91 Figure 42 Rocket on the Lowered Rail Figure 43 Rocket on the Raised Rail Page 91 of 139

92 Robotic System The robotic subsystem is in charge of first picking up the sample payload and then placing it into the rocket. It will accomplish this job by using an assortment of 3D printed pieces and servo motors. This arm will consist of three cylindrical segments with joints connecting each segment, allowing the robot to be able to reach down and pick up the sample, as well as reach over and place the sample into the rocket. Because of the limited movement needed for to pick up the sample and place it in the rocket, the arm will only move in one 2D plane. This eliminates the need for a motor to rotate the arm as well as it makes it more reliable because the position in one direction will be fixed. The first segment of the arm (denoted as segment A) is nine inches long; the second segment (denoted as segment B) is 18 inches long; and the third segment (denoted as segment C), with the gripper fixed to it, is seven inches long. These lengths were determined so that the arm will be able to reach 24 inches down to the ground, as well as be able to place the sample into the rocket while having segment C perpendicular to the ground. These segments will be one inch in diameter and have a hole running through the center for the chords to run through. This design was chosen because the joints of the arm segments will be stronger and easier to create. And because these pieces will be 3D printed, it will allow for the optimization of weight and strength. To utilize this feature, the space between the inner and outer wall will be filled with approximately 80% air with the rest consisting of plastic in a structural efficient honeycomb pattern. The location of these servo motors have also been relocated from where they were in PDR. In the first design, a belt and pulley system was used so that the applied torques could be changed with different ratios of pulley sizes. After further review, with using light, 3D printed segments, it was deemed that the belt and pulley system was not necessary. This design was chosen over the previous due to its lighter weight as well as it was determined that servos could be bought with enough torque to safely raise and lower the arm. The gripper system will also consist of 3D printed pieces, but it will be printed with a 40% infill of plastic so that it is more structurally sound. The gripper system that was chosen utilizes only one servo to operate, which minimizes the chance of failure while still working very effectively. This one servo works by having both fingers geared together so that one cannot move independent of the other. The end of the gripper that will be in contact with the sample will be in the shape of a half pipe so that when the sample is picked up it cannot slip out of its grasp. The half pipe may also have a nonslip coating so that the sample cannot slide out, but that will be implemented on an as needed basis once testing begins. The arm segments along with the gripper have been designed to have all of the center of masses line up in a straight line to try and prevent damage to the servos components, specifically the drive rod. A figure depicting the robotic arm and its gripper system are given below. Page 92 of 139

93 Figure 44 Robotic Arm System Rail System The next subsystem that will be discussed is the rail system. The job of the rail system is to raise the rocket from the horizontal position to five degrees off of vertical. The system described in PDR was going to use high torque, low rpm motor to rotate an axle that was attached to the hinge. This did not offer a large enough mechanical advantage to make it a practical solution. The motor had to impart a large force on the axle to produce the large moment needed to rotate the rail and rocket. Another reason this system was replaced was because a gearbox would be needed in order to get the torque required which would thus increase costs and increase the weight of the system. The system now uses a linear actuator placed 15.6 inches from the base of the launch pad. The other end of the actuator will be connected to the launch rail 22 inches from the pivot point. The pivot point will be a hinge that will be connected to the launch plate. This is farther from the pivot point so it will offer a great mechanical advantage. It will also be a more contained system and be better protected from the blast of the rocket. This system will engage once the hatch door is shut on the vehicle and the sample is safely secured. When it is time to raise the rail the AGSE s computer will send power to the rail s actuator to slowly raise it into position. When it is in the five degrees off of vertical position, a limit switch will be triggered to tell the computer to halt movement of the actuator. Due to the design of the actuator and the friction acting between the gears in the actuator, it will have a natural tendency to hold its position even when power to the system is cut. Even though this setup will allow the rail to be locked into position, a mechanical lock will also be implemented to fix the rail into position. This mechanical lock will consist of a spring, plunger, Page 93 of 139

94 and a mount on the base stand and a hole in a plate attached to the rail. When the rail is raised, this piston will slide along this plate until it reaches the launch position. When that point is reached, a hole in that plate will allow the spring to release its tension and force the piston through the hole effectively locking it in place with the base stand. All of the components of the rail will be fabricated out of either aluminum or steel so that it will not deform due to the stresses of the weights or the temperatures from the blast. Most of these components will be bought with no major modifications needed from companies such as 80/20. What machining is required will be done using the tools available to the team including the aerospace department s machine shop. Calculations were made on this system to determine the force needed to raise the weight of the rail along with the vehicle. Using the script shown in Appendix D, the maximum force needed to lift the rail and vehicle was determined to be 204 lb. The actuator that will be purchased to raise this rail, however, will be capable of exerting a force of 400 lb so that there will be a level of safety as well as it was one of the few actuators that fit the requirements needed for this job. With the subsystems all being connected rigidly, the overall system is more stable due to the distributed loads. There will be runners coming off of the stand to stabilize the rocket and base stand from tipping over due to the wind or other forces acting on it. A figure of the rail actuator is given below. Figure 45 Rail System Linear Actuator Page 94 of 139

95 Igniter System After the rail system finishes its procedures, the igniter system will begin its commands. It will use a linear actuator that is positioned next to the base to raise the igniter into the motor. In between the igniter and the actuator will be a Z piece that reaches under the base and connects the two together. A 24-inch stroke, 35-pound force linear actuator will be purchased for this job. The actuator will be stopped by a built-in limit switch when it reaches 24 inches, which is the distance needed. The rocket will be set 6 inches above the launch pad and the 18-inch igniter must be completely inserted into the motor. The system will begin with the top end of the igniter being right above the launch pad. Therefore, the igniter will need to be raised a little less than 24 inches accounting for the minor thickness of the launch pad. This design has changed since proposal to make the system more reliable and practical. Instead of using a DC motor with a ball and screw rail system, a linear actuator will be used instead. It was assumed before that it would be better if the entire system could fit under the launch pad to protect it from the rocket exhaust. However, after reconsidering the design, the team came to a conclusion that having the system under the launch pad would actually be counterproductive. It would crowd the area underneath the launch pad, which could be dangerous, and it would also be inconvenient when initially setting up the system. Also, fitting a system that raises the igniter approximately 24 inches in a 24-inch high space is difficult. Therefore, the team decided to come up with a mechanism that allows the ignition system to be placed next to the launch pad. Another change from the proposal is that a linear actuator will be purchased instead of building a ball and screw rail system. This ensures accuracy while inserting the igniter. A linear actuator is more accurate in determining how far it lifts the igniter into the rocket motor. It is essential to make sure that the igniter goes deep enough into the motor but not too deep. Another reason why the team decided to use a linear actuator instead of the ball and screw mechanism is because the rocket will be placed at a 5-degree angle. With the linear actuator, the actuator itself can be tilted and aligned to the rocket at approximately 5-degrees. By tilting the actuator, the team can make sure that the igniter ends up in the position it needs to be in. The hole in the rocket motor will also guide the igniter in. With the ball and screw system, the team would not have been able to control the angle as easily as with a linear actuator. The igniter will be attached to a wooden dowel rod by tape. By having a stiff and rigid base, the igniter will easily and accurately be inserted upwards into the rocket motor. The team will purchase a wooden dowel with a ¼ inch diameter. The length will be trimmed to about 25 inches. To account for the thickness of the launch pad, the wooden dowel will be attached to a spring at the bottom. This way, the spring will compress when the wooden dowel is raised to 24 inches completely, not harming the rocket motor. It will also ensure that there is contact between the igniter and the top of the motor better ensuring ignition. A Z shaped aluminum plate will translate the upward force from the linear actuator to the wooden dowel and igniter. This Z shaped plate will allow the linear actuator to be placed next to the launch pad instead of underneath. Being next to the launch pad, the linear actuator will be able to rise 24 inches without anything being in its way. The launch pad will have a small opening in the center for the igniter to be inserted through. The hole will be about ½ inch in diameter to let the wooden dowel and igniter to pass Page 95 of 139

96 through. At the beginning of the launch before the rocket is autonomously raised, the wooden dowel and igniter will already be placed through the hole slightly above the pad. The linear actuator will be protected from the motor exhaust by the Z shaped plate. The plate will be positioned directly in between the rocket and the extended part of the linear actuator. The reason why the Z shaped plate will be built out of aluminum is to serve as a protection from the motor exhaust. Aluminum is a good material because it dissipates heat well, is inexpensive, and is easy to work with. The Z shaped plate not only lifts up the igniter, but also protects the linear actuator to ensure its reusability. A figure of the igniter insertion system is given below, with the insertion system highlighted in orange. \ Figure 46 Motor Igniter Insertion System Page 96 of 139

97 Electronics System There will be one main electronics system that powers the whole AGSE system. This system will run off of one 12V 35AH rechargeable battery that will be located under the robotic arm. For this system, a Raspberry Pi B+ will be used as the computer of this whole system. This board was chosen over other boards such as the Arduino and Beaglebone because of the large amount of GPIO pins it has available. Because 40 GPIO pins are available, it would be able to power all the actuators and servos using relays rather than motor controllers. This reduces the price of the system and also reduces the amount of electrical components. The electronics will have a master kill switch placed between the battery and the computer in order to cut power to the system in case of an emergency. A pause switch will also be connected to the Raspberry Pi so that the official can pause the process in case something goes wrong. For the robotic arm, four servo motors each with a power, ground, and control are connected to the computer. The rail system would then have a linear actuator and limit switch connected to the computer to control that system. Lastly, the igniter system will also have a linear actuator and a limit switch to control its system. There will also be an orange LED used as a safety light and a green led for the all systems go light. The wiring of the electronics can be seen in the graphic below. Figure 47 AGSE Electrical Schematic Page 97 of 139

98 Performance of the subsystems Several calculations were performed when determining the needs of the AGSE. Calculations were performed to find moments, torques, forces, weights, shear displacements, and deformations. These calculations were made to optimize cost and time while utilizing the resources available. Robotic System Performance Characteristics The robotics system had the most calculations of all of the systems. The segments of the robotic system were calculated so that it could reach down and pick up the sample as well as place it in the vehicle. Analysis was also done using MATLAB to find the torque required for each of servos to move their respective segments. Using the code that can be found in Appendix E, the torques of each servo required were found. It was determined that the torque at joint 1 (between base and first arm segment) is 171 oz/in, at joint 2 (between segment A and segment B) is 237 oz/in, and joint 3 (between segments B and C) is 42 oz/in. The torque required for the servo gripper will be determined through testing; however, a servo equal in to size joint 3 s servo will be used to start the trials. The angles that the joints pivot have been determined based on the movements of the robot that minimize the torques required. The angle of joint 1 was determined to rotate 90 degrees, joint 2 will rotate 180 degrees, and joint 3 will also rotate 180 degrees. This will allow the robotic arm to reach all of the positions needed to load the payload into the rocket. The time it takes for the robotic arm to complete all of its tasks will be minimal. The servos chosen for this job did not have data on speed, but the speed of the arm can be estimated by the speeds of similar servos. Using these estimates, the robotic arm should be able to safely load the sample into the vehicle in no more than a minute and a half. Rail System Performance Characteristics A script file was created to determine the force needed to lift the rail and rocket using given data about the geometric angles and lengths involved. Using this file that can be found in Appendix D, it was determined that the force needed to raise the rail and vehicle will be 204 lb. This script also allowed the team to determine the forces needed at any point and showed how it varied with the angle of the rail. This graph can be seen below. Page 98 of 139

99 Figure 48 Relationship between required Actuator force and Rail Angle The speed of this actuator is 0.6 inches per second, so for the actuator to extend to the upright position it will take roughly thirty seconds. Igniter System Performance Characteristics This system has the least amount of calculations needed out of all three systems. This system consists mainly of an actuator whose length was determined by the height of the base off of the ground as well as the depth the igniter needs to go into the motor. The stroke of the actuator needed for this subsystem was determined using this logic to be 24 inches. Because the weights it needs to lift will be very minimal (the weight of the Z piece and the igniter), the force needed to lift them will be very small as well. However, due to the length of the stroke of this actuator, the smallest force of an actuator with that length that could be found was a 30 lb force actuator. The speed of this actuator is the same as the previous one at 0.60 inches per second. This means that it will take the actuator 40 seconds to insert the igniter. This brings the total amount of time needed for the AGSE system to complete its tasks at just over two and a half minutes. Even adding in time for error and the time it spends waiting on the subsystems, it should be easily able to complete the whole process in under ten minutes. Page 99 of 139

100 AGSE Verification Plan and Component Requirements Given below is a table of AGSE system requirements, as well as design features to meet these requirements and verification methods. Table 9 AGSE Requirements and Verification Methods Requirement Design Feature Verification Method Master switch to power all autonomous procedures Capture and placement of sample payload, erection of rocket, insertion of igniter, and activation for launch must be completed in 10 minutes Must have a pause switch to temporarily terminate all actions The pause switch must immediately stop all action when activated Must have an amber or orange safety light indicating the power for AGSE is turned on Safety light must flash at frequency of 1 Hz when power for AGSE is on Safety light must be a solid color when AGSE is paused but power is still supplied All equipment, procedures, and technologies could be implemented in a Martian environment Master kill switch Preprogrammed instructions Pause button Pause button LED light LED light LED light Technologies used The master kill switch will only allow power to the system when closed and will immediately terminate power once opened The instructions for all actions and procedures will be preprogrammed to take place within the 10 minute time limit Pause button will be connected to the computer that controls all AGSE actions and procedures Pause button is directly connected to the computer that controls all AGSE actions and procedures An LED light will be connected directly to the power and computer that controls all AGSE actions and procedures An LED light will be connected to the computer which will cause it to flash when power is on An LED light will be connected to the computer which will cause it to be solid when power is supplied but paused No technologies that could not be implemented in a Martian environment will be used Page 100 of 139

101 All systems will be autonomous Raspberry Pi Raspberry Pi will have preprogrammed instructions to complete all necessary AGSE actions and procedures Sample payload must be picked up Sample payload must be placed in launch vehicle Rocket must be moved from horizontal to launch position Launch platform must be raised to 5 degrees from vertical Launch platform must be pointed away from spectators Igniter must be inserted in rocket Must have an all systems go light to indicate safety verifications are passed and rocket is ready to launch Robotic arm Robotic arm Launch rail, linear actuator, hinge Launch rail, linear actuator, hinge Rail will not rotate to face spectators Linear actuator LED light A pinch grip will pick up the sample payload The robotic arm will have three degrees of freedom in order to pick up and place the sample payload A linear actuator will lift the launch rail to 5 degrees from the vertical There will be a rail limit switch connected to the Raspberry Pi and a lock to ensure placement at 5 degrees from vertical Launch system will be set up so the vehicle points away from spectators at all times The linear actuator will be placed under where the motor will be and will extend to insert the igniter. There will be an igniter limit switch so that the igniter goes the correct distance into the motor A green LED light will be directly connected to the computer and will light up when all actions and procedures have finished and the safety verification is done Rocket must be ignited An Electric Igniter The igniter will be connected to the computer and will be completely inserted in the motor of the rocket Ignitions system must be electric Igniter Voltage will be applied to the igniter causing it to combust Page 101 of 139

102 Must have safety interlock in series with the launch switch that is not installed until rocket is ready for launch The sample payload must be captured and put in the launch vehicle first, then the launch vehicle must be erected, then the igniter must be inserted in the motor A launch button must initiate launch Payload must be ejected at 1,000 feet AGL in descent Launch key Raspberry Pi and setup Launch button Main Parachute Ejection Charge Launching mechanism won't work until key is inserted into control system Raspberry Pi will have preprogrammed instructions to complete tasks in correct order and equipment will be set up to allow the required order The launch button will be used to cause the ignition of the rocket s motor The payload sled will be ejected when the main parachute deploys at 1,000 feet Preliminary integration plan The way the AGSE system was designed requires the integration of all of the subsystems to create a highly precise mechanism. The manufacturing and material choices for these subsystems vary depending on what environment they will be in. The majority of the rail and igniter systems will be made from 80/20 aluminum. The rail itself will consist of two six foot lengths of 80/20 s 1515 t-slotted rail that will be fixed together to make one twelve foot long launch rail. The hinge for this rail will be another one of 80/20 s aluminum products which will allow it to withstand the heat from the blast. The blast plate will be constructed out of a quarter inch thick piece of aluminum with a ten inch diameter so that it will not deform from the blast. The base stand that will support both the rail hinge and the igniter system will be constructed out of various lengths of 1010 t-slotted rail due to the vast amounts of joints and pivots available that can be integrated with this rail. Two aluminum pivots will also be used at each end of the rails actuator to support the weight of the rail and the rocket. A 36 inch by 5 inch piece of quarter inch thick aluminum plate will be bent into the Z igniter piece by the use of a hydraulic or arbour press. The mount for the igniter system s actuator will be constructed by drilling a hole into a piece of 1010 rail and allowing the bottom of the actuator to be securely fastened to it with a bolt. The spring on the other end of the Z piece will be made out of sprung steel, which will be able to handle the harsh environment that it will experience when the rocket launches. There will also be two pieces of aluminum plate that will be attached to the moving and nonmoving parts of the rail system s actuator. These two plates will be able to slide past each other and they will shield the actuator from the blast of the motor. Page 102 of 139

103 Attached to the base will be segments of the 1010 rail sticking out in all four directions to help prevent the rail from tipping over. The 1010 t-slotted rail will also be used to secure the base to the robotic arm s stand. This will consist of two pieces of rail running the roughly six feet from the base to the arm s stand to secure it in position. The robotic arm s stand will also be constructed out of t-slotted rail to create a structurally sound stand that will not bend or move. There will also be two plates below the arm that will be able to hold all of the AGSE s electronic components. All of these 80/20 and other aluminum parts will be secured together using brackets and mounts that 80/20 produces to fit with its products. This will reduce cost and the price of the components because there will be no need for custom machinery. All of the nuts and bolts for these pieces will be made of either steel or aluminum so that they can withstand the stresses involved. The manufacturing and materials involved in robotic system are completely different from the other two systems. The entirety of the robotic arm s structure will be constructed out of 3D printed PLA plastic. This allows the arm to be lightweight yet strong and in almost any shape that can be imagined. The segments of the arm will be one inch in diameter cylinders, with a honeycomb inner structure to reduce weight while not compromising strength. These segments will also have a hole running down the length of them to allow the wires to be protected and hidden. These components will be printed with a 20% infill of plastic, while the gripper system will be printed with 40% infill. The gripper system will be printed with this higher infill due to the smaller parts involved as well as it has less structurally efficient shapes. The arm s segments are designed to overlap at the joint, the points of rotation. This will allow for the body of the servo to be fixed to one segment, while the rotational head of the servo to be fixed to the other segment. This ensures rotation occurs exactly at the joints between the segments. The robotic system will be attached to the aluminum stand using steel nuts and bolts. Precision of instruments and repeatability Due to the design of the system and the components used, it will be effectively precise and repeatable. The robotic arm consists of four servos to move the joints and gripper that all have feedback for very precise positioning. The servos also allow for the robotic arm to be very repeatable because the computer can determine what angle each segment is at compared to the previous segment. Using this robotic arm system, the only loss to precision is when the sample payload is placed into the rocket, and that is due to the friction acting between the gripper and the sample payload. The rail system that is used will be just as precise, if not more precise, than the robotic arm system. The rail system will use a linear actuator system that will stop moving the rail as soon as it reaches the five degree off of vertical point. It will stop at this point by hitting a limit switch that stops the actuator from extending any farther. When it gets to this point, a spring loaded pin will also fall into place locking the rail in position and ensuring it is at the same angle every time. The precision of this rail being in launch position will depend on how accurate the angle can be measured with the equipment available to create this system. That is because the positioning of the limit switch that tells the actuator to stop pushing the rail up is what will determine the accuracy of that angle. The limit switch will be able to be positioned so that the rail will be as precise as angle as can be measured. Page 103 of 139

104 Out of all three systems, the igniter will be the least precise. That being said, it will be more than precise enough to ignite the motor every time. The igniter system will be very precise as well because it also uses a limit switch to stop the actuator when the igniter is in position. The reason this system may be slightly less precise is because precision was traded for repeatability. This was done by attaching a compressible piston between the igniter and the Z piece. This piston will ensure that the tip of the igniter is always in contact with the top of the motor to better ensure ignition. The piston will also be very easily compressible as to not damage the igniter or motor when it reaches the top end of the motor. Electrical Schematics The electrical components are considered in the AGSE subsystem description section of this report beginning on page97, and an electrical schematic is given above in Figure 47. Key Components of the AGSE System The team is using servo motors to actuate the robotic arm. Three will be used to provide motion to the segments. The motors will be attached to the joints. The spinning motion from the servos will cause the segments to rotate. Servos use a small DC motor to create their movement. The servos that are connected to the joint between the blast plate and the first arm, and the joint between the first and second arm segments will have a maximum torque of 333 oz.in and use 6V of electricity. The servos in the joint between the second and third arm segments and the servo in the gripper will have a maximum torque of 104 oz.in and use 6V. The servos will be connected to a relay which will in turn be connected to the computer. The servos will allow the arm to change positions and hold onto the sample and place it into the rocket. To lift the rocket, the team will use a linear actuator. It will be attached to the launch rail at 1.3 feet from the launch pad. It will have an unstretched length of 22.3 inches and will extend and push the rocket to a position of 5 degrees from vertical. The actuator works by extending a rod from the main housing of the actuator. The rail actuator will have a maximum force of 400 lbs. The rail actuator will also work with the locking mechanism to hold the rocket in place for the igniter to be placed in the motor and to launch the rocket. The actuator for the ignition system works just like the one for the rail system. The actuator for the igniter will be tilted at 5 degree from vertical to ensure the igniter will be inserted into the rocket easily and without damage to the igniter or motor. This actuator will have a stroke length of 24 inches. All these systems will be connected to a Raspberry Pi for control and a 12V lead acid battery for power. AGSE/Payload Concept Features and Definition Creativity, significance, and level of challenge The team decided to build the robotic arm from scratch instead of purchasing a pre-made one. To lift the sample payload off the ground and into the rocket compartment, the robotic arm only needs three degrees of motion and does not need to be too complex. Most of the robotic arms on the market were too expensive and had more capabilities than needed for this project. To Page 104 of 139

105 minimize cost, the team will build its own robotic arm. This will add to the creativity and level of challenge for our project. Most of the components of the robotic arm will be 3D printed. Even though the robotic arm will be reusable after a rocket launch, by 3D printing components, the robotic arm could be easily replicated and assembled. Through this process, the team will also be able to learn about CAD, 3D printing, and the mechanisms of a robotic arm. The team will also circuit and program the robotic arm to have it transport the sample payload accurately and safely. The AGSE success will depend on the accuracy of the payload pickup and placement. The payload has to be picked up neatly from the pre-determined place and dropped precisely into the opening in the rocket in a specific time period without damaging the rocket. The perfect positioning of the robotic arm-gripper system and the payload will be key. This AGSE design is very original and unique due to the fact that all systems were designed by and manufactured by the team members. The only exception to this is the gripper system, which was based off of another gripper but will be modified to accomplish the task required and it will be completely constructed by the team. However, because everything was designed and will be made by the team, there is a suitable level of challenge involved, however the whole team is extremely confident that it can be accomplished. Science Value AGSE Objectives The AGSE system will be designed to perform autonomously as an analog to a full Martian sample return vehicle. The system should demonstrate the full set of capabilities that would be necessary to retrieve a sample from a rover or other piece of equipment and then launch it into orbit around Mars. While the rocket used in this system will not be fully to scale with the necessary final Martian launch system, the concepts demonstrated by the team s design will be sufficient to give a proof of concept for the vehicle. AGSE Success Criteria A successful run of the system will see the sample payload lifted from a predetermined position and placed safety into the hatch. Following this, the rocket will autonomously raise to a position 5 degrees off of vertical. The igniter will be positioned fully into the motor of the rocket and then the system will await launching approval from the range safety officer. On the way down, the rocket will release the payload at 1,000 feet AGL, demonstrating the capability of the return vehicle to safely deliver the payload to the Earth s surface. AGSE Approach The team s AGSE system was defined by first identifying the key requirements of the system. Once these requirements were fully outlined, a set of options for all needed components (sample recovery, launch rail erector, and igniter lifting) was constructed. For each of these subsystems trade studies were performed using the possible design options to identify the system that best balanced safety, robust operation, cost, and viability. From these trade studies, the best options were then selected and optimized in order to create the optimal configuration for the total Page 105 of 139

106 AGSE system. Special care was paid to the integration of all components to ensure there were no problems with incompatible components or operations. As the design process is maturing, more simulation is taking place of all operations, especially for the rail system and robotic arm, to confirm that all chosen components will be sufficient to meet the necessary design constraints. As the design process shifts into building and testing, care will be taken to assure that all components are constructed as designed. The team members taking part in the build process for a component will be the same team members who designed said components, or will be supervised closely by members intimately familiar with those systems. AGSE Testing Testing of the AGSE system will take place throughout the building process as individual components are completed to ensure all subsystems are fully operational independently and to avoid any last minute changes or unexpected problems. Testing of the integration of these components, such as the testing required to assure that the system does not have any unanticipated physical barriers to operation, will be a crucial part of the process and will take place as soon as all relevant components are completed for any required operation. Any failure during testing will be thoroughly investigated by the team and solutions to any failure will be closely monitored to ensure they do not interfere with the operation of other unchanged, existing systems. In the testing cases, worst case scenarios with regards to loading and operational hazards will always be assumed to certify the system in the most extreme expected environment. While this operation will provide a reasonable test bed for a Martian Sample Return mission, the conditions at the launch will obviously not be identical to those found on Mars, making this an imperfect test. However, no components will be used that would not function similarly in the Martian environment. No systems requiring Earth s gravity, magnetic field, or atmospheric conditions will be utilized by the team in the completed system. In some cases (such as electronics), Earth based analogs will be used as the radiation and thermal protective measures required on Mars would lead to excessive cost for the system. These components could easily be replaced for a Mars mission to give operational configuration. Successful operation of the sample payload placement, rail lifting, and igniter inserting system will be verified visually by team members with experience as to the correct final configuration of a high powered rocket before launch during the testing phase. On the day of launch, the Range Safety Officer will also confirm the AGSE has met its full set of requirements and is safe to launch. The recovery of the payload separate from the body of the rocket will signal the completed successful operation of this system, with the release height of 1,000 feet AGL being confirmed by the flight altimeters after recovery. AGSE Progress Currently, the team has finalized a baseline design for the AGSE system and has completed extensive simulations of the system to confirm that it will meet all competition requirements and not encounter any physical limitations. Detailed models of the parts have been created and simulations of every robotic component s operation have been performed. These simulations have modelled the dynamics of the robotic systems to ensure all motors meet the minimal loading Page 106 of 139

107 requirements with a significant margin included in all final conclusions. In addition, some test pieces for the robotic gripper have been produced by the team using additive manufacturing techniques to confirm these parts meet the structural strength requirements for the AGSE system. Moving forward, the team will continue to construct individual components of the AGSE system, testing each for strength as they are produced. Integration of these components will occur immediately as subsystems and finally full systems are completed. V) Project Plan Status of activities and schedule Budget Plan Given below are the components the team anticipates using for the project, as well as their quantity and associated costs. Items marked with an asterisk (*) are already owned by the team and will not need to be purchased. Table 10 Project Components and Cost Analysis Component Rocket Components Cost (each) [USD] Quantity Total Cost [USD] Total cost to Team [USD] 48 Long 5.5 diameter Blue Tube " PNC Nosecone (long) x 4 Aircraft Plywood Aerotech L1150 Motor * (1 owned) RMS 75/3840 Motor Case * mm Blue Tube Aeropack 98mm Motor Retainer Body Aeropack 98mm Motor Retainer Cap 28 1* Coupler Bulkheads Centering Rings 1515 Rail Buttons (pack of two) (pack of two) Skyangle 36 Drogue Parachute * Page 107 of 139

108 Iris Ultra 96 Main Parachute 387 1* Iris Ultra 60 Sample Parachute Parachute Deployment Bag /2" Forged steel eyebolts /2" Tubular Kevlar Shock Cord 1.15 per foot 50 feet * (40 owned) /2" nuts /2" washers 8.98 for " Blue Tube coupler /4" Threaded Aluminum Rods 7.09 per 6 feet Stratologger Altimeters * (2 owned) Telemetrum Altimeters Rotary Switches Volt Batteries mah LiPo Battery /4" Quick Link * (4 owned) /2" PVC end caps /4" nuts T-Nuts 6.23 for /4" washers 3.93 for Servo Motors (hatch) Gears (hatch) for set Gear Racks (hatch) for pack of Hatch power supply D shafts (hatch) Flex Sensors (hatch) Microcontroller (hatch) Brackets (hatch) Proline 4100 Epoxy per gallon 1* Nitrile Gloves for Page 108 of 139

109 2 Position Barrier Strips Subscale Rocket Materials Subscale Body Tube Subscale Motor Subscale Fins (3d Printed) Subscale Nose Cone Subscale Parachute * Subscale Altimeter * Subscale Motor Case * Subscale bulkhead & hardware Subscale Rails Buttons 3.07 for Subscale Coupler 4.75 for Subscale Retainer Subscale Centering Rings AGSE Materials Linear Actuator for Igniter System 1515 T-slotted Aluminum Launch Rail lb force/ 18" Stroke Linear Actuator Raspberry Pi B V 40A Relays Servo for Segment A Servo for B Servo for C Gripper Servo V Lead Acid Battery V Battery Charger Red Stranded Wire 100' Black Stranded Wire 100' Green stranded wire 100' Page 109 of 139

110 Wire Connectors Servo Extension Green LED Led Safety Light Kill Switch Pause Switch Limit Switch Blast Plate Hinge for Launch Rail Structure per inch Actuator Shield Actuator Pivot Nub (Rail) Actuator Mount (igniter) Z Piece for Igniter System D Printed Arm Segments * D Printed Gripper Pieces * Educational Outreach Viking Model Rockets for A8-3 Rocket Engines for Estes Portable Launch Pad Travel and Accommodations Total Final System on the Launch Pad Project Totals: Page 110 of 139

111 Funding Plan The Illinois Space Society has secured funding for all costs expected to be incurred throughout this project. Several organizations at the University of Illinois provide funding to student groups to support meaningful student engagements such as this competition. These organizations are the Engineering Council and the Design Council. Additionally, all University of Illinois students pay an Organizational Resource Fee, which is distributed to projects such as Student launch. The team is also contacting additional corporate sponsors for the purposes of securing more funding. Given the fact that current funding sources outweigh projected costs, this surplus will allow more team members to attend launch activities with minimal costs to students and will also allow the team to purchase further educational outreach supplies. A summary of funding sources is given below. Table 11 Funding Sources Funding Source Amount [USD] Illinois Engineering council 1,500 Illinois Design Council 2,000 Student Organization Resource Fee 3,500 Corporate Sponsors 1,975 Total Funding 8,975 The available funding surplus as well as additional sponsorships will allow the Illinois Space Society to manage any small increase in costs. Any budget overruns or replacements parts that need to be purchased will not eliminate the Illinois Space Society s ability to compete in this event. Timeline The current project timeline is listed below, and includes all project milestones as well as self-imposed construction and testing deadlines. The team intends to add significant detail to this timeline in the future, however these dates are highly dependent on how quickly parts may be obtained and cannot yet be determined. The Engineering Open House outreach event takes place on a fixed schedule, however additional educational events are still being scheduled with schools and educators. Page 111 of 139

112 Table 12 Project Timeline Milestone Proposal Due Selection Notification Team Web Presence Established Vehicle and AGSE Design Definition Complete PDR Report, Slides and Flysheet PDR Presentation Subscale Construction Complete Subscale Test Flight Completed Igniter Insertion System Completed and Tested Rail System Complete and Tested CDR Report, Slides and Flysheet CDR Presentation Vehicle Construction Complete Robotic Sample Insertion Arm Complete and Tested Recovery System Ejection Testing Full Scale Test Flight Complete Open House Educational Outreach FRR Report, Slides and Flysheet FRR Presentation Travel to Huntsville Launch Readiness Reviews LRR and Safety Briefing Rocket Fair and MSFC Tours Completion Date October 6th October 17th October 31st October 31st November 5th November 7th-21st December 5th December 20th December 20th January 14th January 16th January 21st - February 4th February 20th February 20th February 21st February 28th March 13th-14th March 16th March 18th-27th April 7th April 7th April 8th April 9th Page 112 of 139

113 Launch Day, Banquet Backup Launch Day Post Launch Assessment Review Winning Team Announced April 10th April 12th April 29th May 11th Gantt Chart A Gantt chart of the project plan is given below. This lists all planned events as well as their current completion percentages. The current time period is bordered by a red outline. Time periods marked in dark purple represent activities that have been completed. Items marked in light purple are planned in order to meet competition requirements. Items marked in light orange are planned in efforts to exceed project requirements and further refine the presentation of this project. Once these additional events are completed, they will be marked in dark orange. With regards to the critical path, many activities are of high importance. Due to the strictly structured timelines of this competition, the vast majority of all deadlines are fixed. Additionally, each design stage of the project must be complete before moving on to the next stage. For example, CDR must be completed before FRR is begun. Due to this cascading effect of the design reports, as well as the significant time and effort required to complete these reports, it is highly critical that the design of the vehicle continue without delay. The final construction of the launch vehicle and AGSE systems are similarly constrained. These have absolute deadlines which must be met to fly in the competition, and must be initiated as soon as possible. Many of the project sub goals (such as individual subsystem constructions) are somewhat flexible in terms of timeline. Although all systems must be complete before the full scale test flight, self-imposed deadlines have been set based on realistic estimations of how quickly these systems may be constructed. These serve as project goals, but not hard requirements. For example the team intends to construct the motor igniter system before the rail lifting system, however this is not a necessary condition for project success. Educational outreach events are also a relatively flexible area in the project plan. The team intends to meet most of the requirements through the Engineering Open House event. However this event occurs extremely close to the outreach deadline. Due to this fact, the team intends to meet requirements before the Open House event, and continue work at the Open House to exceed project requirements. Although delay in outreach activities will increase the risk of not meeting project requirements, delays would not necessarily exclude the team from accomplishing the desired goals. However such a drastic delay in these activities is of course not planned or recommended. Component testing is also an activity with some degree of flexibility. Assuming all tests are completed before the full scale flight test, the timeline of testing activities is not directly relevant to the project completion timeline. However this assumes positive verification of system characteristics through testing. In the event that testing proves a system to be insufficiently designed, redesign and reconstruction will be required. Due to this, it is important that testing be Page 113 of 139

114 completed as soon as possible, even though hard deadlines for test completion are relatively late in the project. In summary it is highly important that the design and analysis process continue at a relatively brisk pace throughout the project. Many intermediate goals and construction steps do not have rigid deadlines, it is extremely highly recommended that all construction and testing activities are carried out as soon as possible. Figure 49 Gantt Chart of Project Plan Page 114 of 139

115 Educational Engagement Plan Engagement Goals The team s plan of actively engaging as many students as possible through hands on demonstrations and activities remains unchanged. The purpose of all engagement activities will not only be to teach the community about physics and rocketry, but also to inspire support and participation in the future of spaceflight technologies. Due to the complicated nature of rocketry principles, engagement activities will focus on hands on demonstrations of the basic physics behind rocket flight. The nature of this mission also allows the team to educate the community on another interesting technological area, specifically robotics. The high relevance of robotics systems to this particular mission allow the ISS team to explore educational areas previously not touched by past outreach events. These activities are planned to occur in a distributed manner throughout the project, and as such the outcome of activities will be evaluated in order to improve future events. The main focus of this feedback will be determining the interest level of those involved, and the understanding of principles demonstrated by the team. This will allow the team to adjust presentations for future activities in order to better educate the community. An initial draft of this feedback form is given in Appendix A. Through the team website the team will also implement a contact system wherein participants of outreach events may request further information or demonstrations from the team. Engagement Opportunities The Illinois Space Society and the College of Engineering provide many opportunities for the team to plan engagement events. Particularly, the Illinois Space Society features an Educational Outreach team which has established relationships with many local schools. This offers a convenient starting point for engagement activities. Particularly, the team intends on contacting schools in Mahomet, Urbana and Champaign Illinois to offer educational services to students. Additionally the team intends on offering hands on demonstrations to students at the University High School on campus and the High Schools previously attended by team members. This allows students to both give back to the local community and the institutions that have previously educated the team. These activities typically take the form of optional after school classes for students, or interaction with school science clubs. The team is currently engaged in the process of contacting these school and determining the interest and availability of instructors. Additionally, the ISS Student Launch Team has contact with local Boy Scout groups through previous engagements, and the team plans on capitalizing on these opportunities for additional engagement. Another major opportunity for engagement is the College of Engineering s Open House on March 13th and 14th. As this is only several days before the educational engagement deadline, the team will strive to complete the required engagement activities before this time. Nevertheless the team still intends to participate in the Engineering Open House. This is a large event held every year and attended by thousands of students and community members. Although not all of these attendees can be directly engaged by the ISS Student Launch Team, the Open House still provides an important opportunity to interact with the community. The team plans on operating continuous activities in order to facilitate indirect interactions with the community. However the team will also use this event to provide direct interactions with students and educators. In order to do this, the team hold scheduled demonstrations at advertised times in order to allow structured hands-on demonstrations. Page 115 of 139

116 The team also has direct contact with a middle school in Moline, Illinois through a family relationship between a team member and an educator. This school is roughly 3 hours away from the University of Illinois campus, which may make direct education events difficult or impossible. It may be necessary to have the educator facilitate hands on activities in the classroom while the team communicates with the students via a video chat. However this may not allow the team to meet the competition outreach requirements. Regardless, the team intends to pursue this engagement activity whether or not it satisfies Student Launch outreach requirements, as it is believed to be a valuable opportunity for both team members and students. Engagement Progress and Plans The team believes the best way to engage students and create an interest in rocketry is to allow the students to build and fly their own rockets. To meet this goal, the team will purchase about 100 Estes Viking model rockets than can be assembled and flown by students. These rockets are designed to allow multiple custom fin configurations, which will allow the team to discuss some basic aerodynamic principles. Although these model rockets will certainly entertain students and engage them in rocketry, the team believes more is required from an educational aspect. The ISS team is researching more hands on approaches to describing physical principles to the students. The main plan is for the ISS team to explain the basics behind Newton s laws of motion, as these are at the forefront of the understanding of basic physics. The team intends to devise a set of simple experiments that are characteristic of one of the three basic laws. These will allow the students to directly interact with the physical principles, and see for themselves how forces affect the motion of objects. By combining the design and building process of the model rockets, and the theoretical demonstrations, the team believes it can provide students with a glance into the full scope of the engineering process. VI) Conclusion The Illinois Space Society is highly committed to the future of rocketry both from an industrial and a hobbyist standpoint. The team is proud to once again compete in the Student Launch competition and intends to do so for the foreseeable future. Team members from many different majors and departments throughout the University have already dedicated several hundred engineer-hours to the design and documentation of these systems. Most importantly, over half of these team members are first or second year students with little to no previous rocketry experience. Under the guidance of the team mentor and more experienced team members, these new team members have already gained a significant amount of valuable insight into both high power rocketry and, more importantly, the real world processes of design and engineering. This personal growth will only be magnified once the construction of the systems themselves is initiated. In previous years, the ISS team has treated this competition as an extracurricular activity for students. Although this is technically still the case, the team intends to put forth a significantly Page 116 of 139

117 higher degree of effort and a more highly defined design than in previous years. Whether it is through writing custom simulation code, presenting hand calculations, or a higher degree of detail in models and drawings, the team has and intends to continue to work put forth significant effort and treat this competition with the attention it deserves. Student Launch provides an opportunity for students to participate in a high profile, real world engineering experience with many technical challenges. This type of opportunity is not common in the world of undergraduate education, and the ISS team does not take this experience for granted. Page 117 of 139

118 Appendix A Illinois Space Society Educational Feedback Form How interesting was the demonstration? (circle one) Not Interesting.A Little Interesting.Very Interesting.Super Interesting How much did you learn from the demonstration? (circle one) Nothing.A Little.A Lot What did you learn from the presentation? What was your favorite part about the demonstration? What was your least favorite part? Do you have any questions to ask the team? Page 118 of 139

119 Appendix B: ISS Tech Team Safety Policy Page 119 of 139

120 Page 120 of 139