Project WALL-Eagle Maxi-Mav Critical Design Review

Transcription

1 S A M U E L G I N N C O L L E G E O F E N G I N E E R I N G Auburn University Project WALL-Eagle Maxi-Mav Critical Design Review 2 Engineering Dr. Auburn, AL January 6th, 205

2 Table of Contents SECTION : OVERVIEW... 4 SECTION.: TEAM INFORMATION... 4 SECTION.2: LAUNCH VEHICLE SUMMARY... 4 SECTION.3: AGSE SUMMARY... 5 SECTION.4: CHANGES MADE TO AGSE/PAYLOAD CRITERIA... 6 SECTION.5: CHANGES TO VEHICLE CRITERIA... 7 SECTION.6: CHANGES TO PROJECT MANAGEMENT... 7 VEHICLE CRITERIA-AIRFRAME DESIGN... 8 SECTION 2.: MISSION STATEMENT... 8 SECTION 2.2: BODY TUBES... 8 Section 2.2.: Payload Body Tube... 9 Section 2.2.2: Main Body Tube... 0 Section 2.2.3: Engine Tube... 0 Section 2.2.4: Construction Methods... SECTION 2.3: NOSE CONE... 2 Section 2.3.: Size and Shape Selection... 2 Section 2.3.2: Manufacturing Method... 4 SECTION 2.4: FINS... 4 Section 2.4.: Shape Selection... 5 Section 2.4.2: Construction Methods (Bulkplates Included)... 6 Section 2.4.3: Size... 8 SECTION 2.5: MOTOR... 8 Section 2.5.: Loki Research K960-P... 8 Section 2.5.2: Cesaroni L SECTION 2.6: MASS ESTIMATES SECTION 2.7: MANUFACTURING Section 2.7.: Tests Section 2.7.: Quality Control Section 2.7.2: Progress SECTION 2.8: VERIFICATION CRITERIA SECTION 2.9: SUBSCALE FLIGHT RESULTS Section 2.9.: Subscale Model Section 2.9.2: Flight Results Section 2.: Influence on Design RECOVERY...38 SECTION 3.: RECOVERY SYSTEM OUTLINE SECTION 3.2: SUBSCALE ANALYSIS SECTION 3.3: REQUIREMENT VALIDATION... 4 SECTION 3.4: PARACHUTES SECTION 3.5: EJECTION SYSTEM... 45

3 SECTION 3.6: ELECTRONICS SECTION 3.7: ATTACHMENT HARDWARE SECTION 3.8: MANUFACTURING SECTION 3.9: SUBSYSTEM TESTING AGSE CRITERIA...60 SECTION 4.: TESTING AND DESIGN OF AGSE EQUIPMENT Section 4..: System Level Design Review Section 4..2: AGSE Structural Design and Materials... 6 Section 4..3: Payload Retrieval Subsystem (PRS) Section 4..4: Launch Vehicle Elevation Subsystem (LVE) Section 4..5: Automated Charge Insertion System (ACI) SECTION 4.2: AGSE ELECTRONICS Section 4.2.: Launch Controller and Function Section 4.2.2: Master Microcontroller Section 4.2.3: Robotic Arm Location and Integration Algorithm SECTION 4.3: AGSE/PAYLOAD INTEGRATION... 9 Section 4.3.: Integration Plan... 9 Section 4.3.2: Compatibility of Elements and Simplicity of Integration Procedures SECTION 4.4: FAILURE MODES AND MITIGATION PLANS SECTION 4.5: PLANS FOR MANUFACTURING AND ASSEMBLY SECTION 4.6: PRECISION OF INSTRUMENTATION, REPEATABILITY OF MEASUREMENT SECTION 4.7: AGSE CONCEPT FEATURES AND DEFINITION Section 4.7.: Creativity and Originality Section 4.7.2: Significance of Design Section 4.7.3: Level of Challenge SECTION 4.8: SCIENTIFIC VALUE OF DESIGN Section 4.8.: AGSE Objectives Section 4.8.2: AGSE Success Criteria... 0 Section 4.8.3: AGSE Requirement Testing Verification Table... 0 Section 4.8.4: System Tests, Variables, and Controls Section 4.8.5: Experiment Process Procedures SAFETY...09 SECTION 5.: CHECKLISTS Section 5..: Final Assembly Checklist Section 5..2: Launch Procedures Checklist... 2 SECTION 5.2: SAFETY OFFICER... 5 SECTION 5.3: HAZARD ANALYSIS... 5 Section 5.3.: Airframe... 5 Section 5.3.2: AGSE Section 5.3.3: Recovery Section 5.3.4: Outreach SECTION 5.4: PRELIMINARY ENVIRONMENTAL EFFECTS Section 5.4.: Vehicle Effects on Environment

4 Section 5.4.2: Environmental Effects on the Vehicle SECTION 5.5: UPDATED ENVIRONMENTS EFFECTS Section 5.5.: AGSE Section 5.5.2: Recovery Section 5.5.3: Airframe Section 5.5.4: Outreach PROJECT PLAN...49 SECTION 6.: TIMELINE SECTION 6.2: CRITICAL PATH SECTION 6.3: BUDGET PLAN SECTION 6.4: FUNDING PLAN SECTION 7: EDUCATIONAL ENGAGEMENT...62 SECTION 7.: GENERAL MISSION STATEMENT SECTION 7.2: DRAKE MIDDLE SCHOOL 7 TH GRADE ROCKET WEEK SECTION 7.3: ROCKET WEEK PLAN OF ACTION SECTION 7.4: ROCKET WEEK LAUNCH DAY SECTION 7.5: ROCKET WEEK LEARNING OBJECTIVES SECTION 7.6: GAUGING SUCCESS SECTION 7.7: AUBURN JUNIOR HIGH 9 TH GRADE ROCKET TEAM SECTION 7.8: TUSKEGEE AIRMEN NATIONAL HISTORIC SITE FIELD TRIP SECTION 7.9: SAMUEL GINN COLLEGE OF ENGINEERING E-DAY SECTION 7.0: AURA MOVIE NIGHT EVENT SECTION 7.: BOY SCOUTS MERIT BADGE UNIVERSITY SECTION 8: CONCLUSION

5 Section : Overview Section.: Team Information Organization Team Name Address Auburn University Project WALL-Eagle 2 Aerospace Engineering Building Auburn, Alabama Mentor Contact Information Name Title TRA # & Level Address Dr. Eldon D. Triggs Auburn University Lecturer and Laboratory Manager Department of Aerospace Engineering trigged@auburn.edu Level 2 Certified, TRA # #259 2 Aerospace Engineering Building Auburn, Alabama Section.2: Launch Vehicle Summary Parameter Size Weight Motor Recovery Rail Size Value 85 inches 25.8 pounds Loki-K960-P CO2 Ejection 5-5 Rail 4

6 Section.3: AGSE Summary The AGSE is titled WALL-Eagle along with the rest of the rocket. First, in order to capture the payload, a retrofitted CrustCrawler robotic arm scans the ground utilizing IR sensors in order to detect the payload. Once found, the robotic arm will retrieve the payload, and place it into the payload bay inside the rocket. Once complete, the arm will then close the outer door to the payload bay. Next, the erection of the launch vehicle will take place utilizing a winch and pulley system to lift the rocket into place. Once complete and locked, after a pause the igniter will be inserted utilizing a telescoping dowel which will be placed underneath the blast plate of the Launchpad. Once confirmed to be safe, the altimeters will be armed, and then the rocket will be launched. 5

7 Section.4: Changes made to AGSE/Payload criteria Payload Bay Door Closing Mechanism Changes were made to the payload bay door and the mechanism that closed and secured the door. Due to testing a decision was made to switch from a microservo mounted in the cavity of the payload bay just below the nose cone, to a torsional spring hinge system held open by a mechanical trip wire. The mechanical trip wire mount will be fastened to the truss 2 inches to the right of the body of the rocket. A pin in the trip wire will hold one end of a nylon strand. The nylon strand will loop through a whole in the door of the payload bay and reconnected to the base of the mechanical trip wire mount. After the payload is secured within the payload bay, a relay will initiate a solenoid that will pull the pin and allow the door to slam shut with sufficient force to secure the door closed. The mechanical trip wire system was found to be more reliable that the microservo. The mechanical trip wire could be armed with little effort and provided enough force to secure the door closed. The microservo would not secure the door effectively and repeatedly, and provided to be a challenge to time within the larger system. Payload integration and Storage Changes to the securing method of the payload include, use of polyurethane spray foam within the cavity of the airframe instead of a mechanical claw system. The polyurethane spray foam proved to be a reliable and low maintenance method to secure and protect the payload. The mechanical claw, while optimal design, would provide unneeded moving parts as well as margin for mission failure. The polyurethane spray foam will fill the entirety of the payload section. A section will be cut out of the foam so that the walls of the foam slope down and intersect at the center the airframe. The foam is cut this way so that the payload, when placed in the payload bay, will roll and rest of the center of the cylinder. Foam strips will be adhered to the door of the payload bay as to prevent any movement of the payload within the cavity of the rocket while in flight. Launch Rail Erection System Changes to the launch rail erection system include, use of a tower on one side of the larger box to winch the trust into launch position. A DC stepper motor was initially designed to provide the means to erect the truss to the desired position, but this designed proved to be complicated in 6

8 implementation. A tower with a wench will proved a reliable, cheap, and easier to design. The wench would also provide less stress to the overall structure but will provide substantial bulk to the structure. Section.5: Changes to Vehicle Criteria Section lengths were changed in order further accommodate the payload and ballast bays, as the original length in PDR was not sufficient to fit the length of parachutes. An alternative primary motor was selected, as it was highly comparable, and even easier to source. The motor section was redesigned in order to accommodate the larger motor. Attachment points inside the rocket were changed accordingly in order to accommodate the new dimensions and shifting of bays. Manufacturing methods were updated. Fin shape was adjusted in order to compensate for the new length and provide higher stability. Added design section for a switch collar in order to make the altimeter bay more accessible. Payload bay fully designed. Section.6: Changes to Project Management Budget was updated to much better reflect status in the current project. Funding was updated to reflect new sources. 7

9 Vehicle Criteria-Airframe Design Section 2.: Mission Statement In order to successfully complete Project WALL-Eagle s mission, the robotic arm must recover the payload, place it within the specified payload bay. And must fly to an altitude of 3000 feet, after successful completion of the AGSE mission, deploy a drogue parachute at apogee to slow the rocket to a safe main deploy velocity, eject the payload bay and main parachute at 000 feet, finally recovering at a safe kinetic energy to the ground. A rendering of the rocket is in Figure 2.. Figure 2.: WALL-Eagle Rocket Section 2.2: Body Tubes Body tubes are essential to effective performance of the vehicle. Since the body tubes comprise the largest surface exposed to the airflow, the aerodynamic properties of the body tubes are highly relevant to the altitude gained by the vehicle. Additionally, as the largest structure in the rocket, the body tubes represent the largest collection of mass in the rocket. With these design parameters in mind, it is critical to select and design body tubes that can survive the stresses of high-powered flight while still remaining light enough to achieve the mission altitude. The body tubes and their lengths are shown in Figure

10 Section 2.2.: Payload Body Tube Figure 2.2: Body Tubes Attached to the nose cone will be a short section, measuring 8.87in which will serve as the payload bay. The payload bay will contain the mechanisms necessary for the capture and safe containment of the payload before and during flight. An aright door will be machined for this section so that the robotic arm may have good clearance to insert the payload into the bay. The door will then be closed by the arm and will feature a self-locking mechanism. This section will be connected to the main section with a 4 inch coupler that will have an outer diameter of 5 in and an inner diameter of 4.5 in. It will be bolted into the payload body tube and a bulkhead will seal the payload bay off from the rest of the vehicle compartments. A diagram showing the components of the payload bay is shown in Figure 2.3. In the figure the payload is the small yellow cylinder toward the aft end of the model. Figure 2.3: Payload Bay 9

11 Section 2.2.2: Main Body Tube The first section behind the nose cone is the main body tube. It houses recovery elements and the avionics bay, along with all the supporting electrical components. The main body tube will measure 3 in long, and have an outer diameter of 5.25 in and an inner diameter of 5 inches. The avionics bay will be 8 in long and have an outer diameter of 5 in with an inner diameter of 4.5 in and can be seen in the figure to the right. It will contain all the electronics necessary for tracking and sending flight data back to the ground crew. All the electronics will be mounted on a custom made ABS bracket to accommodate all the necessary wiring. The avionics bay will be sealed off with bulkheads which will have custom made holes for attaching them to the CO2 recovery subsystem. In the middle of the avionics bay will be the external collar. It will have an inner diameter of 5in, an external diameter of 5.25 in and a length of 2 inches. This switch collar will feature 2 external key switches for the altimeters to reduce the amount of setup time required. Key switches also add safety, reducing team member s potential contact with bare wires, and reliability, ensuring continued connection throughout flight. Section 2.2.3: Engine Tube Figure 2.4: Avionics Bay The last section, following the main body tube, is the engine tube. Its length is 39 in and it has an outer diameter of 5.5in with a wall thickness of 0.25 inches. It houses the ballast tank which will be used in case the center of gravity of the rocket needs to be brought closer to the center of pressure of the rocket or to add mass for a specific altitude goal. Between the ballast tank and the motor is a bulkhead plate and an engine block, both of which secure the motor and keep it separated from the other components. The motor will be secured with a carbon fiber tube manufactured with an internal diameter matching the outside diameter of the motor. A screw cap and retainer will be 0

12 epoxied to the end of this tube to provide proper backing to the motor thus securing it in the vehicle. Centering rings, made out of carbon fiber, will hold the motor mount in place and ensure that the motor remains in the center of the rocket and that thrust is produced uniformly and along the vehicle's central axis. The motor retention system can be seen in Figure 2.5. Figure 2.5: Engine Tube Cutaway Section 2.2.4: Construction Methods With the assistance of GKN Aerospace, the team has elected to utilize the equipment at their facility in Tallasee, Alabama. Filament winding the body of the aircraft provides a unique opportunity to reduce the weight of the rocket. By utilizing unidirectional strength carbon, which is typically much thinner and much lighter than twill-weave or other woven carbon cloths, the team can significantly reduce the weight of the rocket. Normally, this presents trade-offs in tensile or compressive strength properties, as the unidirectional carbon is not as strong when not loaded axially. However, with careful consideration of the ply orientation through combining several different ply orientations within the design, the material can be constructed to be much more versatile in its strength properties, while still maintaining the weight savings provided by unidirectional carbon. By teaming up with GKN, Project WALL-Eagle is able to gain the vast experience of their composite technicians in helping to design the ply design, as well as using their filament winding apparatus to actually construct the body tubes of the rocket. In addition, with the precision of a CNC guided machine winding the fibers, the final finished quality of the body tubes is significantly enhanced from a hand layed-up body tube. This results in a much more aerodynamic surface for air to flow around the body tube, with far less work required to refine the body tube.

13 For the collar, avionics bay and brackets the TAZ4 3d printer will be used. Designs printed from ABS are very accurate, the printer s accuracy is 50 microns, and allows for custom hardware to be created. This also gives the students a chance to design and test components, allowing for quick iterations and producing many similar parts which are really simple to scale. Section 2.3: Nose Cone When selecting the nose cone for our rocket, the team considered several standard designs. The merits of four types of nose cones were considered: ellipsoid, conical, Haack, and ogive. When comparing these cones, the primary characteristics considered were the coefficients of drag, the mass of the cone, and the ease of manufacturing the nose cone. Table 2.: Summary of Nose Cone Trade Study Type of cone Coefficient of Drag Mass Ease of Manufacturing Total Ogive Haack Ellipsoid 2 4 Conical Section 2.3.: Size and Shape Selection The coefficient of drag affects the overall performance of the rocket in flight. The goal for the team was to select a nose cone shape with a low drag coefficient in order to maximize performance. Utilizing the software OpenRocket, the four cone types were compared using the already chosen dimensions for the rocket. The ogive and Haack nose cones compared favorably with each other, with a difference that was negligible, while the ellipsoid and conical cones had relatively higher drag coefficients. The mass of the rocket has a great effect on its performance, with a cone with a lighter mass being preferable. Once again, the OpenRocket software was utilized, in this case to estimate the mass of each nose cone type. The conical nose had the lowest weight, estimated at lb. The ogive and Haack cones were close in weight, at lb and 0.85 lb respectively. The ellipsoid nose cone had the highest weight, at lb. 2

14 As the rocket must be manufactured 'in-house', it is important the nose cone be one that can be made relatively easily with the tools available to the team. A conical nose cone is an extremely simple shape and is relatively easy to manufacture with the mills, lathes, and carbon fiber production tools available at Auburn. The ellipsoid and ogive nose cones, while slightly more complex in design, are still relatively simple and easy to produce. The ogive cone is also commonly used in hobby and professional rocketry, so there is a large amount of off-the-shelf examples available and information on the production of such a nose cone is easy to procure. Haack nose cones are not based on simple geometry, instead being mathematically derived for minimum drag. While still something that the team could theoretically produce, a Haack nose cone would take more time and effort to get right when compared to the other types, for a minimal gain in effective altitude given the team s low speed application. Given the above, the team has decided on an ogive nose cone, as it rated higher than the Haack and ellipsoid nose cones during the team's trade study. While the conical nose cone was equal to the ogive overall when compared in the trade study summarized above, the decision was made to take the ogive due to its better drag performance. While somewhat higher in weight and slightly more complex to manufacture, the difference in drag coefficient was enough to outweigh the conical nose cone's advantages in those other areas. The dimensions of the ogive nose cone designed for this rocket are shown in Figure 2.6. Figure 2.6: Nose Cone Dimensions 3

15 Section 2.3.2: Manufacturing Method To manufacture an ogive nose cone out of carbon fiber a precise 3d printed part will first be created. The ABS nose cone will be used to make a fiberglass mold. This will be a half female mold that will have the nose cone and collar. It will allow for pre-preg carbon fiber sheets to be laid into it creating the nose cone and a flange. This will allow for 2 halves to be assembled and epoxied easily with the flange, which will be removed and sanded smooth once the epoxy is cured. To ensure that the correct shape is obtained, the desired thickness will be removed from the fiberglass mold in order to make room for the flange. To create the model, the ogive tangent nose cone equation was used: P = 2 2 ( R L ) + 2* R Where R is the outer diameter of the nose cone and L is the length. P is the ogive tangent circle which generates the outer curve of the nose cone. R must lie on the radius of the circle that P generates. The mold will allow for many nose cones to be made precisely as long as proper vacuum bagging and curing procedures are followed. Section 2.4: Fins Fins form an integral component of the structure of the airframe. Fins create the large normal force required to stabilize the rocket should the thrust vector not be in line with the center of gravity. This will result in a further destabilization of the rocket if the aerodynamic moment does not counter the gravitational moment about the rocket. In the design of fins, several parameters are capable of being iterated in order to achieve the optimal fin design that creates a large normal force, while minimizing the drag being generated and also Figure 2.7: Various Fin Shapes 4 a) Clipped-Delta Fin b) Ellipsoidal Fin C) Trapezoidal Fin

16 surviving the large amount of stresses encountered by the fins during the flight. Primarily, thickness and cross-sectional shape, overall fin shape, and thickness. Section 2.4.: Shape Selection For the shape of the fins, the team chose between three different shapes, the trapezoidal fin, the ellipsoidal fin, and the clipped-delta fin. They were analyzed based on their aerodynamic properties, as well as their ability to stabilize a rocket. Current research shows, that when compared to its trapezoidal counterparts, the elliptical fin shape has an aerodynamic advantage in that it typically produces lower drag. However, since the traditional aerodynamic analysis uses a flawed model based on assumptions that the fin receives clean, unaltered flow this advantage is negligible at the velocities the rocket will achieve. Additionally, the flawed model also assumes that the tip of the airfoil operates at the same effectiveness as the root of the airfoil, however since the flow at the tip of the fin receives a higher speed flow, the tip of the fin creates a much higher normal force coefficient then the root of the chord. Thus, fins with a larger tip chord, in practice, receive much higher normal force coefficients. This increase in the normal force coefficients presents in the form of a much quicker stabilization of the rocket as it tilts away from a perfectly vertical orientation. In this case, the static stability margin is calculated by the equation: xcg xcp SC.. = D D ref is a reference diameter, in the case of most simulations, the largest diameter on the rocket. While the equation is always valid, most simulations assume an incorrect location of the center of pressure due to the faulty assumption in flow conditions on the fins. In this manner, elliptical fins correct the instability much slower than fins that taper less towards the tip. Furthermore, the elliptical fin shape has complex geometry that would be difficult to manufacture compared to the trapezoidal fin shape and clipped delta fin shape. While all fin designs are easily layed-up as flat plates, and the computer numerically controlled router can accurately cut the curve of the elliptical shaped fins, the complex curve would result in significant deterioration of the edge of the carbon fiber, requiring significant extra work in order to develop a clean, aerodynamic edge. ref 5

17 This extra complexity could potentially lead to mistakes during manufacturing, which could mean a waste of resources. The clipped-delta is a type of a trapezoidal fin that has a trailing edge that is not angled. Although the traditional symmetric trapezoidal fins excel at supersonic speeds, the clipped-delta fins perform with better efficiency at subsonic flight. The trapezoidal fins main advantages over the clipped delta are in stability and its shape. The shape ensures that upon landing the fin tips are less likely to suffer most of the impact since the corners are angled away. Thus, with the increase in the difficulty in production of the elliptical shaped fins, coupled with their decrease in corrective stability and the trapezoidal fins ease of manufacturing and convenient shape the trapezoidal fin was chosen as the fin shape. The design parameters discussed in the paragraphs above are summarized in Table 2.2. Table 2.2: Summary of Fin Shape Trade Study Ease of Type of Fin Stability manufacturing Drag Total Trapezoidal Clipped Delta Elliptical Section 2.4.2: Construction Methods (Bulkplates Included) Within the rocket, bulkplates provide the structural stability for couplers and other internal interfaces. As such, it is paramount that they be constructed out of robust and sturdy materials. Fins also must be extremely sturdy, as they experience high loads on the exterior of the rocket due to high dynamic pressures and large vibrational stress. Once again, given the abundance of carbon fiber at the team s disposal, and the high strength to weight characteristics of carbon fiber, the team has elected to use carbon fiber for the bulkplates. However, this presents significant tradeoffs in weight and cost for the team, as well as being slightly harder to machine and much harder to integrate with bonding systems. Despite the significant tradeoffs, the team believes that with correct application of adhesive and other fasteners, as well as the low amount of material required to build bulkplates and fins, the ease of manufacturing and strength properties of carbon fiber produce a much superior end product to be integrated into the rocket. 6

18 Since carbon fiber is the selected material, the construction method is to manufacture flat plates using a compressive curing technique. The compression machine, while inherently limited to only flat applications, provides exceedingly good ply consolidation in composite manufacturing. In addition, given the relatively low number of projects applicable to its use, the compressor has much fewer demands for usage, enabling a much less strict schedule of use for manufacturing flat plates. The bulkplate dimensions are show in Figure 2.8. They will consist of an inner and an outer plate to seal with the collars and bays inside the vehicle. The avionics bay bulkheads will have the CO2 subsystem attatched to them. Once the flat plates are constructed, the team will use a CNC controlled router in order to cut the correct fin or bulkplate dimensions out of the carbon fiber plates. This allows for a variable thickness, as well as rapid prototyping using 3D modeling techniques. Figure 2.8: Bulkplate Dimensions 7

19 Section 2.4.3: Size In order to successfully stabilize the vehicle, the fins must be large enough so that they can create a normal force acting on the fins to counteract the moment of the mass rotating the rocket about its center of gravity. Having determined the ideal shape of the rocket s fins, the sizing must then be determined in order to achieve a subsystem that functions as intended. The overall dimensions of each individual fin are shown in Figure 2.9. To ensure that the fin is attatched appropriately, it will be epoxied to the motor mount, engine tube and motor centering ring. Section 2.5: Motor Figure 2.9: Fin Dimensions Motor selection is a highly important parameter to the success of the mission, therefore careful consideration must be applied to ensure optimum motor performance. Section 2.5.: Loki Research K960-P The rocket motor initially selected for the competition in proposal was the Aerotech K780R-P. Due to the limited availability of this motor the K960-P was chosen for its similarities and availability. The K960-P s specifications are listed below in Table 2.3 (next page). 8

20 Table 2.3: Motor Specifications Manufacturer Loki Research Motor Designation K960-P Diameter 54mm Length 498 mm Impulse 949 N-sec Total Motor Weight 3.85 lbm Propellant Weight 2.05 lbm Propellant Type Redline Average Thrust 225 Pounds Maximum Thrust 345 Pounds Burn Time.95 sec Figure 2.0: K960-P Thrust Curve This motor was chosen based on OpenRocket simulations, as it provides an initial thrust-to-weight ratio above the required 5: ratio required to create a stable rocket. As shown in Figure 2.0: K960-9

21 P Thrust Curve, the motor achieves a higher than average thrust early on in the thrust profile, reaching the required 5-to- thrust ratio in less than a quarter of a second. The maximum altitude achieved by the K960-P was 3352 feet at the current estimated weight, with an average wind speed of 5 mph. Given the increase of 2.97% however, the projected altitude was 3079 feet, displayed in Figure 2.. While still above the mission altitude of 3000 feet, it is assumed that the actual altitude of the rocket will be slightly lower than the simulated height, as the imperfections in the model OpenRocket uses tends to be more idealistic than the rocket can attain. Additionally, the average wind speeds may be higher than the simulated 5mph, further reducing the expected altitude of the rocket. Simulations of the rocket s flight at higher wind speeds reach well below the 3000 feet target. Thus, should the weather conditions be off the nominal, the rocket will still be below the maximum altitude. Figure 2.: Altitude vs. Time K960-P 20

22 With this in mind however, a mass increase of 2.9% is well within the mid-range of the expected mass increase. Should the rocket increase more than that, to a respectable 25% mass increase, the primary selection of motor would not be capable of lifting the vehicle and payload to the required 3000 feet. Therefore, a secondary motor has been selected in order to ensure that the competition altitude can be met, given any mass increase. Finally, cost and availability of the motor is also a critical aspect of the motor selection. The selected competition motor can be purchased by several different vendors, including the team s primary vendor, Chris s Rocket Supplies. Section 2.5.2: Cesaroni L-60 The secondary motor selected for the competition should the mass increase surpass what is capable of the primary motor is the Cesaroni Technologies Incorporated L60, with properties summarized in Table 2.4. Manufacturer Table 2.4: L-60 Properties Motor Designation L-60 Diameter 75 mm Length 395 mm Impulse N-sec Total Motor Weight 8.7 lbm Propellant Weight 3.5 lbm Propellant Type Redline Average Thrust 37 Pounds Maximum Thrust 97 Pounds Burn Time 5. Seconds Cesaroni Technologies Incorporated Once again, the flight profile for the rocket is simulated using OpenRocket. Given a mass increase of 25% the altitude yielded is 3245 feet. In this scenario, additional mass could be added utilizing 2

23 the ballast tank in order to achieve the optimal mass for increased altitude, and an increase in the size of the fins would be necessary in order to compensate for the shift aft in the center of gravity, resulting in a loss of stability caliber. Once again, given a mass increase of 25%, the thrust-to-weight ratio of the rocket is higher than the 5: required ratio for stable, steady flight of the vehicle. Also, the motor once again reaches the required thrust very quickly within the burn profile of the rocket, as shown in Figure 2.2Figure 2.2. This results in a high stability of the rocket as it exits the launch rail. Figure 2.2: L-60 Thrust Curve Once again, the motor is readily available for purchase should the vehicle mass grow more than anticipated, the secondary motor should produce enough total impulse to deliver the vehicle to the desired altitude and thus successfully complete its subsystem requirements. Section 2.6: Mass Estimates The mass of WALL-Eagle and all of its subsystems was calculated using optimal mass calculations from OpenRocket. Since most of the parts will be manufactured using carbon fiber, a brick sample 22

24 was created to have an accurate density measurement. This density test is exceedingly important given the method of mass estimation. Since construction methods vary drastically from each manufacturer, as well as different resin and cloth systems varying, it is highly important to get an accurate model of the density. The test brick was made in the Composites and UAV Laboratory by pressing layers of preimpregnated carbon fiber together on a polished aluminum surface and molding it into a brick. After the desired shape was achieved it was cured in the Blue-M Convection oven, using manufacturer s recommendations for curing. Once the cured piece was completed, it was machined to have as close to even, parallel sides so that an accurate volume measurement could be made. The block s dimensions were measured with calipers to receive measurements within 0.00 in. and since a parallel surface was very difficult to achieve by grinding an accurate model was made using Solid Edge to calculate volume of an oddly shaped object. After weighing the piece a new density was calculated plus a 2% adjustment for any inconsistencies, such as air pockets. The new density was used in the OpenRocket model to accurately measure mass, since the software calculates volume on its own, an accurate density for each part is the last critical parameter to obtain an accurate mass estimate. Having determined an accurate density for the carbon fiber of the rocket, and the structure of the rocket being the most significant portion of the weight of the structures of the rocket, the team used estimates from last year s rocket to determine the initial size estimate of the rest of the subsystem components. The team believes that this model presents an estimate that is sufficient until prototypes of each system are developed during the Critical Design portion of the program. As the program develops, the model will attain a higher and higher accuracy in its simulation. 23

25 Table 2.5: Overall Mass Budget Section Mass (lb.) Percentage of Total Weight Structure % Supporting Equipment % Electronics % Recovery % Motor % Total N/A Mass Growth % Mass Allowance % Shown above in Table 2.5 is the overall mass budget of the system as calculated by open rocket with the assumptions detailed in the previous paragraphs. Additionally, the mass budget is further detailed into each independent section in Table 2.6, where the components in the table are described in the list below. Structures The airframe of the rocket including fins and nose cone, as well as centering rings, bulkheads and the engine block. Supporting Equipment Internal bays that support structure and provide a frame for different subsystems to be mounted to. Includes the avionics bay, payload bay, and the ballast tank. Electronics All electrical components and wiring necessary to complete mission. Recovery Equipment necessary to deploy parachute and safely land vehicle. Motor Includes the motor as well as the casing and retention. Payload Bay The section to be used for payload, which will support all mechanical and electrical components of the subsystem. Avionics Bay The section that will contain onboard electronics for reading flight data and sending it to the ground crew, including all of the altimeters and tracking equipment. Ballast Tank Will be used in case more mass is required for stability or for optimal altitude. Bulkhead Seals off the motor compartment from the rest of the subsystems Engine Block Supports the motor and keeps it from moving into the other sections of the rocket 24

26 Centering Ring Aligns the motor and ensures that thrust is delivered in the same direction, relative to the rocket Table 2.6: Detailed Mass Budget Booster Section Upper Section Component Mass(lb.) Component Mass (lb.) Structure 3.2 Structure 3.83 Ballast Tank Payload Bay 3.07 Bulkhead Avionics Bay Engine Block 0.35 Recovery Centering Ring 0.38 Nose Cone 2.3 Fins Electronics.4 Recovery 2.22 Motor 3.85 Electronics 0. Total Total 3.48 The current configuration has an expected mass growth of 3.7 pounds, resulting in a total of 3% of the overall rocket mass. While being lower than the recommended 25% mass increase common in manufacturing processes, the team believes that the initial model of the mass is accurate enough to warrant not including the extra 3%. In addition, special consideration has been taken to ensure that the bottom portion of the rocket is modeled accurately, with any error leading to a heavier bottom portion of the rocket. Therefore, any mass increase is expected to be due to manufacturing, and therefore most likely forward of the current calculated center of gravity. As mass increases in the upper section of the rocket, the center of gravity moves upward and further away from the center of pressure. Should too much mass be moved forward, and the rocket become statically over stable, a ballast tank is positioned in the bottom of the tank, relatively close to the current center of gravity. Therefore, it is possible to move the center of gravity back towards a stable rocket, but should the rocket not gain the expected mass increase, the ballast tank can still be utilized to bring the rocket back to an optimal weight in order to deliver the vehicle and payload to the correct altitude. Mass is inherently one of the most critical parameters in the vehicle s design as most of the other subsystems rely on the mass for their design, from the recovery system to the motor selection. The recovery, and many other systems, are flexible and can be iterated with the changes in mass. Motor selection however, is extremely limited, and thus the selection of the motor is detailed in the next section. 25

27 Section 2.7: Manufacturing The majority of the construction of WALL-Eagle has already been practiced for the creation of the parts for the subscale. The subscale launch proved that all components will behave as they should and no problems should be encountered with the full scale build. The same manufacturing practices will be used along with the lessons learned to ensure that a quality product is made. Section 2.7.: Tests Tests have been performed on the 3d printed ABS collars to ensure that the parts have some strength. A picture of test setup is shown in Figure 2.3. A inch section was cut from 2 different samples, same dimensions and print settings, but different print jobs. These were then loaded into a continuous compression loading test apparatus to determine the failure load. The first sample failed at 58 lb and the second sample at 65 lb for a ¼ in thick coupler. A figure of the failed specimen is to the right. The results obtained from this test show that the collars should sustain regular flight and should be re-usable. The carbon fiber components have been field tested and performed as expected, they are durable and lightweight. Further testing is planned, such as tensile testing and other physical property testing, along with integration tests with other systems, such as recovery and AGSE. Section 2.7.: Quality Control Figure 2.3: Collar Testing To ensure all parts are precise, proper measurement techniques will be used along with calipers which measure in units of 0.00 inches. All machined parts will be created in Solidworks in order 26

28 . to check fit and clearances. These files can then be used to print and CnC precision parts in the manufacturing lab. Safety will be followed during assembly processes and parts will constantly be checked for fit and accuracy to ensure that a quality vehicle is produced. Section 2.7.2: Progress Currently all Manufacturing processes are well defined and practiced, except for the nose cone construction. A model of the desired nose cone has been created and printed. The current objective is to use this model to create a half female mold, the details of the process are outlined in the nose cone section. Production of full scale parts is already underway Section 2.8: Verification Criteria The requirements for the vehicle, along with their execution and verification method are summarized in Table 2.7. The following paragraphs expand upon the execution, giving specifics of the tests, demonstrations, or analysis being performed in each of the executions. As the design of the vehicle is further iterated and refined, the verification table shown below will be updated to reflect the self-imposed requirements demanded of a high-performance vehicle. Table 2.7: Verification Tables for the Vehicle AU Specific Number Requirement Section/Number Requirement Statement AU- Vehicle. The vehicle shall not exceed 3,000 feet AGL AU-2 Vehicle.2 Vehicle shall carry one commercially available, barometric altimeter AU-3 Vehicle.2. Altimeter shall report the official competition altitude via a series of beeps to be checked after flight. AU-4 Vehicle.2.2. At LRR, a NASA official shall mark the altimeter that will be Verification Method Test Analysis Demonstration Testing Inspection Demonstration Inspection Testing Inspection Demonstration Execution of Method Launch vehicle and check altimeters Purchase and calibrate one commercially available altimeter Test the altimeter to verify it creates audible beeps Complete safety check at LRR 27

29 used for official scoring. AU-5 Vehicle At the launch field, a NASA official will obtain the altitude by listening to the audible beeps AU-6 Vehicle At the launch field, all audible electronics except official altimeter shall be capable of being turned off. AU-7 Vehicle.2.3. The official, marked altimeter shall not be damaged. AU-8 Vehicle The team shall report to the NASA official to record altitude on day of launch AU-9 Vehicle Altimeter shall not report altitude over 5,000 AGL AU-0 Vehicle Rocket shall fly at the competition launch site AU- Vehicle.3 Launch vehicle shall be recoverable and reusable AU-2 Vehicle.4 Launch vehicle shall have a maximum of four independent sections 28 Inspection Demonstration Inspection Demonstration Testing Inspection Analysis Testing Demonstration Demonstration Testing Demonstration Testing Analysis Demonstration Inspection Demonstration Ensure beeps are audible, launch successfully Ensure all electronics can be turned off and back on Design the electronics housing to prevent damage to altimeter The team is timely and organized in gathering data and reporting to NASA official Design and test launch vehicle to meet altitude requirement Team will launch the rocket at the appropriate site on launch day Trajectory simulations and testing will ensure the launch vehicle is recoverable and reusable Team will design and build launch vehicle that can have, but does

30 not require, four independent sections AU-3 Vehicle.5 Launch vehicle shall be limited to a single stage AU-4 Vehicle.6 Launch vehicle shall be capable of being prepared for flight at the launch site within 2 hours AU-5 Vehicle.7 Launch vehicle shall be capable of remaining in launchready configuration for hour AU-6 Vehicle.8 Launch vehicle shall be capable of being launched by a standard 2 volt DC firing system AU-7 Vehicle.9 Launch vehicle shall use commercially available solid APCP motor system AU-8 Vehicle.9. Final motor choices shall be made by CDR AU-9 Vehicle.9.2 Any motor changes after CDR shall be approved by NASA RSO AU-20 Vehicle.0 The total impulse provided by launch vehicle shall not 29 Demonstration Demonstration Testing Demonstration Demonstration Demonstration Demonstration Demonstration Team will design and build a singlestage launch vehicle Team will be timely and organized to ensure vehicle is prepared on time Batteries shall be tested with full electronics to verify their life Vehicle will be designed and tested to be launched by the standard 2 volt DC system Vehicle will be designed around commercially available, certified motors CDR will determine which motor the team will use for competition If the change is made to increase safety margin, NASA RSO will allow the change Launch vehicle impulse will be designed to not exceed 5,20

31 exceed 5,20 Newtonseconds (L-class) AU-2 Vehicle. Provide an inert or replicated version of motor matching both in size and weight to the launch day motor AU-22 Vehicle.2 Pressure vessels on the vehicle shall be approved by RSO AU-23 Vehicle.2. Pressure vessels on vehicle shall have a minimum factor of safety of 4: with supporting design documentation included in milestone reviews AU-24 Vehicle.2.2 The low-cycle fatigue life shall be a minimum of 4: AU-25 Vehicle.2.3 each pressure vessel shall include a solonoid pressure relief valve that sees the full pressure of the tank AU-26 Vehicle.2.4 Full pedigree of the tank shall be described including application for which the tank was designed, and the history of the tank, including the number of pressure cycles put on the tank, by whom, and when Demonstration Analysis Testing Inspection Analysis Testing Analysis Testing Inspection Analysis Testing Inspection Demonstration Newtonseconds. The team will bring an exact inert or replicated version of motor on launch day. Inspection of pressure vessel by RSO standards by testing. The team will inspect the safety factor of the pressure vessels through testing with documentation. Testing of the low-cycle fatigue. Inspection of each pressure vessel and testing of the pressure relief valve to see does it work as inspected. The team will inspect the tank along with documentation of testing and history. 30

32 AU-27 Vehicle.3 Successfully launch and recover a subscale model of the full-scale rocket prior to CDR. Subscale model shall perform as similarly as possible to the fullscale model, full-scale shall not be used as subscale model AU-28 Vehicle.4 A full-scale rocket must be successfully launched prior to FRR in its final flight configuration. The rocket flown at FRR must be the same rocket flown on launch day. Demonstrate the launch vehicle's stability, structural integrity, recovery systems, and teams ability to prepare the launch vehicle for flight. (A successful flight is defined as a launch in which all hardware is functioning properly AU-29 Vehicle.4. Vehicle and recovery system shall have functioned as designed AU-30 Vehicle.4.2 The payload does not have to be flown during the full-scale test flight AU-3 Vehicle.4.2. If payload is not flown, mass simulators shall be used to simulate the payload mass. Demonstration Testing Testing Demonstration Testing Testing Inspection Inspection Demonstration A demonstration of the launch will be exhibit through testing. A test of the rocket will be exhiit demonstration all hardware functions properly. Testing of vehicle will show how recovery system functions. Inspection of rocket payload by team will be flown at fullscale test flight. Payload will be flown. 3

33 AU-32 Vehicle The mass simulators shall be located in the same approximate location on the rocket as the missing payload mass. AU-33 Vehicle If the payload changes the external surfaces of the rocket or manages the total energy of the vehicle, those systems shall be active during the full scale demonstration flight AU-34 Vehicle.4.3 Full-scale motor is not required to be flown during the full-scale test flight, however it is recommended AU-35 Vehicle.4.4 Vehicle shall be fully ballasted during fullscale test flight AU-36 Vehicle.4.5 Vehicle or components shall not be altered after final demonstration flight without permission from NASA RSO AU-37 Vehicle.5 The cost of the competition rocket and the Autonomous Ground Support Equipment (AGSE) may not exceed the budget of [$0000 or $5000] Inspection Demonstration Testing Inspection Demonstration Testing Demonstration Demonstration Demonstration Inspection of the rocket payload will be done by the team to ensure it is properly placed. Demonstration of the adaptability of the systems notice to payload changes of the external surfaces through testing. Inspection of the motor will be done by the team to ensure it is flown through fullscale testing. Testing of the vehicle will demonstrate it being fully ballasted. The team will demonstrate that it did not alter any components or vehicle after demonstration flight. The team will demonstrate its budget of the competition rocket to validate its cost. 32

34 AU-38 Vehicle.6. Launch vehicle shall not utilize forward canards AU-39 Vehicle.6.2 Launch vehicle shall not utilize forward firing motors AU-40 Vehicle.6.3 Launch vehicle shall not utilize motors that expel titanium sponges (Sparky, Skidmark, MetalStorm, etc.) AU-4 Vehicle.6.4 The launch vehicle shall not utilize hybrid motors AU-42 Vehicle.6.5 The launch vehicle shall not utilize a cluster of motors. Demonstration Demonstration Demonstration Demonstration Demonstration Inspection The team will demonstrate how the launch vehicle does not utilize canards. A demonstration of the launch vehicle will demonstrate it not utilizing forward firing motors. The team will demonstrate that the motor does not expel titanium sponges through test flight. The team will exhibit how the launch vehicle does not utilize hybrid motors. A demonstration and inspection of the launch vehicle to validate it does not use a cluster of motors. To ensure compliance with requirement AU-, the vehicle will have a test launch with the goal of attaining the 3000 ft apogee requirement of the competition. After the launch, the altimeter will be checked; should the vehicle fail to adhere to the requirement, modifications to the design will be made to correct any issues and the vehicle will be retested. To ensure compliance with requirement AU-3, the altimeter will be checked after a test launch of the vehicle to ensure that the altimeter reports the altitude reached via a series of beeps. 33

35 To ensure compliance with requirement AU-6, the switch that controls the vehicle's electronics shall be activated and deactivated to ensure that the electronics properly turn on and off on command. To ensure compliance with requirement AU-7, the altimeter shall be checked for damage after each test launch of the vehicle. Should any damage occur to the altimeter, the housing for the altimeter will be modified to ensure the altimeter will survive future flights, and the vehicle will be test launched again. To ensure compliance with requirement AU-9, the vehicle's altitude will be monitored during test launches. If the vehicle exceeds 5,000 ft AGL during test flight, steps will be taken as necessary to bring the vehicle's flight back into the acceptable altitude range. This may include adding/removing ballast weight, choosing a different engine, or similar measures. To ensure compliance with requirement AU-, the vehicle will undergo a test launch, and must be recovered intact and in a reusable condition. If the vehicle is not recoverable/reusable after this test launch, design changes be made as necessary will be made to ensure future iterations meet the requirement. To ensure compliance with requirement AU-5, the vehicle will be placed on its launch pad in launch-ready configuration for at least once hour as a test of the electronic system's battery life. To ensure compliance with requirement AU-22, any pressure vessels on the launch vehicle will have to meet the RSO's standards through standard testing. To ensure compliance with requirement AU-23, any pressure vessels on the launch vehicle will be put through testing to ensure that they meet a minimum factor of safety of 4. The results of these tests will be well documented and presented during milestone reviews. To ensure compliance with requirement AU-24, any pressure vessels on the launch vehicle will be put through low-cycle fatigue testing, and must have a minimum fatigue life of 4:. To ensure compliance with requirement AU-25, any pressure vessels must have solenoid pressure relief valves; these valves must be tested to ensure they function as intended. 34

36 To ensure compliance with requirement AU-27, a subscale model of the launch vehicle shall be built and launched before CDR. This model will be a separate vehicle from the actual launch vehicle, and will be designed to be as close to the actual launch vehicle in performance as possible. To ensure compliance with requirement AU-28, the final version of the launch vehicle will be completed before FRR, and will go through at least one full, successful launch to demonstrate the vehicle's adherence to general competition requirements. To ensure compliance with requirement AU-29, the recovery systems shall be fully demonstrated during the test flight listed under AU-28. To ensure compliance with requirement AU-33, if the payload changes the external surface of final vehicle design or alters the total energy of the vehicle, then those systems will be active during the test under AU-28. To ensure compliance with requirement AU-35, the vehicle must be fully ballasted during the fullscale test under AU-28. Section 2.9: Subscale Flight Results Section 2.9.: Subscale Model For the subscale flight, a scaling factor of 80% was applied to the full-scale rocket in order to correlate the modeling to real-world results. The scaling factor was chosen in order to closely resemble the full-scale flight. At 80%, the forces are highly realistic to what the full-scale would experience. Unfortunately, due to the testing schedule on the recovery system, the primary fullscale recovery system was not available for the January th subscale-launch. Thus, the blackpowder backup was utilized, since the full-scale purchased CO2 system requires significant design changes to the bulkplates to implement. In addition, the rocket was flown without payload or ballast bays. This was due to insufficient room for parachutes when included. Therefore, the OpenRocket Design was as follows: 35

37 Figure 2.4: Subscale Open Rocket Model With a predicted apogee of 2288 feet, and a static stability margin of.72 calibers. At 9.9 pounds, while the length was approximately 80% of the full-scale rocket, the weight was not. This can be attributed to removing several significantly weighted portions of the rocket. Section 2.9.2: Flight Results On January th, the team flew the subscale rocket at Phoenix Missile Works in Sylacauga, Alabama. The subscale rocket performed closely to the modelled rocket, with an apogee of 229 feet. Unfortunately, the flight was not fully successful for a multitude of reasons. Firstly, the packing length was barely sufficient to fit the parachutes without inclusion of the ballast tank and the payload bay when scaled. Additionally, the rocket failed to deploy a main parachute due to a 2500 Altitude vs. Time Graph 2000 Altitude (ft) Time (seconds) Figure 2.5: Altitude vs. Time for Subscale 36

38 failure of the nosecone to separate from the rocket. Since all shear pins were broken and both charges deployed upon recovery of the vehicle, the team is attributing the failure of the nosecone to separate to the masking tape applied to the nosecone in order to increase the friction between the nosecone and the body tube. The altitude vs. Time graph for the flight is shown in Figure 2.5. With the problems detailed, the rocket successfully deployed the drogue at apogee, but failed to deploy the main. This resulted in the rocket falling at a much faster descent rate than planned. Thanks to the solid design in the structural portion of the rocket however, the only damage sustained was a broken bond line in one fin, as well as a shear failure in the coupler. However, since the first subscale launch was not a completely successful subscale launch, the team is intending on launching a secondary subscale rocket for flight on January 3 st. This rocket will reflect the sizing changes reflected in the full-scale design, as well as the team s CO2 ejection system. Section 2.: Influence on Design Given the failures during the subscale launch, the team has made several drastic design changes. Primarily, the large increase in length of the full-scale rocket. With the extremely low amount of space present in the subscale rocket when packing the various parachutes, the full-scale design clearly required more space in order to pack the parachutes as well as the necessary bays in order to ensure mission success. Additionally, the team has increased the amount of ejection charge when switching to black powder, in order to ensure successful deployment of the main and payload bay parachutes, even in the event of additional friction between the nose cone. 37

39 Recovery Section 3.: Recovery System Outline The Auburn Student Launch team is using a dual-stage recovery system with a drogue deploy at apogee (target height approximately 3000 ft.) and a main deploy at 000 ft. The payload ejection requirement will be met by ejecting the payload bay and nosecone section together in conjunction with the main deploy charge. Section separation and parachute ejection will be achieved through the use of a custom CO2 charge system designed by the Auburn Recovery Team. The rocket will be recovered in three sections. The booster section and avionics bay are fully recovered under the main parachute and are connected by the shock cord of the drogue parachute. The nosecone and payload bay remain attached and are recovered together under a third parachute. This system is described in Figure 3.: Stages of Recovery. Figure 3.: Stages of Recovery 38

40 During recovery, the booster section weighs 9.9 pounds, the avionics bay section weighs 4.4 pounds, and the nosecone and payload bay section weighs 4.9 pounds. The final expected decent velocity is 5 feet per second under drogue, below the threshold velocity for safe main deploy 20 feet per second. The decent velocity for the booster section and the avionics bay section under main is 7 feet per second, with the nose cone and payload bay falling under a separate parachute with a velocity of 5 feet per second. The kinetic energy on impact of the booster section is 44.4 ft-lbf. The kinetic energy of the avionics bay and associated body section at impact is 9.7 ft-lbf. The kinetic energy of the payload bay and nosecone on impact is 7. ft-lbf. The kinetic energy of each section is well below the designated maximum of 75 ft-lbf. This data is summarized in Table 3.. Table 3.: Launch Velocities and KE Booster Section Avionics Bay Section Nosecone and Payload Weight 9.9 lbm 4.4 lbm 4.9 lbm Decent Velocity - Drogue 5 ft/sec 5 ft/sec 5 ft/sec Decent Velocity Main 7 ft/sec 7 ft/sec 5 ft/sec Impact Kinetic Energy 44.4 ft-lbf 9.7 lbm 7. lbm Section 3.2: Subscale Analysis The Auburn Student Launch Team participated in the Phoenix Missile Works launch on January 0, 205. The team launched a 4 inch diameter subscale rocket. Because the priority of this flight was to test the stability of the rocket design and not to test the individual subsystem components, the team stuck to basic launch procedures. The CO2 deployment system being developed for the full scale does not fit into the smaller subscale body therefore the team used basic black powder charges in this launch. The main and payload parachutes for this launch were commercial chutes while the drogue was custom made by the team for this launch. The team used two PerfectFlite Mawd altimeters with the main deploy set at 900 feet and 700 feet on the primary and backup altimeters, respectively. The flight had a successful drogue deploy at apogee however it did not deploy a main and therefore hit the ground only under drogue resistance. The launch failed at the main deploy due to a combination of three factors that led to unsuccessful section separation between the nosecone and the body tube containing the main chute forward of the avionics bay. The first problem was too many shear pins for the charge being applied. Four 39

41 shear pins were used at the launch site. While all the shear pins broke, the force required to break the pins was too high and did not leave enough force to properly separate the sections. The second issue was too little black powder. Using a rocketry rule of thumb of gram per inch of rocket diameter, the main deploy black powder charges were 4.5 grams. This proved to be not enough because although both charges successfully deployed at their desired altitudes, the section still did not separate. The team believes that although we correctly estimated the charge for the rocket diameter, the amount of black powder did not reflect the length of the section that needed to be pressurized to successfully separate the sections. The section of body tube between the forward end of the avionics bay and the end of the nose cone was approximately 4 inches, a considerable length of carbon fiber body tube. The third and final issue preventing section separation and a successful deploy was too much friction between the collar of the nose cone and the rocket body tube. The fit of the two pieces was too tight and the layer of masking tape, as is common practice in many launches of this scale, added too much additional friction. The team believes that this combination of errors is what led to the failure of the launch. Had any one of those components been eliminated the team believes, based on ground testing, material data, and previous flights, that the launch would have still been a success. To prevent failures of this type in the future, changes are being made to the teams launch day procedure. As the custom CO2 system is designed for the 5 in diameter rocket body and the subscale avionics bay did not have room for the Peregrine CO2 system, the team has to make a hasty switch to black powder. In switching to a backup plan we discovered that we need to have a more detailed procedure for how much black powder to use for that specific launch, rather than an educated estimate made on site. With this in mind, the team has come up with exact calculations for the black powder needed in all future subscale and full scale launches, should the CO2 system be unavailable in any way. The team has also developed a chart that will be included in launch day supplies that give a more accurate estimate of the amount of black powder required for different lengths of 4 inch and 5 inch body tubes. This chart also contains the recommend number and size of shear pins that should be used with each associated charge. The separation failure experienced in the subscale launch is incredibly unlikely to occur with the CO2 system, dud to the gradual pressurization and the size of the canisters. The sizes of canisters commercially available for use with our systems, the Peregrine and the custom system we are 40

42 developing, are 8 gram and 2 gram. The 2 gram charges will always be used as they will not damage the rocket and will always provide enough charge to separate the rocket and eject the parachutes, even if deployed in perfect tandem, due to the gradual nature of CO2 pressurization. Section 3.3: Requirement Validation The Auburn depth has done an in-depth analysis of the requirements outlined in the Student Launch Handbook. The method of validation for all recovery requirements is outlined in Table 3.2 Table 3.2: Recovery Requirement Analysis and Validation Requirement Number Requirement Method of Validation 2. Deployment of Recovery Test Deployment System Devices 2.2 Ground Ejection for Drogue & Test Deployment System Main Parachute 2.3 At Landing, Max KE of 75ft-lbf Calculation and subscale testing for each Independent Section 2.4 Recovery system Electrical Circuits Independent of Payload Electrical Circuits 2.5 Recovery System Must Contain a Redundant, Commercially Available Altimeter 2.6 Exterior Arming Switch for each Altimeter 2.7 Dedicated Power Supply for each Altimeter 2.8 Arming Switch Capable of being Locked in the ON Position 2.9 Removable Shear Pins used for Main & Drogue Parachute Compartment 2.0 Electronic tracking Device Installed in Rocket to Transmit the Location of the Tethered Vehicle or any Independent Section to a Ground Receiver 2.0. An Active Electronic Tracking Device shall be connected to any Make Separate Electrical Circuits for Recovery and Payload Add Redundant Altimeter to Recovery System Add Exterior Arming Switch for each Altimeter Put Separate Dedicated Power supply for each Altimeter Make Sure Locking Mechanism Locks the Switch when Turned ON Put Removable Shear Pins for Main and Drogue Parachute Place Electronic Tracking Device in the Tethered Vehicle and to any Independent Section and Test the Signal Location Attach Electronic Tracking Device to any Independent 4

43 Independent Rocket Section or Payload Component The Electronic Tracking Device shall be fully Functional during Official Flight at Competition Launch Site 2. Recovery System Electronics shall not be affected by other On- Board Electronics during Flight 2.. Recovery System Electronics must be placed in a Separate Compartment away from any other Radio Frequency/ Magnetic Wave Producing Device 2..2 Recovery System Electronics Shielded from all On-Board Transmitting Devices 2..3 Recovery System Electronics Shielded from all On-Board Transmitting Devices Producing Magnetic Waves 2..4 Recovery System Electronics Shielded from any other On- Board Transmitting Devices Section & Payload Section and Test the Signal Location Test and Bring extra Electronic Tracking Devices at the Official Flight Test Recovery System Electronics along with the other On-Board Electronics to Ensure Signal Strength Recovery System Electronics will be placed in a separate compartment away from any other Radio Frequency/ Magnetic Wave Producing Device And Tested to Ensure Signal Strength The Recovery System is designed be shielded by the avionics bay. The Recovery System is designed be shielded by the avionics bay. The Recovery System is designed be shielded by the avionics bay. Section 3.4: Parachutes The dual stage recovery system with the addition of an ejected payload requires the use of three parachutes. All three parachutes will be designed and constructed by the Auburn team. The parachutes are constructed from ripstop nylon fabric sewn together with nylon thread. Currently the parachutes are designed to be semi-ellipsoidal in shape, based on historical data from the team using semi-ellipsoidal parachutes in the past. Extensive wind tunnel testing and analysis will compare semi-ellipsoidal parachutes to hemispherical, cross-form, and other custom shapes of parachutes. In addition, the team will be comparing parachutes with spill holes to parachutes without. Traditionally in high powered rocketry the standard parachute material used is ripstop nylon. Project WALL-Eagle however refuses to settle for the standard and instead has decided to prove 42

44 which material would be the most effective for the project. This was done using a trade study which took into consideration several key factors vital to mission success. For parachute material the primary factors for material selection were cost, ease of access, and tensile strength. Cost is always a major factor when working a project with a fixed budget. 30 inch wide by foot long sections of each material were used when looking for price estimates. These sections can be cut and sewn into whatever shape needed. Ease of access is another factor which should always be in consideration. The Auburn team will be making their own parachute this year. Because of this excess material will be needed for practicing sewing and hemming techniques. If the material cannot be easily found, it will restrict the capabilities of the team. Tensile strength is the last factor considered in this trade study. The mission s success depends on having a material that can adequately slow down the rocket and payload to a safe speed. Table 3.3 shows the materials and decision factors used in the trade study. Decision Factors Materials Table 3.3: Parachute Material Trade Study Decision Factors Cost (30 in wide/ft) Ease of Access Tensile Strength (ksi) Ripstop Nylon $.50 High 96 Silk $0.67 Medium 72.5 Kevlar $2.93 Low 525 Terylene $.84 Low 0.8 Table 3.4 shows the rubric created to rate the materials based on the decision factors deemed important to the project. Table 3.4: Parachute Material Trade Study Rubric Scoring Cost (30 in wide/ft) Ease of Access Tensile Strength (ksi) > $.50 Low 50 2 $.50 Medium 00 3 $0.75 High >00 43

45 All that was left to do was decide which factors were most important to mission success based on a scale from to 3. Most of these materials are relatively cheap when compared to the rest of the budget. For this reason cost was rated a 2 out of 3. Ease of access was deemed the most important factor by recovery team members, as being able to work and practice with the materials in order to get the best parachute possible will be vital to ensuring mission success. Tensile strength was also rated a 2 out 3 as most of these materials are known and successful parachute materials and all could suffice if used properly. The trade study results are show in Table 3.5. Factors Table 3.5: Parachute Material Trade Study Results Weight (-3) Ripstop Nylon Silk Kevlar Terylene Cost Ease of Access Tensile Strength Raw Score Ranking As shown, ripstop nylon came out as the number selection for the parachute material. This supports the well-practiced fact of ripstop nylon being the typical selection for high powered rocketry parachutes. The Auburn team has used ripstop nylon before and already has a vendor from which to get materials. Because of this and the information from the trade study, ripstop nylon will be used to make all of Auburn s parachutes. After motor burnout the rocket weighs approximately 23 pounds, 9.9 pounds in the booster section, 4.4 pounds in the avionics bay section and 4.9 pounds in the payload and nosecone section. All remaining weight is in the recovery system (parachutes, harnesses, shock cords, connection hardware, etc). For the rocket to reach a velocity for safe deploy a drogue parachute of 20 inches in diameter is required to reduce the rocket to 5 ft/sec descent velocity. Assuming a standard parachute Cd of.5 for a semi-hemispherical design, the following formula was used to calculate parachute diameter: 44

46 FF_DD = /2 ρρ CC_DD AA vv^2 Note that the parachute shape study is still in progress and should a different shape parachute be chosen as a result of that study, the parachute will be sized so that it produces the same drag force as the semi-hemispherical chutes, despite it having a different drag coefficient. At 000 feet, the main and payload parachutes eject. The main chute will be responsible for recovering the booster section and avionic bay section. The sections are tethered by the shock cord of the drogue chute and weigh a combined 4.3 pounds. For this weight and a final descent velocity of 7 ft/sec, the main chute needs to be 99 inches in diameter, assuming a semi-hemispherical shape. This results in an impact KE of 44.4 ft-lbf and 9.7 ft-lbf for the booster section and avionics bay section, respectively. The nosecone and payload bay fall under the influence of their own parachute. As the payload of any mission is assumed to be delicate cargo it should recovered with the utmost care. To reduce the impact kinetic energy and to reach a final descent velocity of 5 ft/sec, a 65 inch diameter parachute is required. Again the diameter for the main and payload parachutes are determined assuming a standard semi-hemispherical shape with a drag coefficient of.5. Should the parachute shapes change as a result of the shape study, the impact kinetic energy and final descent velocity will not change. The drag values calculated are final, assuming the weight estimates are final. Should maximum ballast be required for any reason, the designated parachutes will still recover the rocket components within the safe range of impact kinetic energy (up to 75 ft-lbf). Section 3.5: Ejection System The Auburn Student Launch team has decided to make the switch from a black powder ejection charge system to a CO2 ejection system after in-depth analysis of both systems. The CO2 system is safer, minimizing the need for team members to handle black powder. In addition the system is more scalable. Black powder charges can fail in high altitude/low oxygen environments whereas a CO2 system will be equally reliable at any altitude up to 27,000 feet. After experiences working with black powder charges in previous launches, the team believes that a CO2 system is a more elegant and precise solution to the challenges presented in the recovery of a rocket. 45

47 Since the CO2 inside a standard 2 gram cartridge is partly gaseous and partly liquid, it is a very complex system and using the ideal gas law is not an appropriate approach. The team is currently working on creating an accurate analytical model for this complex system. For now, the team will rely on other s experimental data and extensive ground testing. The viability of using a CO2 ejection system has been demonstrated repeatedly in the field and several premade commercial systems are available. Rouse Tech provides a chart that indicates which standard CO2 cartridge is required for different rocket configurations, as shown in Figure 3.2 Figure 3.2: Recommended CO2 Charges For Rocket Body Sections For a rocket diameter of 5 inches and parachute compartments up to 22 inches in length, a single 2 gram CO2 cartridge is recommended. This will be the baseline for ground testing. The team will determine the minimum amount of CO2 required as well as the amount of stress applied to the airframe via ground testing. For the scale of this project, standard 8 gram and 2 gram CO2 cartridges would be the most cost efficient way to achieve ejection. However, the team is also experimenting with refillable CO2 canisters that would allow for a customized charge to better suit the needs of each ejection. The refillable cartridges will be compared to standard sizes on a basis of cost and desired degree of confidence in deploying parachutes without damaging rocket body. 46

48 Figure 3.3: CO2 Cartridge Assembly - Exploded View The Auburn USLI team has designed a custom CO2 ejection system. While this system is inspired by Tinder Rocketry s Peregrine CO2 Ejection system, Auburn s custom system is designed specifically for our needs and has increased reliability and fault tolerance. Figure 3.3 shows the exploded view of Auburn s CO2 ejection system. 47 Figure 3.4: CO2 Ejection System

49 The ejection system is triggered by igniting a small Pyrodex charge with an e-match. The force of the charge pushes a 2 gram CO2 cartridge into a pin that punctures the cartridge. The CO2 is then released, pressurizing the parachute compartment with enough force to separate the rocket and deploy the parachutes. The entire ejection system consists of three of these CO2 assemblies as shown in Figure 3.4 While only one 2 gram CO2 cartridge is needed to separate the rocket, having three separately controlled CO2 assemblies provides dual fault tolerance for the CO2 ejection system. The team will carry out ground testing to determine if puncturing three cartridges simultaneously will pose a risk to the structural integrity of the airframe or damage to the parachutes. The rocket will use two ejection systems total, one placed on either side of the avionics bay as shown in Figure 3.5. Figure 3.5: Avionics Bay Layout One system will deploy the drogue parachute and eject the payload bay at apogee, approximately 3000 feet in altitude. The second system will deploy the main parachute at 000 feet in altitude. Each altimeter will control both CO2 assemblies on each system, providing redundancy. The CO2 cartridges are placed flat along the exterior of the avionics bay bulk plates instead of down the length of the rocket body. This makes the system becomes much more space efficient and minimizes interference with the parachutes. In addition, the team has decided to move the CO2 systems outside of the avionics bay. This provides an extra layer of protection to the electronics housed within the avionics bay; the altimeters will be shielded from the temperature changes that occur when the cartridges are punctured as well as protected from any form of misfire or cartridge explosion. However, by moving the ejection systems out of the avionics bay moves them into the body tube sections containing the parachutes. The parachutes will be protected by their deployment bags. The ejection system s own casing is being designed to sufficiently minimize risk 48

50 of damaging the parachutes. In addition, the team is constructing deployment bags to further protect parachutes from harm during launch and ejection. This will be verified through extensive ground testing. The team has decided to use Pyrodex instead of black powder due to its increased efficiency as a propellant and increased safety. Pyrodex is designed to be a volume-for-volume substitute for black powder. Other similar commercially available CO2 ejection systems use a charge of 0.5 grams, which provides a baseline for ground testing. Pyrodex is not as sensitive of an explosive as black powder, which makes it much safer. However, the decreased sensitivity means that Pyrodex charges need a higher level of confinement to ensure a complete burn. This additional confinement is provided in the design of the ejection system, which has each charge completely sealed in a charge cap within the main housing of the ejection system. Additionally, each charge cap and each lift piston has an O-ring along the circumference that creates a vacuum seal around each Pyrodex charge. Confining the charges to this degree further ensures the safety of the parachutes. The following launch procedures will be followed to ensure proper functioning of the ejection system. Before assembling the system, first clean all the parts thoroughly to remove residue from previous use. Then, insert the igniter into the hole on the bottom of the charge cap so that the ignitor head rests inside the reservoir of the charge cap. Place one drop of epoxy into the center hole of the charge cap to seal the ignitor into place. Once the epoxy has dried, measure out grams of Pyrodex and fill the reservoir of the charge cap. Cover the reservoir with painters tape and cut off all excess tape. There must not be any tape hanging over the edge of the charge cap. Next, load the charge cap into the base of the inner casing cylinder so that the painters tape faces the inside of the cylinder. Load the inner casing cylinder bottom first into the charge half of the casing, making sure the ignitor wires come through the holes at the base of the casing. Use a silicon based lubricant to lubricate the outer edge of the lift piston. Slide the lift piston into the case cylinder already inside of the casing with the curved side facing out. Use the butt end of the 2 gram CO2 cartridge to push down the lift piston completely to the bottom where it meets the charge cap. Place the alignment collar over the tip of the CO2 cartridge. Repeat these steps for each chamber of the charge half of the casing. Next, place a spring into each of the chambers of the pin half of the casing. Laying both halves of the casing flat on a table, with the holes facing 49

51 each other, slide the two halves of the casing together, making sure the CO2 cartridges enter their proper chambers. Holding the two halves together, secure the casing halves shut with the casing screws and nuts. Now that the system is loaded and assembled, it is ready to be placed in the rocket and have the ignitors wired to the altimeters. After launching successfully, carefully inspect the system to ensure that all charges have fired and CO2 cartridges are punctured and stable. While securely holding the system closed, carefully remove the casing screws. Once all screws are removed, slowly open the casing by separating the halves. Remove all components and clean thoroughly before reusing. Discard spent cartridges and ignitors. If the ejection is not a success for any reason, the team will take care in handling rocket components. The risks that student team members are exposed to a much less than those of a black powder system but undeployed canisters and systems will still be handled with extreme caution until cleared by the safety officer and launch range officials. Section 3.6: Electronics The team is using two altimeters to meet redundant system requirements. The primary altimeter will deploy CO2 charges at apogee for drogue deployment and 000 ft for main deployment. In black powder systems, the second altimeter is typically set to deploy at an altitude slightly off the desired altitude (our team typically uses 00 feet below desired altitude) to eliminate the risk of simultaneous deploy of both black powder charges. Simultaneously deploy with black powder is a situation that must be prevented as it could potentially cause considerable damage to the rocket such as fire damage and pressure failure of the rocket body itself. However, the new CO2 system does not pose the same risks to the rocket structure. The section pressurization is not explosive as with black powder. Rather, the section pressurizes gradually as the CO2 escapes the canisters after they are punctured. This means that as soon as the section reaches the force required for ejection, the force will break the shear pins and section will separate. The pressure cannot build up past this point because once the sections are separated there is not structure for containing this pressure. Therefore, all canisters can be deployed simultaneously. It is unlikely that the event will be exact simultaneous due to calibration errors between the two different types of altimeters, but even if 50

52 a) AltusMetrum Telemetrum altimeter b) PerfectFlite StratoLogger Figure 3.6: Altimeters they were the system can handle the charge. This means that our second altimeter could also be set to deploy the drogue at apogee and the main at 000 ft. However, due to restrictions on the setting of our chosen backup altimeter, the backup main charges will be set to deploy at 900 feet. For the main altimeter, the team has chosen the AltusMetrum Telemetrum altimeter, shown in Figure 3.6a. This altimeter provides dual deployment support for the main and drogue parachutes. The flight will be subsequently tracked through an accelerometer and barometric pressure sensors that remain reliable up to a maximum altitude of 00,000 ft. In addition, the Telemetrum will provide the advantage of having an integrated GPS. The Telemetrum allows for GPS tracking of the The altimeter itself will be powered by a lithium polymer battery. The second altimeter will be a PerfectFlite Mawd, shown in Figure 3.6b. This is a switch from the original plan to use a PerfectFlite StratoLogger as the second altimeter. The team has made the switch to using the Mawd, or minialit/wd, due to availability to the team and our history using Mawd altimeters in previous launches, including the subscale launch. The Mawd sensor is valid up to 25,000 feet, lower than the Stratologger but well within the range for flights in this competition. The Mawd stores over five minutes of flight data and communicates with a set of audible beeps. It is easy to program on-site should the altitudes need to be changed for any launch day circumstances. 5

53 As a fallback, should either of these altimeters be proven to be deficient, a PerfectFlite Stratologger altimeter will be used in the deficient altimeter s place. The team also owns additional PerfectFlite Mawd altimeters that can be used in case of errors with the chosen altimeters. In addition to the GPS tracking provided byt the Telemetrum altimeter, standard commercial RF trackers will be added to each section of the rocket. The Auburn University team s final launch at the competition in 204 highlighted the need for excessive redundant tracking of each individual component of the rocket. Three RF trackers will be used in total: one in the payload bay, one attached to the main chute, and one in the booster section. In addition the teams contact information will be clearly displayed on all components of the rocket, should for any reason the components become separated and misplaced during launch. Section 3.7: Attachment Hardware The parachutes are attached to their respective bulk plates using a system of U-bolts and Quick Links. The team has made the switch from bent eyebolts used in past years to U-bolts to minimize interference between the contention equipment and the shock cords/ shroud lines to prevent tangling, an issue faced by the leach in previous launches. In addition, U-bolts eliminate the risk of straightening the loop of a bent eyebolt during deployment. U-bolts are a much more cost efficient option than forged eyebolts and provide the added benefit of distributing the load between two points of contact with the bulk plate. Shock cords will be attached to Quick Links using bowline knots and stopper knots as required to tie up loose ends. Tension testing will be conducted to confirm that the knots will not reduce the structural integrated of the shock cords below loads reached during flight. Using trade studies in the design of the recovery system is a useful tool when comparing and selecting materials. This allows the recovery team to find the correct material without over or undercompensating for the task at hand. It also allows for a good reference point to come back to if the team discovers that a stronger or more durable material is needed. 52

54 Selecting the proper material for shock cords is a particularly challenging task to do for each project. This year the team is moving to use a CO2 deployment system, which removes the risk of burning the shock cord. However, if an effective CO2 system is unable to be developed in time the recovery team has decided to fall back on what it knows and use black powder. Knowing this, flame retardance remained a key factor when selecting shock cords. Decision Factors Materials Table 3.6: Shock Cord Trade Study Decision Factors Tensile Strength (lbs) Melting Point ( F) Average Price ($/per ft) /2" Tubular Nylon /6" Tubular Nylon " Tubular Nylon /8" Tubular Kevlar /4" Tubular Kevlar /2" Tubular Kevlar As seen in Table 3.6 the team s primary decision factors for selecting shock cords are tensile strength, melting point, and the average price per foot of cord. While there were only two different types of materials listed there were several thicknesses of each material compared. These two materials are notoriously known to be used in high powered rocketry and the reasons why are easily seen by the property values shown in the table above. Scoring each material was the next step in of the shock cord trade study. A rubric was created to show the scoring values which were put into the trade study table. The rubric is shown in Table 3.7. Scoring Table 3.7: Shock Cord Trade Study Rubric Tensile Strength (lbs) Melting Point ( F) Cost ($/per ft) > > 4500 >

55 Factors Tensile Strength (lbs) Melting Point( F) Average Price ($/per ft) Weight (-3) Table 3.8: Shock Cord Trade Study Results /2" Tubular Nylon 9/6" Tubular Nylon " Tubular Nylon /8" Tubular Kevlar /4" Tubular Kevlar /2" Tubular Kevlar Raw Score Ranking After determining the score of each material, the next step was determining which factors were most important to the project. Tensile strength was seen as the most important factor as the shock cord will be put through tension when the recovery event occurs. This is an unavoidable circumstance and having a strong shock cord in tension is vital to successful recovery. Also, while black powder is not the primary consideration for separating rocket components this year, it was still a major factor when considering the importance of mission success should a CO2 method of ejection not be deemed possible this year. For this reason, the melting point was also considered a concern of primary concern in the trade study. Price was considered to be of moderate significance as these materials are relatively cheap in comparison to the rest of the budget and all these materials are easily bought from local or online vendors. The final scores of each material can be seen in Table 3.8. As shown the ½ tubular Kevlar comes in first when considering all of the design factors discussed. However, due to the switch to a CO2 deployment system and the reduction in height requirements, this material is slightly too much for what the team needs. Also, considering the lack of elasticity and the abrasiveness of Kevlar it poses the risk of snapping the cord or zippering the rocket. For these reasons, the tubular Nylon has been chosen as the primary candidate for the shock cord of the rocket. Should the team need a stronger material, ½ tubular Kevlar will be the teams go to material. The rocket requires 30 feet of shock cord connecting the booster section and the avionics bay section. The shock cord connecting the main parachute to the avionics bay is 5 feet, the reduced 54

56 length as a result of the shock cord being one sided, only connecting the parachute to one rocket section. Section 3.8: Manufacturing The recovery subsystem can be divided into three main categories: parachutes, ejection charges, and attachment hardware. Each of these categories has its own set of manufacturing and fabrication procedures. The parachute category involves sewing the drogue, main, and payload parachutes and attaching all of the required shroud lines and shock cords. After the shape, size, and design of each parachute is finalized a paper template is created from the 2D model of each gore made from the shape coordinates plotted in Excel and mapped using CAD software. The template is used to cut the appropriate number of gores from the rip-stop nylon fabric. The gores are pinned together in pairs in preparation for sewing. The seams are French slip style, providing strength and dual fault tolerance in the parachute seams. To construct a proper French seam, the wrong sides of the fabric (the inside of the parachute) are sewn together with a quarter inch seam allowance. The seam is then reversed and the right sides of the fabric (the outside of the parachute) are sewn together with a ¾ inch seam allowance that encases the raw edges of the fabric and the previous seam. The gore are sewn in pairs and the pairs are sewn together until eventually forming the halves of the parachute. The halves are joined in an additional French slip seam down the middle of the chute. This main seam further seals the ends of all the other seams by encompassing the tips of all the other gores within the fold of the French seam. The thread is utility nylon thread chosen for its strength properties. After the parachutes are sewn, the shroud lines are attached. Shroud lines are typically made to be one and a half times the diameter of the parachute to which they are attached. There are two options for shroud line materials. For smaller chutes, such as the drogue and payload chutes, 550 Paracord is used for its smaller size and lighter weight in comparison to tubular nylon. For the main chute ½ inch tubular nylon is used for shroud lines. A different method is required for 55

57 attaching each of the two different types of shroud lines. For paracord shroud lines, the shroud lines are attached to a parachute seam beginning four inches up from the edge of the parachute. The paracord is straight stitched to the inside of the seam flap, then the flap is folded over and a wide zigzag stitch over the entire width of the cord. This sandwiches the paracord between the seam flap and the outside layer of fabric, creating multiple surfaces of contact between the paracord and the fabric. For ½ inch tubular nylon, the shroud lines have a larger surface area and therefore can be attached simply by using heavy wide stitches. For larger chutes such as the main, the shroud lines should begin at least 6 inches from the edge of the parachute. The team used different fabrication methods when designing the CO2 ejection system and when testing prototypes. One of the first methods used in this process was 3D printing. Recently, the team gained access to a 3D printer and it quickly proved to be an invaluable tool for every iteration of the design process. Using ABS plastic, the team was able to easily construct models and parts for the ejection system which made the design of the system evolve much faster and much more fluidly. The final two-part casing of the ejection system consists of 3D printed parts. After 3D printing several prototypes, the team decided to machine all inner components out of aluminum using our in house machinist. The smaller pieces, such as springs, O-rings, screws, and nuts are store bought. Section 3.9: Subsystem Testing Ground testing, component testing, and subscale testing are absolutely critical to this year s recovery system. The introduction of the CO2 ejection system requires vigorous testing to verify that it is, in fact, more reliable than a black powder system. The handmade parachutes and custom deployment bags will require extensive testing to ensure successful deployment. The new ejection system testing begins with testing of an existing system, the Tinder Rocketry Peregrine ejection system. The team hope to obtain pressure, timing and reliability data through ground testing. Our custom built system will then be tested to ensure that it can reach a comparable level of successful deploys through testing. In addition, the system will be tested using custom cartridges to achieve more specific ejection pressures. This data will be compared with that of 56

58 standard cartridges to see if the introduction of a custom sized charge is necessary and if refillable cartridge provide a notable risk of pressure leakage or other modes of failures. One of the many features the team was anxious to test with the subscale launch was the CO2 ejection system. Unfortunately, due to the size of the 2 gram CO2 cartridges, Auburn s custom system could not be scaled down to fit inside the subscale frame. So instead, the team decided to implement a commercially bought CO2 ejection system as a proof of concept. The system that was selected was the Peregrine Exhaustless CO2 ejection system by Tinder Rocketry. The Peregrine system works in a manner almost identical to Auburn s custom system, however it orients the CO2 cartridges vertically within the rocket (perpendicular to the bulk plates). This allows the Peregrine system to fit comfortably within our subscale rocket. The Peregrine system also houses each CO2 cartridge assembly separately and expels the CO2 gas through the bulk plate to the next section, meaning that each assembly is placed within the avionics bay. First the team performed ground testing on the Peregrine system by assembling it, attaching it to a small bulk plate, and securing it in a safe testing area before triggering it. This test was to ensure that the system was working properly and that the charges would not damage anything nearby, since they were to be installed within the avionics bay. After a few attempts, the ground tests were successful, and the team became confident enough to integrate the Peregrine System into the subscale rocket. There are two ways the team is ground testing the CO2 ejection system. The first method is assembling the system, securing it in a safe testing location, and activating the system. With this method the team can ensure that the charges fire properly and the CO2 cartridges are punctured and properly venting. By carrying out this testing method, the team determined the viability of the design, as well as the proper amount of lubrication and Pyrodex powder used to ensure proper operation of the system. The optimal amount of Pyrodex for each chamber of the system is approximately grams. The second method of ground testing the CO2 system is assembling the CO2 system and installing it in the recovery section of the rocket. Then after packing the payload and main parachutes and attaching the nose cone, the section is oriented vertically and secured in a safe testing location. The system is activated. In a successful test, the ejection system will push off the nose cone and expel the parachutes from the recovery system. A shape study is currently being conducting comparing different shapes and designs of parachutes. foot diameter models are being made of all design options. The shapes currently in the test are 57

59 hemispherical with and without spillholes, semi ellipsoidal, with and without spillholes, sky angle, and flat. Each model will be tested in the Auburn University 2 foot by 2 foot wind tunnel. Testing will determine the coefficient of drag for each design and will compare them to weight, material cost, and ease of manufacturing. Until the results of the study are final, the team is working with semi-ellipsoidal parachutes. Tensile testing will be conducted on the parachute fabric and the tubular nylon used for the shock cords and shroud lines to obtain data confirming that the system will not break under the expected flight loads. Nylon machine screws will be subjected to shear testing individually and the data will be confirmed by successful separation of rocket sections in ground testing and subscale launches. Because the ejection system is CO2 rather than black powder, flammability is less of a concern for the parachutes and their corresponding shock cords and shroud lines. However, since our backup system involves black powder charges, testing will be done to ensure that parachutes will not incur damage when charges are deployed. The subscale flight served as a test of the black powder system. The failures of the main deployment in the subscale flight showed that we need to refine the calculations of black powder charges and the associated shear pins. In addition, it showed that the design could benefit from the addition of nomex blankets to further protect the parachutes from black powder charges. In a successful deploy black powder damage is minimized but when the sections do not successfully separate, the parachutes are exposed to flames within the close quarters of the rocket body tube and can see significant melting and flame damage. We would like to preserve as many components of the rocket as possible, even when experiencing failures so additional fireproofing when using the black powder system will ensure the reusability of parachutes and their harnesses. After comments made in PDR, the team performed pack tests to determine the amount of space that the parachutes and their associated harnesses will take up within the rocket. Using various parachutes available to the team, it was determined that inch length of 5 inch diameter rocket body tube is required for approximately every 9 square feet of parachute fabric. This length includes the shroud lines but does not include the attachment hardware at the end of the harness. 5 feet of shock cord plus the quick links on either end require approximately 2.5 inches of 5 inch body tube length. Using this data, the main parachute requires 9 inches of body tube with

60 inches for shock cord. The payload chute requires 3 inches of 5 inch body tube with 2 inches of body tube for attachment hardware and shock cord. The total section length required between the avionics bay and the end of the payload bay is 6.5 inches. The designed length is approximately 9 inches allowing for a 5% growth in length at time of packing which should be more than sufficient, especially with the addition of parachute bags preventing an expansion of length. 59

61 AGSE Criteria Section 4.: Testing and Design of AGSE Equipment Section 4..: System Level Design Review The step-by-step requirements and design tasks for the AGSE system are summarized in AGSE Task Sequence Checklist below. Each requirement will be checked off as it is described throughout this subsection. AGSE Task Sequence Checklist AGSE System Tasks 0 minutes Scan an area of the ground and locate the payload Capture and transport the payload into the payload bay of the launch vehicle Secure the payload and seal the payload bay hatch on the launch vehicle Erect the launch vehicle to a position 5 degrees off the vertical Insert and secure igniter into the motor Verify that the sequence is complete and pause for inspection and launch preparation Payload Retrieval Subsystem (PRS) 5 minutes Launch controller initiates AGSE operations and solid orange safety light begins blinking PRS robot arm begins the payload search sequence until the payload has been detected Gripper orients with payload, payload is securely captured, and capture is verified Payload is transported to the launch vehicle payload bay and placed inside PRS robot arm returns to starting position and terminates payload search sequence Payload bay hatch seals and locks, tightly securing the payload in the launch vehicle Launch Vehicle Elevation Subsystem (LVE) 3 minutes LVE is initiated to begin lift sequence to raise rocket from the horizontal to the vertical position Launch vehicle is raised to 5 degrees from the vertical Truss is mechanically locked into upright position to prevent further radial movement 60

62 LVE terminates and the vertical position of the rocket is verified Automated Charge Insertion Subsystem (ACI) 2 minutes ACI is initiated to begin igniter insertion sequence Extendable telescopic rod is elongated to the top of the motor and sets charge in place ACI terminates with charge securely in place All systems pause and orange launch controller safety light stops flashing AGSE and rocket are inspected by the LCO Section 4..2: AGSE Structural Design and Materials The AGSE structure is designed to be functional, lightweight, easy to assemble, and intuitive to integrate all AGSE subsystems. Figure presents a computer generated model of the final structural design of the AGSE. Figure 4.: AGSE structural design model 6

63 The larger U-shaped box will be made of plywood and will have the following dimensions displayed in the figure below. Figure 4.2: U-shaped box dimensions for the base of the AGSE in inches The second box will support the top of the truss in the horizontal position, as well as serve as the base for the PRS. A schematic for the second box and its dimensions are displayed in the figure below. 62

64 Figure 4.3: Support box dimensions for the base of the PRS in inches These two boxes will be connected and held together by an A-frame of square aluminum tubes that are connected by 3D printed fitted connectors that are specifically designed to connect base structures together easily. AGSE Materials Trade Studies Trade studies were conducted for a variety of materials for the launch rail truss and the boxes at the base of the AGSE system. AGSE Base Materials Selection Firstly, a trade study was used to determine material with which to construct the box of the AGSE. This trade study allowed the AGSE team to determine the most appropriate material to use in construction of the box. If a new material is needed, the trade study can easily be referenced. 63

65 Four materials were considered for construction of the box: carbon fiber, aluminum, fiberglass, and plywood. The tensile strength, ease of construction, average price, and density of each material was compared. Since the box will not be subjected to a large amount of stress, the most important factors in choosing a material were average price and density. The initial comparison of materials can be seen in Table 4.. Decision Factors Materials Table4.: Box Materials Comparison Chart Tensile Strength (psi) Ease of Construction Average Price ($/lb) Density (lb/ft 3 ) Carbon Fiber Hard Aluminum Medium Fiberglass 3450 Hard Plywood 4500 Easy The next step in the trade study was scoring each material property. Each property was broken up into five ranges and the ranges were ranked from -5, with 5 being the most desirable. Table 4.2: Weighted scoring criteria for box material trade study Scoring Tensile Strength (psi) Ease of Construction Cost ($/lb) Density (lb/ft 3 ) <= 0000 Very Hard >.25 >200 2 <= Hard <=.25 <=200 3 <= Medium <=.00 <=50 4 <= Easy <=.75 <=00 5 >40000 Very Easy <=.50 <=50 After determining the score of each material property, it was necessary to determine which properties were the most important to the project. Tensile strength was deemed low priority since 64

66 the weight of the rocket on the large supporting area of the box will result in low stresses. Ease of construction was also rated as low priority because there is a large amount of time available to construct the box. Average price was the most important factor. Due to budget constraints, saving money on the box will allow more money to be spent on the robotic arm and materials used in construction of the rocket. The density of the material was ranked as medium priority as a lighter material will be easier to transport and set up. Factors Table 4.3: Box Material Trade Study Weight Carbon Fiber Aluminum Fiberglass Plywood (-3) Tensile Strength (psi) Ease of Construction Average Price ($/lb) Density Raw Score Ranking As seen in Table 4.3, plywood ranked first in the trade study and was chosen to be used for construction of the box. It is an extremely cheap and lightweight material that will fulfill everything required of the AGSE box. A thin layer of carbon fiber will be added to the top of the box in order to protect it from the flame emitted by the rocket on launch. If budget allows, the entire box will be made out of carbon fiber to greatly increase the factor of safety concerning the strength of the box. AGSE Truss/Launch Rail Material Trade Study Another trade study was used to determine the best material to use in construction of the truss supporting the rocket. Four materials were considered for construction of the truss: carbon fiber, aluminum, fiberglass, plywood, and PVC. The tensile strength, ease of construction, average price, density, and modulus of elasticity of each material was compared. Since the truss will be supporting the full weight of the rocket and will need to resist a strong bending moment, the 65

67 modulus of elasticity was the most important factor in choosing a material. The initial comparison of materials can be seen in Table 4.4. Decision Factors Materials Table 9.4: Truss Material Comparison Chart Tensile Strength (psi) Ease of Construction Average Price ($/lb) Density (lb/ft 3 ) Modulus of Elasticity (psi) Carbon Fiber Hard E+07 Aluminum Hard E+07 Fiberglass 3450 Hard E+07 Plywood 4500 Easy E+04 PVC 3053 Very Easy E+05 The next step in the trade study was scoring each material property. Each property was broken up into five ranges and the ranges were ranked from -5, with 5 being the most desirable. Table 4.0: Weighted scoring criteria for truss trade study Scoring Tensile Strength (psi) Ease of Construction Cost ($/lb) Density (lb/ft 3 ) Modulus of Elasticity <= 0000 Very Hard >.25 >200 <=E+04 2 <= Hard <=.25 <=200 <=E+05 3 <= Medium <=.00 <=50 <=E+06 4 <= Easy <=.75 <=00 <=2E+07 5 >40000 Very Easy <=.50 <=50 >2E+07 After determining the score of each material property, it was necessary to determine which properties were the most important to the project. Tensile strength was rated medium priority. The design of the truss will cause several members to be in tension, therefore the material chosen must have a high tensile strength. Ease of construction was chosen to be low priority because there is a large amount of time available to construct the truss. Average price was ranked medium priority. The truss is critical to the success of the entire project, therefore it was determined to be worthwhile to use a stronger, more expensive material. Density was also ranked medium priority. An electric motor will be lifting the truss into launch position after recovery of the payload, so a lighter truss will result in lower costs as a cheaper, less powerful motor can be used. The most important factor was determined to be the modulus of elasticity of the material. The truss will be subjected to a large bending moment while supporting the rocket in a horizontal position and while lifting the 66

68 rocket into the launch position. A material with a high modulus of elasticity will be required to ensure the success of the project. Factors Table 4.: AGSE Truss Material Trade Study Weight Carbon Fiber Aluminum Fiberglass Plywood PVC (-3) Tensile Strength (psi) Ease of Construction Average Price ($/lb) Density Modulus of Elasticity Raw Score Ranking As seen in Table 4.6, carbon fiber ranked first in the trade study and was chosen to be used for construction of the truss. It is a very strong and lightweight material that will easily support the rocket. However, if carbon fiber is unable to be used due to budgetary limitations, PVC will be used for construction. Section 4..3: Payload Retrieval Subsystem (PRS) The first major assignment that the AGSE must be able to complete concerns the capture and transportation of a small cylindrical payload from the ground to the inside of a launch vehicle that has been positioned horizontally. The cylindrical payload will be a sand-filled PVC pipe with dimensions of ¾ inches in diameter and 4.75 inches in length. The payload will weigh approximately 4 ounces, and the ends will be capped with domed PVC end caps. The payload will have no other components or mechanisms installed for capturing purposes, and it will be placed somewhere on the ground within reach of the robotic arm outside of the mold line of the rocket while the rocket is oriented horizontally. 67

69 The team knew from the beginning that a robotic arm with optical sensors would be used to complete the mission requirements. A robotic arm would offer multiple degrees of freedom to work with, reduce risks of mission failure, and present technical challenges to the team that would expand our knowledge of autonomous robotic technology in a Martian environment. In choosing the design of the robotic payload retrieval system, the team took into consideration the requirements that were presented in the Request for Proposal (RFP) as well as the risks involved in completing the mission and additional opportunities that using a robotic arm presented. Table presents the design criteria required in the RFP and additional capabilities that are important to the team. Table 4.7: Robotic arm design criteria Robotic Arm Design Criteria Baseline Requirements Must be able to reach both the ground and the payload bay of the launch vehicle Must be able to securely capture the payload from the ground Must transport the payload to the payload bay of the launch vehicle Must be able to insert the payload into the payload bay of the launch vehicle System must not rely on Earth-based operating technology Performance and Derived Requirements Cannot inhibit or disrupt any other AGSE subsystems All robotic arm components must be quickly reusable The arm must be modifiable to suit a variety of applications The robot must accept sensors and other controllable electrical equipment The system must be able to sense that the payload has been obtained The gripper must be large enough to hold the payload at multiple points of contact The gripper must have non-slip surfaces to prevent the payload from sliding out of grip The system must be simple to assemble The cost of the entire assembly must be at a reasonably good value The robot arm must have at least 4 degrees of freedom System must be able to withstand Martian environment Several options existed for the robot arm. The team had the option to either fabricate the arm entirely in-house or purchase a commercially available arm. After performing some online research for commercially available technology, it became quickly evident that it would be more advantageous for the team to successfully meet all baseline requirements and more by purchasing and modifying a commercially available robot arm rather than building one from scratch. There are many commercially available robotic arms and parts that have been optimized to be durable, 68

70 reliable, and manufactured to perform tasks similar to the requirements stated above. As the team has limited access to the materials and technology that is required to build such similarly advanced robotic technology, it was decided that a commercially available robotic arm should be purchased and modified to meet the baseline requirements and so much more. In choosing the best technology to integrate into the Project WALL-Eagle design, it was important that the system could be modified to fit the design and provide avenues to ensure mission success, hence the Performance and Derived Requirements presented in Table 4.7. The commercially available choices were narrowed down to three options, which are presented next. Option : Lynxmotion AL5D 4 Degrees of Freedom Robotic Arm Combo Kit (BotBoarduino) Some major points for what makes this selection a viable option are described as follows. This robot arm can be controlled via an Arduino board. It has 4 degrees of freedom and a reach length of 0.25 inches, which would be sufficient enough to move a small PVC tube from one nearby location to the rocket, but it the reach would limit the area in which the payload may be retrieved. The price is reasonable ($309), so it leaves plenty of room for potential modification and expenses elsewhere in the AGSE. The Lynxmotion website also provides all assembly/coding/cad information that would be required to operate and integrate this arm into the system. This site also offers sample code available for running all available types of Lynxmotion robot arms. This selection meets the baseline criteria and some of the other performance criteria as indicated by checkmarks below. 69

71 Robotic Arm Design Criteria Baseline Requirements Must be able to reach both the ground and the payload bay of the launch vehicle Must be able to securely capture the payload from the ground Must transport the payload to the payload bay of the launch vehicle Must be able to insert the payload into the payload bay of the launch vehicle System must not rely on Earth-based operating technology Performance and Derived Requirements Cannot inhibit or disrupt any other AGSE subsystems All robotic arm components must be quickly reusable The arm must be highly modifiable to suit a variety of applications The robot must accept sensors and other controllable electrical equipment The system must be able to sense that the payload has been obtained The gripper must be large enough to hold the payload at more than two points of contact The gripper must have non-slip surfaces to prevent the payload from sliding out of grip The system must be simple to assemble The cost of the entire assembly must be at a reasonably good value The robot arm must have at least 4 degrees of freedom System must be able to withstand Martian environment Although the Lynxmotion robot arm has the ability to meet all baseline requirements and some of the Performance and Derived Requirements, the extent for how well it meets those requirements is limited. For example, the length of the unit could become an issue if the payload were to be dropped and recaptured from an unknown position. Also, there are not many options to modify the device with advanced capabilities that would be more reliable and suitable to the design. Most importantly, better options exist that meet more requirements and provide higher performance standards than the Lynxmotion option. Option 2: RobotShop M00RAK V2 Modular Robotic Arm Kit (No Electronics) This arm can lift up to 500g at a full reach of 24 inches. Servos and electronics are sold separately, but the item description recommends servos/electronics to purchase for this robot arm, which may 70

72 all be purchased from the same site. Since the PVC tube will be 4 oz. (roughly 5g), an arm capable of lifting 500g is more than enough to easily lift and place the payload within the rocket. Its recommended controller is the same as the one for the Lynxmotion robot, so it has all of the same control benefits and limitations as that model. The price for this robot arm is $ While the cost is greater than that of the Lynxmotion option, the extended reach would be vital for its functionality and ability to complete the mission. The arm could potentially be extended if necessary. This model also does not come with a gripper, so a gripper would need to be purchased and installed separately. The primary advantage of this model is its customization abilities. Many different design configurations could be constructed and tested for optimal performance, but the cost of purchasing all of the required components and electronics separately increases dramatically (total costs would be several hundreds of dollars more). Robotic Arm Design Criteria Baseline Requirements Must be able to reach both the ground and the payload bay of the launch vehicle Must be able to securely capture the payload from the ground Must transport the payload to the payload bay of the launch vehicle Must be able to insert the payload into the payload bay of the launch vehicle System must not rely on Earth-based operating technology Performance and Derived Requirements Cannot inhibit or disrupt any other AGSE subsystems All robotic arm components must be quickly reusable The arm must be highly modifiable to suit a variety of applications The robot must accept sensors and other controllable electrical equipment The system must be able to sense that the payload has been obtained The gripper must be large enough to hold the payload at more than two points of contact The gripper must have non-slip surfaces to prevent the payload from sliding out of grip The system must be simple to assemble The cost of the entire assembly must be at a reasonably good value The robot arm must have at least 4 degrees of freedom System must be able to withstand Martian environment Although this option may be modified to meet all baseline and many of the additional requirements, the costs of all the required additional components would be too substantial for the team to acquire. 7

73 Option 3: CrustCrawler AX-2A Smart Robotic Arm The CrustCrawler AX-2A Smart Robotic Arm is loaded with features and customization capabilities. This arm can lift 2-3 lbs. and can reach a length of approximately inches with 5 degrees of freedom. Although the arm only needs to lift 4 ounces, the strength capabilities will enhance the speed at which the arm may complete its tasks without adding too much stress on the servos. The arm is completely compatible with any microcontroller, computer control system, and any programming languages (including MATLAB and LABVIEW. The frame is constructed from aluminum, so the structure is stable, reliable, and lasting, and in combination with the included dual-actuator servos at each joint, the kinematic accuracy of the arm measures from mm to 3mm. The arm measures position, voltage, current, and temperature feedback that can be utilized in a variety of experimental applications. Additionally, the sensor-engineered gripper is designed to accept IR detectors, cameras, pressure sensors, and more. The product includes all of the necessary hardware, software, electronics, and power supplies needed to operate the arm. The system comes completely disassembled so the arm may be built and modified to any specifications that are necessary. The system comes with computer programs, LABVIEW VI Packages, cables, and a CM-530 controller, which are all included in the price. The CrustCrawler AX-2A Smart Robotic Arm package costs $830.00, which is within the budget that the team allocated for purchasing a robotic arm. As can be seen below, the option would be capable of meeting every requirement listed in the design criteria. 72

74 Robotic Arm Design Criteria Baseline Requirements Must be able to reach both the ground and the payload bay of the launch vehicle Must be able to securely capture the payload from the ground Must transport the payload to the payload bay of the launch vehicle Must be able to insert the payload into the payload bay of the launch vehicle System must not rely on Earth-based operating technology Performance and Derived Requirements Cannot inhibit or disrupt any other AGSE subsystems All robotic arm components must be quickly reusable The arm must be highly modifiable to suit a variety of applications The robot must accept sensors and other controllable electrical equipment The system must be able to sense that the payload has been obtained The gripper must be large enough to hold the payload at more than two points of contact The gripper must have non-slip surfaces to prevent the payload from sliding out of grip The system must be simple to assemble The cost of the entire assembly must be at a reasonably good value The robot arm must have at least 4 degrees of freedom System must be able to withstand Martian environment Selection: Option 3 - CrustCrawler AX-2A Smart Robotic Arm The Crustcrawler AX-2A Smart Robotic Arm would present no predictable limitations for what it would be required and designed to accomplish. The team decided to purchase the CrustCrawler AX-2A Smart Robotic Arm to integrate into the AGSE system. Although this arm is the most expensive of the options that were considered, the features, tools, and components that are included in this purchase are justified by the value that the package offers for the team. Everything that the team needs to operate and integrate the arm is included in the purchase, and the abilities to easily modify the arm with sensors, detectors, and grippers, along with its strength and reach, will ensure that the AGSE subsystem will be able to complete its requirements with minimal risks. The CrustCrawler AX-2A Smart Robot Arm that the team has chosen to integrate into Project WALL- Eagle is presented in Figure 4.4 below. 73

75 Figure 4.4: CrustCrawler AX-2A Smart Robot Arm Final Robot Arm Design and Function The robotic arm will be constructed as intended by the manufacturer with the exception of three modifications that the team has developed in order to meet all requirements described in the statement of work and those described in Table 4.7. The robot arm has already been purchased, assembled, and measured, excluding the design modifications. More details on how the measurements influenced the design modifications are described in the paragraphs that follow. Modification #: Extending its reach The unmodified robot arm has a maximum extended reach of 20 inches with the gripper open and 22 inches with the gripper closed. However, those measurements include the base of the robotic arm and the first joint, which vertically fixed and cannot be used to extend the reach of the arm toward the ground from the box. In addition, the arm will be used to reach over the horizontally positioned launch vehicle in order to close the payload bay hatch. Therefore, to ensure that the arm can reach a slightly more than reasonable area of the ground from its mount, a 3-D printed 74

76 extension to the length of the arm has been developed to ensure that the arm can reach the ground at a distance of a least foot from the box. The custom extension piece will be made of ABS plastic, which will certainly be lightweight and durable enough to withstand the gravitational moments created by the weight of the arm and the payload. This modification will be tested extensively and verified by successfully capturing and transporting the mock payload from different distances to measure the modified arm s strength capabilities at farther-away positions than the unmodified arm was designed for. Modification #2: Modified gripper The gripper that was developed as the default gripper for the unmodified robotic arm only had two points of contact with the payload. When the unmodified version was built, the gripper was able to lift small 2-8 ounce objects with ease. However, some of the heavier objects rotated in the plane perpendicular to the plane in which the gripping device opened and closed. This raised concern for the gripper s ability to hold a cylindrical payload without rotating about the two points of contact. The gripper was tested on a few cylindrical objects around the room (including an unopened soda can). The arm was able to lift the objects and move them, but the objects would rotate while in the gripper. To prevent rotation the team decided to choose a more customizable gripper that would offer more points of contact to the payload. Figure 4.5 below presents the preliminary modification for the modified gripper. It will provide three points of contact, and the gripping surfaces will have nonslip padding installed, preventing any sliding inside the grip due to little friction. 75

77 Figure 4.7: The modified three-point gripper in progress Modification #3: Infrared Sensing Technology In order to ensure that the requirement to capture the payload and insert it into the launch vehicle is successfully completed, the team decided to create a system that could autonomously detect and repeat search and capture procedures in case the gripper were to drop or miss the payload before reaching the payload bay. To do this, the team has begun intense development of a program that will incorporate the use of infrared emitters and detectors installed on the gripper that will be able to sense the payload on the ground and (re)capture it. The robot arm will essentially scan an area of the ground within reach to locate the payload, orientate the gripper to the orientation of the payload, and capture the payload from any position in that area. Pressure sensors installed on the gripper will be able to send feedback to the controller to signal whether or not the payload has successfully been captured. Section 4..4: Launch Vehicle Elevation Subsystem (LVE) 76

78 The Launch Vehicle Elevation Subsystem (LVE) is comprised of a winching mechanism that will be connected to pulleys, a tower, and the top of the launch rail in order to slowly, safely, and effectively lift the launch rail and the rocket to a vertical position at 5 degrees off the vertical. Figure 4 shows the setup of the winch mounted to the front of the AGSE and the tower that it will be support the pulley for the cables. A tower will be mounted above the winch motor to hoist the cable through a pulley that will route the cable to the top of the launch rail. The tower will have a height of 0 ft in order to ensure that the launch rail will be easily lifted and locked into position when raised to the vertical. Figure 4.8: The winch setup mounted on the base of the AGSE At the top of the launch rail, a lightweight aluminum mount will be installed to connect the winching cables to the top of the launch rail. Figure 4.7 presents a schematic for how the mount will be mounted to the top of the truss. 77

79 Figure 4.9: The winching cable mount at the top of the launch rail/truss The mount will be connected through the webs of the double-t carbon fiber truss as shown in the following figure. 78

80 Figure 4.0: Mounting mechanism for the winching cable supports The truss that is used to lift the slotted T-rail and rocket will be made from carbon fiber, and it will have a double-t beam cross section to provide strength to the structure over a 9.6 ft length. The truss will have dimensions according to the following. Figure 4.: Dimensions of the carbon fiber truss Mounted to the center of the carbon fiber truss will be a T-slotted launch rail which will serve as the guiding rail for the launch of the rocket. Its cross section and dimensions are presented in the following figure. 79

81 Figure 4.2: Dimensions of T-slotted launch rails for launching the rocket Section 4..5: Automated Charge Insertion System (ACI) The Automated Charge Insertion System (ACI) is the third autonomous subsystem of the AGSE that will objectively insert the igniter into the motor. The motor has a length of 9.6 inches, and therefore, the ACI must be able to accommodate this length. It is necessary to autonomously insert the igniter into the rocket once the rocket is in the vertical position. A retractable telescopic linear actuator igniter insertion system is the most straight forward and effective way to accomplish this task. A toothed plastic rod with the igniter attached to it will be driven by a DC electric motor with a toothed gearhead such that the igniter is driven into the rocket engine. Such systems were frequently used for retractable car antennas, and the technology is commercially available. The team is manufacturing a custom retractable linear actuator that will be mounted beneath the blast plate near the base of the launch rail approximately 8 inches below the base of the rocket. A schematic of how the linear actuator will be positioned on the blast plate relative to the rocket in its collapsed position may be viewed in Figure 4.. In Figure 4.2, a schematic of the retractable actuator may be seen as it has been fully extended into the motor. 80

82 Figure 4.3: Igniter inserter in the collapsed position Figure 4.4: Igniter inserter fully extended into the motor An e-match will be attached to the tip of the actuator, and the e-match will be connected to the launch controller. The launch controller will only be enabled when all systems have completed their sequences. Once the igniter is inserted, the system will pause. Once all sequences have been successfully completed, inspected and verified, a green light will light up on the primary AGSE controller. When this light turns on, the launch controller may be activated once all recovery avionics have been enabled. When the rocket is inspected and deemed to be safe for flight, the launch sequence will be activated, the igniter will ignite, and the rocket will fly. A schematic for the motor that will drive the extendable dowel up into the motor may be seen in Figure 4.3 below. 8

83 Figure 4.5: Schematic for igniter insertion actuator The dowel will be comprised of three six inch sections and a 7.6 inch section that will extend telescopically to the top of the motor. The inner most piece will be the 7.6 inch section with the igniter attached to the tip of it. The dowel will extend through a hole in the blast plate while the actuator and electronics will be below the blast plate. This will ensure that the actuator is protected from the burning motor at ignition and that the igniter will be in its correct position. This igniter insertion system is designed such that the igniter will be securely placed in the motor and will be thin enough to ensure that the motor is not choked. In addition, the residue from the burning motor raises the risk of tarnishing the steel and causing the unit to become unusable. In order to prevent the risk of insertion failure, the dowel will be easily replaceable in the event that the burning motor damages the dowel during launch. Section 4.2: AGSE Electronics Many electronic components are necessary for the AGSE system to function correctly. Each AGSE function will be supervised by the master microcontroller which will be an Arduino Mega. Having 82

84 one master microcontroller minimizes risk by allowing the system to be aware of its status at any given point in time. The master microcontroller will have a dedicated internal power supply (a 9V battery) while other AGSE equipment will be powered by a 2V car or motorcycle battery. The AGSE sequence begins at the launch controller. The launch operator first turns on the master switch and then presses the start button on the launch controller. When the start button is pressed, the master microcontroller sends a command to the microcontroller on the robotic arm to locate the payload. The robotic arm then autonomously locates the payload using an infrared detector/emitter pair, picks it up with a grabber claw, deposits the payload in the rocket s payload bay, and moves clear of the rocket. When this is accomplished, the robotic arm s microcontroller sends a signal to the master microcontroller indicating that the payload has been integrated into the rocket. When the master microcontroller receives this signal, it activates a servo that pulls a pin that releases the payload bay door and allows the torsion springs on the hinges of the door to close the payload bay with enough force for the door to latch. Once this happens, the master microcontroller will drive a continuously rotating servo that will slowly wrap a cable around an extended servo shaft. This cable will be wrapped over a pulley at the top of a winch tower, and the other end of the cable will be attached to the tip of the launch rail. Because the cable will be attached so far from the pivot point, the servo will provide enough torque to retract the cable and raise the launch rail so that it is 5 degrees from the vertical. When the launch rail reaches the desired position, the blast pad will depress a push-button switch which will indicate to the master microcontroller that the desired position has been reached and that the servo should be set to a minimum torque value (to keep some tension on the cable). After this, the master microcontroller will drive the linear actuator to insert the igniter into the rocket motor. The system will then pause and wait for the launch command. When the safety switch is turned off and the launch button is pressed, the master microcontroller will actuate a relay that will drive a large current through the igniter and activate the solid rocket motor. Section 4.2.: Launch Controller and Function The launch controller is the handheld unit that is utilized to start, pause, and reset the AGSE sequence and is also used to command the rocket to launch once the sequence is complete. Once the AGSE sequence begins, the only component that can be interacted with is the launch controller. The launch controller will consist of an orange LED indicator, a green LED indicator, two push- 83

85 button switches, and two rocker switches. The first rocker switch will serve as a master switch that controls the flow of electricity from the 2V battery. The positive wire from the battery will be routed through this master switch such that power to all AGSE components can be directly activated or disabled from the launch controller. One of the push-button switches will serve as the start/pause/reset/button. This button is pressed to start the AGSE sequence in the beginning and is pressed again to pause the sequence. When the AGSE sequence is running, the orange LED will blink at a frequency of Hz. When paused, the sequence can be resumed by pressing the button once again. When the sequence is paused this LED will glow solid orange. If the system is paused and the button is depressed for more than 5 seconds, the master microcontroller will assume a reset of the entire AGSE sequence and proceed as such. The second push-button is the launch button. When all AGSE processes have been verifiably completed, the second rocker (the launch safety switch) will be set to the ON position, and the launch button will be activated. The green LED will turn on to indicate that all systems are ready for launch. The Launch Control Officer (LCO) will then be able to launch the rocket by pressing the launch button. A total of six wires will run from the AGSE system to the launch controller. One wire will be connected to the system ground, one will be the positive voltage line from the battery, two wires will be needed for the LEDs, one will be needed for the start/pause/reset button, and one for the launch button. The electrical schematic for the launch controller is shown below in Figure 4.5. Note that all AGSE components will be connected to the same ground voltage. 84

86 Figure 4.6: Electrical schematic for the launch controller Section 4.2.2: Master Microcontroller An Arduino Mega will be used as the AGSE master microcontroller and is presented in Figure 4.6 below. The Arduino was selected because of its low cost and extensive open source libraries. The experience that some AGSE team members had working with Arduinos also made the Arduino Mega the preferable microcontroller. 85

87 Figure 4.7: Arduino Mega microcontroller for the AGSE The Arduino will be connected to the launch controller, launch rail elevation system, robotic arm microcontroller, and payload bay hatch closing mechanism. The diagram below illustrates how the Arduino Mega will be connected to these various components. An Arduino Uno is used in the diagram in place of the robotic arm microcontroller. The LEDs connected to the Uno in the diagram represent the IR detector and emitter attached to the robot arm microcontroller. This design depends largely on the electronic components contained within the microcontrollers and servos, eliminating the need to solder large numbers of components on a circuit board and reducing risk. An electrical schematic for the entire AGSE may be viewed in Figure 4.7 below. 86

88 Figure 4.8: Electrical schematic for entire AGSE Button presses are detected by measuring the voltage at a certain point in the button circuit. For both buttons, the orange wires allow the Arduino to measure the voltage right after the resistor. If the button is depressed, the voltage will be high (logic high). Otherwise the voltage will be zero (logic low). Servos are provided a positive voltage line and a ground line, and the speed of the servos is controlled by a pulse wave modulated (pwm) signal sent by the master microcontroller. A relay is used to provide a large amount of power to the igniter. When the Arduino actuates the relay, power flows directly from the battery to the igniter. A single digital input/output port of the master microcontroller is connected to a digital input/output port on the robotic arm microcontroller (the Arduino Uno Figure 4.7). The master pin is set to output mode with a low 87

89 voltage and the arm microcontroller is set to input mode. When it is time for the arm to locate and integrate the payload, the master microcontroller sends a high voltage signal to the arm microcontroller via the wire and immediately after, it switches the pin to input mode. When the arm microcontroller receives this signal, it begins the payload location and integration sequence. After this sequence is done, it sends a signal through the wire to the master microcontroller. When the signal is received, the master proceeds to close the payload bay hatch, erect the rocket, insert the igniter, and wait for the launch command. The operator can pause this sequence at any time by pressing the pause button. If the pause button is pressed during the payload retrieval phase, the master microcontroller will send another signal to the arm microcontroller, and the robotic arm will pause wherever it is. The flow chart for the master microcontroller is as follows in Figure 4. 8 on the following page. 88

90 Figure 4. 9: AGSE electrical flowchart for mission sequences via the master microcontroller 89

91 Section 4.2.3: Robotic Arm Location and Integration Algorithm When it receives the signal from the master microcontroller, the robotic arm will proceed to locate the payload and deposit it in the payload bay of the rocket. The simplest way to do this would be to place the payload in a predetermined location and have the robotic arm to travel to that location and pick up the payload. However an algorithm to scan the area and find and pick up the payload will be implemented for added redundancy. This algorithm will prove useful if the payload is in an unexpected location or if the payload is dropped. An infrared emitter/detector pair will be mounted above the claw of the robotic arm. The robotic arm will move over the target area in a grid pattern in increments of.5 in. Because the payload is.75 in thick, moving no more than.5 in. at a time will ensure that the arm will pass over the payload at least once no matter what as long as the payload is in the target area. After each movement, the IR emitter will flash and the IR detector will record the level of IR reflection from this flash. The white PVC pipe that the payload is made of will reflect more light than the darker ground, so when a large amount of IR reflection is received, the claw of the robotic arm must be over the payload. By sweeping the area and recording the IR measurement at each point in the scan, the location and orientation of the PVC pipe can be determined. Scan points with IR reflection above a certain threshold will be assumed Figure 4.20: Infrared sensors binary mapping of payload to be above the PVC pipe, while points with IR reflection below this threshold will be assumed to be over the ground. By determining the two points that are furthest apart and drawing a line 90

92 between these points, the location of the longitudinal axis of the PVC pipe will be estimated. The arm will then move the claw to the midpoint of this line and orient the claw so that the grip axis is parallel to the estimated longitudinal axis of the payload. The claw will then descend and grab the payload. Note that while the two furthest points may not lie exactly along the longitudinal axis, the amount of error that would occur would be insignificant and would not prevent the arm from being able to properly grab the pipe. Once the payload has been acquired, the arm will deposit it in the payload bay of the rocket and then move out of the way of the rocket. If the payload is dropped at any point during this procedure, the arm will re-execute the searching algorithm and attempt to recover the payload. Figure 4.9 above shows the binary map of locations over the PVC pipe and the estimated longitudinal axis of the pipe for a sample payload. Section 4.3: AGSE/Payload Integration In order for the payload to be properly integrated within the launch vehicle, a dedicated payload bay was designed. The payload bay was designed to ensure that the payload would be secured until final delivery. As outlined in Section in the Statement of Work, the payload is a PVC tubing filled with sand that will weigh 4oz with a length of 4.75 inches and a diameter of.75 inches. The payloads dimensions served as the driving design parameter of the payload bay. Another design consideration that was vital to consider was the overall 5 inch diameter of the rocket. Furthermore, ensuring ease of integration and compatibility of all components involved was another driving design consideration. In order to accomplish this step, an integration plan was developed to ensure that all mission requirements were met while still ensuring a simple integration process. Section 4.3.: Integration Plan All systems are controlled by a master microcontroller in the form of an Arduino Mega. This will serve as a safety precaution by allowing the system to be aware and in control of all systems of the AGSE in the case of component failure. This is important to consider in order to ensure that in the event of any errors, all systems can be immediately shut down to prevent any catastrophic events. 9

93 Launch Rail and Truss The Launch Rail and Truss System is one of the essential components in ensuring mission success. It consists of two supporting blocks designed to house all the electrical and mechanical components needed to raise and launch the rocket. In addition, the supporting blocks will house the robotic arm and the corresponding electrical system. Therefore all systems must not only be easily integrated and compatible, but they must also be able to be shielded from any potential damage from the launch and the environment. The two support blocks are connected via trusses that will serve as structural supports but also to separate the launch platform from the robotic arm system. It will consist of the robotic and main support structures. Robotic Support Structure The robotic support structure (the smaller base structure) houses the robotic arm s microcontroller and all supporting electronic equipment including a power source in the form of a 9V battery. The robotic arm is mounted on a tiered surface of the robotic support structure that allows it to easily transport the payload from the environmental surface to the payload bay. As a secondary function, the structure also functions as a support for the launch rail and vehicle while is oriented in the horizontal position. This further allows the robotic arm easy access to the payload bay by positioning the payload bay in an optimal position in relation to the robotic arm further increasing the compatibility and ease of payload integration. As stated above the robotic arm s microcontroller is directly linked to the master microcontroller to ensure proper and safe procedures. Additionally, the robotic arm not only transfers the payload, but also securely closes the payload access hatch when the payload is secure within the payload bay. Main Support Structure The main support structure serves a multitude of functions. The main support structure (the larger, U-shaped base) houses the launch rail erection system, the ignitor insertion system, and serve as the launch pad for the launch vehicle. In order to ensure that all these systems were compatible, an easily implemented integration process was implemented. This consisted of linking all systems to master microcontroller to ensure that all systems operated correctly and at precisely the correct time and ensuring ample space for all components and their supporting systems Launch Vehicle Elevation System 92

94 The LVE consists of motor that will raise the launch rail the desired position. The launch rail will be counterweighted to ensure that the rail is secure. A mechanical locking system also helps to ensure that the launch rail is securely in place. In order to achieve to desired inclination for launch, a blast plate is welded onto the base of the launch rail at precisely 5 degrees from the relative to the horizontal. As the motor raises the LVE, the blast plate will come into contact with the top surface of the main support structure until it is completely flush with the surface. This ensures that the rocket is at the proper inclination. Furthermore, the blast plate also protects the internal components from any adverse effects from the launch. Automated Charge Insertion System The ACI consists of a power antenna motor unit that utilizers a toothed insertion system through a series of gears that will insert the ignitor into the rocket motor via a small hole in the blast plate. Again to ensure ease of integration, the ACI is controlled by the master microcontroller to initiate insertion only after the rocket is in the proper position and inclination. Robotic Arm As previously described in the robotic support structure section, the robotic arm has its own microcontroller given the complexity of the system. However, this microcontroller is directly linked to the master microcontroller for safety purposes and to ensure that all systems function as intended. The robotic arm utilizes several algorithms for a variety of environmental situations based on if the payload is located at a predetermined location or not. The robotic arm utilizes infrared emitters and detectors mounted above the claw to detect the payload and transport the payload into the payload bay. The arm will have 5-6 degrees of freedom and will have the ability to be directly controlled from the master microcontroller in order to properly execute the mission. Payload Access Hatch/Payload Bay The payload access hatch is cut out of the airframe and is hinged in order to allow for the payload bay to be easily accessible. The usage of a microcontroller and a microservo to control the hatch was determined to be an unnecessary element because of the additional electronics systems that would be required and the restrictive volume of the payload bay. This would have hindered proper payload integration. Therefore, a completely mechanical system was devised in order to eliminate the need for a dedicated electrical system and allow for easier integration. The payload bay is filled with a polyurethane spray foam that serves to secure the payload along the center axis of the rocket 93

95 to ensure stability during flight. This is a foolproof system that is easily accomplished and assembled that allows for a proper and efficient payload integration. Furthermore, the payload bay hatch will be closed by the robotic arm using a preprogramed algorithm. In order to ensure that the hatch is securely closed and the payload is firmly situated along the center axis of the rocket, a mechanical lock will be lock the hatch into place. Section 4.3.2: Compatibility of Elements and Simplicity of Integration Procedures All systems were checked for compatibility. This included ensuring that all electrical systems functioned correctly through the master microcontroller. Furthermore, systems were checked to ensure that they would work in sync with one another and that individual systems could be independently controlled while others functioned for safety purposes. Systems were checked to ensure that there would not be preemptive systems startups. Additionally, to ensure simplicity of integration, all electrical systems were evaluated for their importance and eliminated if necessary in order to simplify the overall design of the system. This was the driving design decision for utilizing a purely mechanical system for the payload access hatch. Furthermore, eliminating nonessential electrical systems reduced the overall cost of the AGSE which was another driving consideration in the design. All systems were also designed around the dimensional considerations of the rocket, the payload bay, and the supporting structures. This was essential in order to ensure that not only would all the systems would function in coherence with one another, but also simply fit within their given areas. All components of each individual subsystem is easily assembled and disassembled for efficient transportation and construction. This was accomplished by once again eliminating any component that was not determined to be necessary and introducing a modular system that allowed for different designs to be tested in the initial phases of the design process. This modular system was essential in determining the most efficient and effective method to ensure proper payload integration. Section 4.4: Failure Modes and Mitigation Plans The AGSE system relies on the coordinated functioning of numerous subsystems. As such there are many ways in which this system can fail. Possible causes of failure include: 94

96 . Insufficient power supplied to AGSE components- Power must be supplied to the master microcontroller, robotic arm, payload sensors, launch rail erection system, igniter insertion system, and launch controller. If any one of these components is unable to receive power, mission failure will result. Insufficient power could be caused by a failure to set up power cables correctly or power supplies that are too small. This failure mode can be mitigated by designing robust, adequate capacity power supplies that are easy to set up and are connected to the AGSE components in a sturdy manner. Additionally, AGSE components should be able to resume operations and complete the AGSE sequence in the event of a power surge or temporary power failure. 2. Programming errors in AGSE master microcontroller- A single master microcontroller will be utilized to drive all the AGSE processes. As such, it is critical that this microcontroller function correctly. A programming error in the master microcontroller could cause the entire system to fail. For example, if the payload detection algorithm is faulty, the arm may not properly grab the payload. Programming errors can be mitigated by using a debugger, running numerous test cases on the code, having multiple people review the code for errors, and running the entire system numerous times to ensure adequate functionality. 3. System takes too long to complete sequence- The AGSE must integrate the payload, raise the rocket, and have the rocket ready for launch within 0 minutes. If the system is unable to do this within 0 minutes, mission failure will occur. Each component of the AGSE system should be designed such that the entire system takes less than 5 minutes to complete the full sequence under normal operating conditions. A failure of the robotic arm to locate and grab the payload on the first attempt is the most likely cause for delay. Low battery levels may also make the system operate more slowly than intended. For these reasons, the system must be thoroughly tested and designed to complete operations in significantly less than the maximum allotted time. 4. AGSE components not synchronized properly- AGSE components must be properly synchronized and must operate in the correct order. For example, the hatch cannot be closed until the payload is integrated into the rocket. To simplify synchronization, a single microcontroller will be utilized to drive all processes. This will minimize the need for communication among different microcontrollers. Because the AGSE sequence is fairly linear (only one step is completed at a time), it should be fairly simple for an Arduino or Raspberry Pie microcontroller to control the entire sequence. The master microcontroller will not begin a process until the previous required process is complete. This will ensure that processes are executed in the correct order. 5. AGSE equipment not set up properly- Failure to properly set up the ground equipment on the day of the launch could lead to AGSE failure. To mitigate this, AGSE equipment should be designed in a robust, modular manner such that it is easy to set up and difficult to break. Parts should fit together smoothly and the need for precise adjustment or arrangement should be minimized. Parts should lock together in a manner that leaves little room for human error. Additionally, AGSE team members should practice setting 95

97 up the system in a precise and timely manner to minimize the chances of assembly errors on the day of the competition. 6. System does not respond to user input- The AGSE operator will be in charge of issuing the stop, pause, and reset commands to the system. If the system does not respond to these commands, mission failure will occur. The controller shall be hooked up to the AGSE master microcontroller in a robust manner and must be thoroughly tested to ensure that operator commands will be received. 7. AGSE equipment damaged by environmental factors- If there is drizzle or high humidity on the day of the launch or during testing days, AGSE electronic equipment may be damaged. To the extent possible, sensitive electronic equipment should be covered or shielded such that environmental impacts will be minimal. Integrating a significant portion of the electronic equipment into the boxes that will constitute the AGSE platform will minimize the probability of environmental damage. Section 4.5: Plans for Manufacturing and Assembly All structural material for the AGSE and payload has been acquired. Plywood was the decided material to make up the two boxes. The plywood will be cut to the dimensions of the boxes and assembled by the last week of January. Aluminum pipes will act as the support and connection between the two boxes. The aluminum pipes will be attached to the box with custom fitters, which will be 3-D printed to link the aluminum pipes to the plywood boxes and to each other. Assembly of the truss is underway. The truss will be a flat steel plate with rails welded to the bottom of the plate. Once the rails are welded, the blast plate can be attached to the base in the designed position. Alignment of the blast plate is crucial, and will have to be measured and welded with minimal error. Measurement before and after welding the blast plate is mandatory to insure that the angle has not varied to an insufficient degree. The ACI equipment and materials have been acquired and will be constructed during the third week of January. The optimal design will function similar to a car antenna, and will act as the igniter inserter. If the antenna completes all of the required tests, it will be mounted under the blast plate and will slide through a small hole in the plate until it reaches the top of the inside of the motor. 96

98 Testing of optimal foam is underway. Once the desired properties are established for the foam, the foam will be applied to the inside of the payload bay. The payload bay will then be carved to the designed slope, insuring that the payload will rest in the center of the airframe. Applying foam to the door will allow the payload to be secure within the payload bay, and will prohibit erratic motion of the payload. Section 4.6: Precision of Instrumentation, Repeatability of Measurement The AGSE will include various instruments that will aid in the capture and securing of the payload, along with the erection and arming of the rocket. Table 4.7: Infrared sensors binary mapping of payload Instrument Model Operating Range Precision IR emitter and detector Robotic Arm Truss and blast plate Master Microcontroll er N/A CrustCrawler AX-2A N/A Arduino Mega Igniter inserter N/A Wavelength of nm Forward voltage 800mV 5kV Lift 2-3 lbs Reach of inches 5 degrees of freedom 85 degrees motion Support 80 lb Extensive source library 54 digital input/output pins 6 analog pins Ocular extension of 3 ft Replaceable mast.40% 3 mm 3 5 mm 0.25% - 2 mm These instruments have been chosen do to their commercial success. They have proven their repeatability within a reasonable margin of error, and the relevant data is transcribed in Table

99 Section 4.7: AGSE Concept Features and Definition The design of the AGSE/ Payload concept centered on several key design considerations. First and foremost were to meet the mission requirements as outlined in the statement of work. Other design considerations included the dimensions of the payload, dimensional restrictions in the manufacturing capability of the rocket, safety procedures, cost, and ease of integration among other factors. All of these factors had to be individually and collectively examined in order to determine the most efficient and effective method of ensuring mission success. Therefore, since there was bound to be several iterations of the design, it was imperative that the design be modular so that individual systems of the AGSE could be replaced, analyzed, and tested separately and function as a component of the entire system as a whole. Critical concepts and features include the ability of the entire system to be controlled by one master microcontroller. This serves as a safety feature while also allowing all systems to report to a central command controller that will allow for ease of implementation and usage, while allowing for isolating any arising issues to a specific subsystem. Therefore the master microcontroller is the most integral part of the entire AGSE system. Further essential features include the robotic arm/claw that will allow for the transport of the payload from the environment into the payload bay along with the corresponding electrical systems will be critical to the success of the mission. Infrared sensors are critical in allowing the robotic arm to detect the payload. Properly written algorithms to control the robotic arm and other systems will also be essential mission success. Multiple algorithms will be included in order to allow adaptability to the robotic arm depending on the environmental conditions and payload positioning. The launch rail lifting/erecting system also is an integral component of the AGSE system fulfilling a mission requirement while additionally providing a methodology of launching rockets autonomously from incredibly long distance. The simple mechanical design of many of the subcomponents of the system allow for easier integration, safety precaution against electronic failure, and cost reduction. The same design process was applied while testing systems for the payload access hatch and the payload bay. Using a simple hinged system, while a polyurethane foam core would secure the 98

100 payload along the center axis of the rocket, eliminated the need for any complicated and expensive electrical systems by simply utilizing a mechanical lock to secure the hatch and using an existing electronic system. The robotic arm, to close the hatch. This again allows for easier payload integration, eliminates any possible failure from electrical systems, and reduces the overall cost of the system. Section 4.7.: Creativity and Originality Project WALL-Eagle is designed to be a lightweight, easy-to-assemble system that would be conceptually easy and cost-effective to transport to Mars and be operable remotely. Ideally, the system could be light enough and easily assembled and disassembled. Although it seems counterintuitive that the system is lightweight when it needs to be able to lift a large rocket, this design would not need extra mass to be transported with it to weigh down the system. The idea is that a variety of mass elements may be deposited into the empty boxes to weight down the system. Also, if the boxes are large enough in scale, all sorts of important payloads may be collected, stored, and protected inside the boxes until they are ready to be transported or used up. Utilizing infrared emitters and detectors in order to determine the location of the payload is a unique approach to the mission requirement. By using infrared systems, the detection system provides an alternative when compared to the more common camera system. The simplicity of the overall design is unique when compared to other systems. Eliminating other electronic systems proves to be more cost effective than other designs which is other design feature. Our design is perhaps one of the few designs that truly considered cost and cost reduction as one of the primary design considerations allowing for our design to be widely accessible and economically despite any changes in budget. Section 4.7.2: Significance of Design Relatively low cost and proven reliability are components that greatly add to the significance of this project. Proven reliability was evident in the design phase as one of the considerations was choosing a robotic arm. By selecting a already proven robotic arm, it eliminated the need for the expensive and time consuming process of developing a completely new robotic arm, further reducing cost which was a significant factor in the overall design of the AGSE. Additionally, the 99

101 system was design to be as simple as possible so that it could be easily maintained while enduring the Martian environment. Furthermore, being able to operate in such an inhospitable environment would raise any cost, especially in shielding and protecting electronic systems, which is why the design was simplified to discard any electrical systems that were not essential. Section 4.7.3: Level of Challenge The level of challenge is high as there are a plethora of design constraints the made some more creative and outlandish designs to be dismissed despite their promise. Another major technical challenge is being able to design a system that would not only survive the harsh Martian atmosphere, but also the even harsher transit to arrive at Mars. Shielding from radiation and any other possible adverse effects had to be taken into consideration when designing the AGSE. This was another factor in trying to simplify the need for any shielding necessary to protect electronic systems as that would add additional cost which was unacceptable under the design considerations. Further challenges include designing the system to be reliable and serviceable remotely. More specific challenges include ensuring the ignition insertion system (ISS) operates correctly as the system must insert the ignitor through a small hole or making sure that any one of the systems is available for commands from the master microcontroller. Section 4.8: Scientific Value of Design Section 4.8.: AGSE Objectives AGSE Objectives System The entire system must secure the payload inside the rocket and have the rocket ready to launch in under 0 minutes. Launch Vehicle Elevation Subsystem (LVE) Launch pad will support the entire weight of the AGSE and rocket. House and protect important electronics and motors from environment. Raise rocket from horizontal to launch position 5 degrees from the vertical. 00

102 Support and guide rocket during launch to allow stable flight. Capable of lifting launch vehicle weighing 30 lbs. Payload Retrieval Subsystem (PRS) Scan and detect payload location on ground. Capture the payload. Deliver the payload to the payload bay in the launch vehicle. Return to resting position. Must be able to reacquire payload if dropped by the arm. Automated Charge Insertion System (ACI) Must move the igniter into the motor once rocket is in launch position. Must move the igniter into the motor until it reaches the top of the fuel grain. Will stop moving the igniter once it reaches top of fuel grain. Must withstand exhaust from launch vehicle. Must be reusable. Section 4.8.2: AGSE Success Criteria AGSE Success Criteria System Payload secured and rocket in launch position in under 0 minutes. Launch Vehicle Elevation Subsystem (LVE) Vehicle is launched without failures in the supporting structure. After vehicle is launched, all components housed inside launch pad are undamaged. Payload Retrieval Subsystem (PRS) Payload located and captured by the arm. Payload placed inside the payload bay of launch vehicle. Payload is in the correct orientation inside launch vehicle. Automated Charge Insertion System (ACI) Motor stops when igniter reaches top of fuel grain. Rocket launches successfully. Igniter Insertion System is undamaged and ready to be used again. Section 4.8.3: AGSE Requirement Testing Verification Table 0

103 The below table lists the requirements set forth in the NASA SL handbook in the Statement of Work (SOW). Everything mentioned in the table is in reference to the Autonomous Ground Support Equipment (AGSE). The st Requirement column gives the reference to the specific requirement listed in the SOW. The 2nd Means to Meet column gives a brief description of what will be done to make sure the requirement is met. The 3rd Verification Method column will provide the method to verify that the incorporated system functions as desired. The options of this column are: Inspection, Analysis, Testing, and Demonstration. Requirement Means to Meet Verification method Status AGSE AGSE will be fully autonomous AGSE Any pressure vessel used in the AGSE will follow all regulations set by requirement.2 in the Vehicle Requirements section AGSE AGSE equipment must be able to operate in a Martian environment AGSE Sensors that rely on Earth s magnetic field cannot be used AGSE Ultrasonic or other sound-based sensors cannot be used AGSE Earth-based or Earth orbit-based radio aids cannot be used The AGSE has been developed in a way such that human interaction is not necessary for tasks to be completed Requirement has been reviewed and it has been determined a pressurized vessel is not necessary Earth-based reliant systems are not being used Sensors used do not rely on earth s magnetic field. AGSE does not rely on sound-based sensors. AGSE does not rely on Earth-based radio aids. Testing and Demonstration Pressurized vessel is not being used Design methods and Construction System was designed in such a way that sensors used are independent of Earth s magnetic field Sound based sensors are not being used Earth-based and Earth- Orbit based radio aids are not used Incomplete Complete Complete Complete Complete Complete 02

104 AGSE Open circuit pneumatics cannot be used AGSE Air breathing systems cannot be used AGSE Payload bay must be of proper dimensions and must be able to seal properly AGSE The payload may be placed anywhere in the launch area for insertion, as long as it is outside the mold line of the launch vehicle when placed in the horizontal position of the AGSE AGSE The payload container must utilize a parachute for recovery and contain a GPS or radio locator AGSE does not utilize open circuit pneumatics AGSE does not utilize air breathing systems Payload bay has been designed to meet payload dimension requirements and the hatch has been designed to fully seal upon closing Payload was placed away from the mold line of the launch vehicle s horizontal position. Payload has a parachute for descent and GPS Open circuit pneumatics are not used Air breathing systems are not used A dummy payload was made to official dimensions and used for testing. After testing the door was checked for proper closure and lock. Proper spacing was measured and then testing of the system was conducted Parachutes were constructed for the payload bay and a GPS was placed inside Complete Complete Complete Incomplete Complete 03

105 AGSE Each team will be given 0 minutes to autonomously capture, place, and seal the payload within their rocket, and erect the rocket to a vertical launch position five degrees off vertical. Insertion of igniter and activation for launch are also included in this time. AGSE Each team must provide the specified switches and indicators for their AGSE to be used by the LCO/RCO AGSE A master switch to power all parts of the AGSE. The switch must be easily accessible and hardwired to the AGSE AGSE A pause switch to temporarily terminate all actions performed by the AGSE. The switch must be easily accessible and hardwired to the AGSE. AGSE has been developed in such a way that all necessary procedures can be completed in at least 0 minutes. Proper switches and indicators have been placed in the system. A master power switch has been placed into the system A pause switch has been placed in the system that can terminate all actions for the AGSE. System tests were conducted and timed Guidelines were reviewed and switches and indicators are in place The switch was tested and used to power the system on and off The pause switch was tested during AGSE operations to verify system does stop Incomplete Incomplete Incomplete Incomplete 04

106 AGSE A safety light that indicates that the AGSE power is turned on. The light must be amber/orange in color. It will flash at a frequency of Hz when the AGSE is powered on, and will be solid in color when the AGSE is paused while power is still supplied. AGSE An all systems go light to verify all systems have passed safety verifications and the rocket system is ready to launch An orange color safety light that indicates the AGSE power is on has been placed into the system. It is solid in color when the AGSE is paused and flashing with a Hz frequency when it is powered The all go systems go lighting system was designed in way that light only comes on after all procedures have been completed Testing and demonstrations were done to verify light operates properly Testing and demonstrations were done to verify light operates properly Incomplete Incomplete Section 4.8.4: System Tests, Variables, and Controls The unmodified robotic arm has been tested to ensure it is suitable for the tasks required of it. It was controlled manually and used to pick up a mock payload weighing one pound. The arm was used to rotate the mock payload in the air and move it to numerous locations around the arm. Because the arm was able securely hold the mock payload while maneuvering to various positions, it was deemed suitable for handling the required payload of four ounces. A plan has been devised to thoroughly test the assembled AGSE in the interest of ensuring all subsystems work properly and are capable of accomplishing the mission. The LVE and ACI will be tested independently and programmed to complete their tasks in seconds and 0-20 seconds respectively. A control test will be performed in which a four ounce payload made of PVC will be placed on a black surface one foot away from the base of the robotic arm. The arm will start in its resting position and the payload will be oriented so that its longitudinal axis intersects the base of the arm. The AGSE will be activated and the test will conclude after the robotic arm 05

107 closes the payload bay door and returns to its resting position. This test will be timed for comparison with all additional tests. The first series of tests will vary only the position and orientation of the payload. The surface color and the starting position of the arm will remain constant. The tests will involve the arm locating, acquiring, and depositing the payload into the payload bay of the launch vehicle. After each test, the payload will be translated and rotated to a new position within the arm s reach. The next series of tests will use a green surface color. The arm will be subjected to the same tests, varying only the position and orientation of the payload and recording the duration of the test. Each test series will involve a different surface color while the arm acquires the payload in different positions. The arm will be forced to drop the payload during each series to ensure it is able to automatically reacquire the payload. The data acquired from this testing will be used to determine the placement and orientation of the payload that yields the fastest successful results. It will also be used to determine how the surface color affects the infrared sensor system. The receiver will be calibrated to provide the highest accuracy in acquiring the payload on various surface colors. Section 4.8.5: Experiment Process Procedures The AGSE consists of a robotic arm for payload retrieval, a winched truss system to raise the launch vehicle, an ignitor insertion system, and two boxes acting as the base of the launch pad. The boxes are connected by square aluminum tubes seated in 3D printed fitters allowing for easy transportation, assembly, and disassembly. The two boxes also house electronics and motors necessary for the success of the mission. Once the AGSE is deployed and the payload is placed, the launch sequence is initiated. The robotic arm is equipped with an infrared emitter and receiver mounted above the claw. The difference in reflectivity between the grass and the white surface of the payload will allow the arm to accurately detect and retrieve the payload. The arm will use an algorithm to sweep the area in a grid pattern in increments of 0.5 inches. The payload is 0.75 inches thick, so increments of

108 inches ensure that the arm passes over the payload at least once. At each interval, the infrared emitter will pulse and the receiver will detect the amount of light reflected back. When the infrared receiver detects more than 50% of the light sent from the emitter, the claw is over the payload. The arm will continue sweeping the area to acquire multiple readings where the infrared light being reflected is above the threshold of 50%. Using this data, a line will be drawn between the two points furthest apart to estimate the longitudinal axis of the payload. The claw will move to the midpoint of this line and orient itself so that the gripping axis is parallel to the estimated longitudinal axis. The claw will then descend and grab the payload, using pressure sensors to verify that the claw is holding the payload. The amount of error introduced by estimating the longitudinal axis in this manner is not significant enough to prevent the claw from acquiring the payload. Once the arm has acquired the payload, it will move to a set of coordinates so that the claw with the payload is above the open hatch of the payload bay. If the payload is dropped by the claw before it reaches the programmed coordinates, the arm will begin the scanning algorithm again. When the claw reaches the programmed coordinates with the payload, it will deposit the payload into the launch vehicle. The door of the payload bay will start at rest in an open position 5 degrees from the closed position. The robotic arm will be programmed to move to a position behind the door and push it closed. The claw will then gently press down on the closed door to ensure it is latched and secured. Once this routine is completed, the arm will move to a resting position away from the truss and launch vehicle. When the arm reaches its resting position, it will send a voltage signal to the master controller marking the completion of its phase. Upon receiving the voltage signal, the master controller will send another voltage signal to initiate the LVE. The LVE will begin raising the rocket into launch position. It will take the LVE approximately -2 minutes to raise the launch vehicle from a horizontal position to the launch position 5 degrees from the vertical. When the rocket is in launch position, the blast shield underneath the rocket will rest against the top of the launch platform (larger box). This will ensure that the rocket cannot be lifted past 5 degrees from the vertical. Once the vehicle is in the launch position, a voltage signal is sent to the master controller. After it receives the voltage signal from the LVE, the master controller will send a voltage signal to the ACI. The ACI is a telescoping linear actuator attached to the LVE directly underneath the motor of the launch vehicle. With the igniting charge attached, the ACI will extend into the motor and stop when the ignitor reaches the top of the fuel grain. After the rocket is inspected by the RSO, 07

109 the all systems go switch will be switched on to activate the launch button, and at the end of a 5- second countdown, the launch button will be pressed to launch. Once the rocket leaves the launch pad, the AGSE has completed its mission and can be reset. 08

110 Safety Section 5.: Checklists Section 5..: Final Assembly Checklist Final Rocket Assembly Initial Check-off Points Check rocket tube for any structural imperfections acquired during transport. Check rocket tube for structural integrity and flight readiness. Check nose cone and payload capsule for any structural imperfections acquired during transport. Check payload capsule for proper functioning and mission readiness. Check parachutes for any imperfections that could be a problem during recovery operations. Check parachutes and parachute bags for flight readiness. Check avionics for proper functioning. Check carbon dioxide expulsion system for flight readiness. Check motor casing for any structural imperfections acquired during transport. Check motor mount, motor casing, and thrust plate for flight readiness. Check the couplers for structural integrity and flight readiness. Check fins for structural integrity and flight readiness. Check shock cords for flight readiness. Assemble fin section with motor mount, motor casing, and other motor mounting items. Pack drogue chute and assemble carbon dioxide expulsion system. Pack main chute and insert into position. Assemble avionics bay and attach to the motor section of the rocket. 09

111 Attach parachute to payload bay. Attach payload bay and nose cone section to the rest of the rocket. Check all connections and assemblies on the rocket. Insert rocket motor into motor casing. Complete final check of the assembled rocket. Auburn USLI Safety Office signature Auburn USLI President signature X X 0

112 Final Launch-pad Assembly Initial Check-off Points Check launch-pad base for any structural imperfections acquired during transportation. Check launch rail and frame for any structural imperfections acquired during transportation. Check raising motors for proper functioning and mission readiness. Check robot arm for any structural imperfections acquired during transportation. Place launch-pad base at the final position for the mission. Insert launch rail raising motor into position and attach to the rotating axil for. Attach launch rail frame to rotating axil for raising. Attach mechanical safety for launch rail system. Attach counter weight to launch rail frame. Attach electronic igniter insertion/ ignition system onto counter weight in proper location. Attach robot arm to launch-pad base. Run a full systems check on robot arm to ensure proper functionality. Run a full systems check on rocket raising system to ensure proper functionality. Complete final check of assembled launch-pad. Auburn USLI Safety Office signature Auburn USLI President signature X X

113 Section 5..2: Launch Procedures Checklist Final Construction Check Initial Check-off Points Check for proper connections between nosecone and payload capsule. Check payload capsule for proper operation and final flight readiness. Check the main body tube for final flight readiness. Check launch lugs for proper operation. Check fins and fin connections for final flight readiness. Check engine mount for final flight readiness. Overall rocket construction readiness check. Final Launch-Pad Check Initial Check-off Points Check launch rails for proper operation and no foreign debris. Check robot arm for proper operation and final procedural readiness. Check payload capture device for proper operation and final procedural readiness. Check rocket raising arm for proper operation and final procedural readiness. Check system for automatic insertion of the electronic ignition for proper operation and final procedural readiness. Check charge status of electrical system for the launch-pad. Overall launch-pad readiness check. 2

114 Final Wireless Controller Check Initial Check-off Points Check batteries charge status to ensure proper function throughout launch sequence. Check all emergency stop and safety switches for proper functioning. Check AGSE program initializer for proper functioning. Check AGSE steps feedback indicators for proper functioning. Check extra functions of the wireless controller for proper functioning. Overall wireless controller readiness check. Final Overall Systems Check Initial Check-off Points Final overall check of rocket construction. Final overall check of launch-pad construction. Final overall check of AGSE readiness. Final overall check of wireless controller functioning and readiness. Final overall check of personnel and observers readiness. Final overall launch readiness check. Auburn USLI Safety Office signature Auburn USLI President signature X X 3

115 Launch Procedures Check Initial Check-off Points Place rocket on the launch rails ready for mission process. Have unnecessary personnel move to safe location for launch process. Have qualified personnel place electronic igniter on insertion/ ignition device. Remove mechanical system safeties. Turn on wireless transmission receiver at the launch-pad. Have all personnel move to proper launch operations locations. Ensure all safeties are set on wireless controller before turning on the controller. Turn on wireless controller. Ensure launch area is cleared for system operation. Initialize mission process. Receive proper feedback on proper instillation of payload. Initialize raising of rocket to proper launch position. Receive proper feedback on rocket reaching proper launch angle. Check with range officer to ensure range is all clear and ready for launch. Initialize electronic ignition insertion system. Receive final all clear for launch readiness. Initiate motor ignition. Check for proper ignition. 4

116 Section 5.2: Safety Officer Austin Phillips will serve as the ideal choice for a safety officer. A senior in aerospace engineering at Auburn University, Austin is a fully trained and certified EMT and firefighter in the state of Alabama. Working full-time as a firefighter for the City of Auburn as well as being a student at Auburn, Austin is well versed in crisis-management and safety practices. His extensive training makes him an invaluable resource towards maintaining safety throughout the competition. In addition, having a High Powered Level certification, and being very close to completing his Level 2 certification, Austin is well versed in the challenges and safety hazards that are associated with the construction of a high-powered rocket. Austin s role is to facilitate the safe operation of all aspects of the program from start to finish. With this goal in mind, Austin develops the checklists as well as the safety protocols that guide the team during all processes involved with the team. In addition, Austin also helps to analyze the risks associated with meeting deadlines and other timetables associated with the project. Section 5.3: Hazard Analysis Section 5.3.: Airframe Safety is taken into consideration for every part of building the rocket. There are steps that will be taken by the airframe team to ensure the safety of the members while they construct the airframe for the rocket. There are three different areas that we will look at while considering failure modes for safety protocols for airframe: operations, materials, and construction. Operations Transport Not properly transported Airframe damaged it Transportation Storage Stored in wet area Stored in dirty area Ground Operations Cracks in the carbon fiber Gaps between different parts 5

117 Excess epoxy Lack of epoxy Launch Cracks in Airframe Airframe breaking apart Construction Autoclave Materials left in Autoclave by Previous user Drain strainer not properly cleaned Explosive breakage of glass vessels Burns to hands and other body parts Lacerations to hands and other body parts Trauma to users eyes Materials catching on fire Breathing toxic fumes Autoclave not set on correct setting Aluminum mandrel Hands caught in mandrel Burns form touching mandrel after it comes out of autoclave Injury due to torque of mandrel while wrapping material Materials Carbon Fiber Allergic dermatitis from coming in contact with carbon fiber Skin irritation from coming in contact with carbon fiber Respiratory irritation from breathing in particles Trauma to users eyes from fragments of carbon fiber Carbon fiber should be kept away from electrical equipment Epoxy Trauma to eyes from epoxy coming in contact with eyes Setting up before work is completed Mixing too much epoxy 6

118 heating up and melting through container Not properly disposed of All of these failure modes for operation, construction, and materials have been taken into consideration and the proper mitigations have been put into effect to ensure the safety of team members and the environment. Mitigation tables for failure modes within airframe are listed below. Personal hazards that could occur during the construction of the airframe and during the launch have been assess to ensure the safety of team members and people in the area around the launch site. Mitigation tables have been put in place to make team members aware of these hazards to minimize the risk of them occurring, these mitigation tables are listed below. Along with the mitigation tables team members are required to read over the MSDS sheets that pertain to the material or machine that they are working with. To prevent personal hazards while operating the autoclave each team member should be knowledgeable about how the autoclave operates by reading over the operator s manual for the autoclave, alone with looking over the mitigation table that has been put in place. Risk Mitigation Table: Airframe Potential Risk Potential Effect Impact Risk Mitigation Risk2 Airframe not properly transported () Airframe not properly stored (2) Cracks in Airframe (3) Gaps between airframe and other Damage to airframe 4 3 Make sure Airframe is properly secure in a dry and clean area Damage to airframe 4 3 Airframe will be stored in a clean and dry area Breaks on launch 4 3 Airframe will be injuring team inspected during members construction and before launch to ensure there are no cracks Failure during 4 3 Airframe will be launch causing inspected during construction and 7

119 parts of the rocket (4) injuries to team members Lack of epoxy (5) Airframe breaks apart during launch Integrity of epoxy on fins attached to motor mount (6) Integrity of epoxy on fins attached to airframe (7) 3D printed altimeter bay is properly mounted (8) Damage to fins and motor mount during launch and landing Damage to fins and airframe during launch and landing Damaged to altimeter bay upon impact of landing before launch to ensure that there are no gaps in between parts 4 3 Airframe will be inspected during construction and before launch for lack of epoxy 4 3 Epoxy between fins and motor mount will be inspected during construction and before launch 4 3 Epoxy between fins and airframe will be inspected during construction and before launch 4 3 Altimeter bay will be checked after 3D printing for defects and will be inspected before launch to make sure it is properly mounted Risk Mitigation Table: Autoclave Potential Risk Potential Effect Impact Risk Mitigation Risk2 Debris flies up into users eyes () Material left in autoclave (2) Door not properly closed (3) Wrong cycle selected (4) Explosive breakage when opening (5) Trauma to the users eyes Damage to autoclave and new material Damage to material inside autoclave Damage to material inside autoclave Damage to body from explosive 3 3 Wear safety glasses or face shield while operating autoclave 4 3 Always look inside the autoclave to make sure the previous user did not leave anything inside. 2 3 Make sure the door is fully closed. 2 3 Make sure the correct cycle has been selected. 4 2 Wear proper PPE and always keep hands, 8

120 Touching materials (6) hot Materials catch fire (7) Overcooking materials (8) Toxic Fumes (9) Unauthorized (0) use Burns to hands and body Damage to the autoclave and materials will occur. Possible risk of fire spreading to the rest of building and causing harm to individuals is possible Materials become unsuitable to be used with the rocket. Material waste also occurs Respiratory issues can occur if toxic fumes are breathed in Damage to Autoclave, materials, and to personnel head, and face away from opening. 3 3 Wear proper PPE such as heat and cut resistant gloves, rubber apron, and rubber sleeve protector. 5 4 In the case of a fire a fire extinguisher must be kept in the same room as the autoclave and be easily accessible. If fire spreads contact 9 immediately. 4 4 Properly authorized individuals will use the Autoclave. All temperatures will be checked before use and a worker will be present in the same room as the autoclave until Autoclave has completed. 5 5 Respirators will be used when working with materials that give off toxic fumes when cooked. Proper ventilation of the area is required when the autoclave is working 5 3 Room where autoclave is located is locked up by authorized personnel. Autoclave is also locked to prevent unauthorized use. 2 9

121 Risk Mitigation Tables: Carbon fiber Potential Risk Potential Effect Impact Risk Mitigation Risk2 Allergic reaction from coming in contact with carbon fiber () Debris flies up into users eyes (2) Toxic particles (3) Skin irritation 3 4 Wear proper PPE when handling carbon fiber Trauma to the users eyes Respiratory irritation Electrical shock (4) Burn or electrocution 3 3 Wear safety glasses when working with carbon fiber 3 3 Wear proper breathing apparatus when working with carbon fiber 4 2 Carbon fiber is electrically conductive so it should be kept away from electrical equipment or machinery Risk Mitigation Tables: Epoxy Potential Risk Potential Effect Impact Risk Mitigation Risk2 Improper Ventilation () Vapors generated can cause headache, nausea, and hurt the respiratory system Skin Contact (2) Can cause skin irritation Proper Storage (3) Degradation of Epoxy Resin Spilling and leaking (4) Hardens on work bench or lab equipment 5 5 Keep Lab ventilated at all times when working with epoxy. Also wear ventilation masks when working with epoxy. 2 5 Wear proper lab clothing when working with epoxy. If epoxy gets on skin was off with soap and water 4 3 Store in cool dry place around 40 F to 20 F 2 4 Handle Epoxy carefully. If spilled, use paper towel to clean up and stop leakage. Use warm

122 Fire Hazard (5) Epoxy gets in users eyes (6) Damage to lab area and surrounding equipment Trauma to the users eyes water and soap to clean 5 3 Keep epoxy away from high heat sources. If fire starts use Foam or carbon dioxide to put out. Keep fire extinguisher in the same room 3 3 Wear safety glasses while using epoxy Epoxy setting up before work is finished (7) Epoxy burning through container (8) Epoxy not properly disposed (9) Waist of epoxy that is not used Damage to work environment Damage to work area and environment 2 3 Never mix too much epoxy at one time 2 3 Never mix epoxy and leave it unattended and be aware of how hot the epoxy is as it starts to set 2 3 Always properly dispose excess epoxy when finished with work. 2 When constructing the airframe there are environmental concerns that will be addressed. These concerns include how the airframe affects the environment and also how the environment affects the airframe. A risk mitigation table has been put in place for airframe environment effects to make team members aware of the impact they can have on each other. This mitigation table has been listed below. Risk Mitigation Tables: Airframe Environment Effects Potential Risk Potential Effect Impact Risk Mitigation Risk2 Harmful toxic Damage to 4 3 Make sure that fumes released into environment and autoclave is always environment by breathing air properly ventilating autoclave () 2

123 before turning on and operating Epoxy not properly Damage to work 2 3 Always properly disposed (2) area and dispose excess epoxy environment when finished with work. Airframe stored in Damage to airframe 3 3 Always make sure wet environment (3) from being wet that airframe is stored and transported in a clean and dry environment Airframe not Hazard to the 3 2 Airframe will be tract recovered on launch environment from during the launch to (4) carbon fiber and ensure that it will not epoxy be lost Section 5.3.2: AGSE During the process of building a rocket, safety is constantly kept in mind. With the design concept for this year s payload integration techniques being the autonomous ground support equipment, it is being thought of even more so. There will be guidelines implemented to ensure the safety of the members of the AGSE team while the construction and testing of the system is occurring. There are three different sections that are being looked at while considering failure modes for safety protocols for AGSE: operations, materials, and construction processes. Operations o Mission Processes o Testing Personnel Risks (Operator and Observers) Environmental Risks (Macro and Micro) 22

124 Vehicle Risks (Launch, Flight, and Recovery) Controller Risks (Electrical and Mechanical) Construction o Hand Tools o Soldering Equipment o Drill Press o Band Saw o Autoclave Personnel Risks Environmental Risks Vehicle Risks Materials o Carbon Fiber o Aluminum o Epoxy o Electric Motor o Copper Wires o Flux and Soldering Materials Personnel Risks Environmental Risks Failure Risks Operations: Mission Processes 23

125 Risk Mitigation Table: Mission Process Potential Risk Potential Effect Impact Risk Mitigation Risk2 System does not react to submission of wireless transmission. () System does not complete mission. 3 Checks will occur to make sure proper electrical charge will be delivered to the rocket system and Robot arm does not find the payload. (2) Robot arm does not find the payload bay door for insertion of payload. (3) Incomplete insertion of the payload into the payload bay. (4) Incomplete closure of the payload bay door. (5) Launch rail does not raise up at all. (6) System does not complete the mission. Robot arm harms payload bay structurally, or the system does not complete the mission. Robot arm harms payload bay or bay door structurally, or the rocket launches with unexpected aerodynamic feature causing unexpected flight characteristics. Payload could fall out during flight failing to complete the mission, or the rocket launches with unexpected aerodynamic feature causing unexpected flight characteristics. Mission is unable to be completed. wireless control. 3 Extensive testing will be done on the robot arm and controlling program to ensure proper functioning. 4 2 Extensive testing will be done on the robot arm and controlling program to ensure proper functioning. 4 2 Extensive testing will be done on the robot arm and controlling program to ensure proper functioning. 4 3 Extensive testing will be done on the robot arm or door control system to ensure proper functioning. 4 Extensive calculations and testing will be done 24

126 Launch rail does not raise up to the required angle of eighty-five degrees. (7) Rocket does not leave launch-pad. (8) Motorized FM antenna electronic igniter insertion/ ignition module receives interference that causes early ignition. (9) Igniter does not receive correct voltage for ignition. (0) Wireless controller fails to receive proper feedback on part Mission is unable to be completed, rocket could launch in wrong direction or inclination. Launch-pad and/or surrounding area catches of fire endangering other rockets and/or observers, and mission is a failure. Improper timing for ignition changes burn type for the rocket motor causing unexpected flight characteristics, or sends the rocket flying in the wrong direction or inclination. Rocket does not launch resulting in mission failure. Causes the system to either stop at that step or induces a step to react before to ensure proper available torque to ensure raising of the launch rail. 4 2 Extensive testing will be done on the raising system and program to ensure proper angle for the expected launch. 5 2 Launch rails will be examined prior to launch to ensure nothing will hinder the ejection of the rocket. Also extensive testing will occur to ensure the proper functioning of the launch rails and launch lugs. 4 The antenna will be shielded to ensure no interference by FM radio waves, also igniter requires much higher voltage to ignite than the induced voltage by the radio waves. 3 2 Extensive testing will be done to ensure proper voltage is delivered to properly ignite the electronic igniter. 4 2 Extensive testing will be done to ensure proper feedback is given by the parts of the 25

127 of launch sequence. () the step is expected. launch sequence, and redundant safeties will be inserted between the steps of the launch sequence. Operations: Testing Risk Mitigation Table: Testing Potential Risk Potential Effect Impact Risk Mitigation Risk2 Improper weight for counter weight test. () Robot arm strikes ground or launchpad frame at a high rate of travel. (2) Launch rail raising motor raises the truss at an unexpected acceleration. (3) Wrong movement performed by robot arm. (4) Causes the rail truss to raise at an unexpected acceleration causing structural damage or damage to the tester. Structural damage occurs requiring new piece being bought or fabricated, potentially sending fragments of arm material at testers with harmful potential. Truss receives some structural damages that require a new truss to be fabricated. Robot arm harms payload bay or bay door 3 3 Careful calculations will be done to ensure the tested counter weight will be in a proper range for the desired acceleration. 3 Careful coding will be done to ensure that the robot arm will be able to know what its limits in directional travel are to ensure a collision as such described will not occur. 4 Careful calculations will be done to ensure proper motor selection for expected load, and extensive testing will be done to ensure proper control of the motor. 4 Initial testing will be done on a dummy structure 26

128 structurally, or damages some other critical component or structure. that will not be able to harm the robot arm and allows multiple trials to be run so movements can be made more precise. Construction (refer to tool specific tables for specific risk mitigations) Risk Mitigation Table: Construction Potential Risk Potential Effect Impact Risk Mitigation Risk2 Launch rail truss has an anomaly in material. () Robot arm is improperly assembled. (2) Improper material used during fabrication of AGSE system. (3) Improper connection is used between the launch rail truss and rotating axil. (4) Causes reduced structural integrity potentially leading to a failure during a launch or testing. Causes improper strength in arm or improper movements in the arm. Unexpected strength properties causes a failure in part when expected load is applied. Causes the truss to fail under loading and causes structural damage to the AGSE system. 4 Proper fabrication practices will be followed to ensure the best product is made. 3 Proper instructions will be followed during the construction process with an additional construction check occurring before any operation occurs. 4 Careful calculations and trade studies will be performed to ensure the proper material is chosen for each part. 4 2 Proper research and calculations will be done to ensure proper connection is selected. Materials (Refer to specific risk mitigation tables in the airframe and recovery sections) 27

129 Risk Mitigation Table: Other Materials Potential Risk Potential Effect Impact Risk Mitigation Risk2 Improper force applied to aluminum part. () Aluminum part shears or buckles under loading causing failure in the structure. 3 2 Proper calculations will be performed to ensure the proper sizing for aluminum Aluminum shards are sent flying on an uncontrolled path. (2) Electric motor has improper voltage and amperage sent through the system. (3) Improper wire selection. (4) Excessive amounts of flux and soldering materials are used. (5) Person machining the aluminum gets burnt or cut by the sharp, hot shards, or the shards catch some substance in the workshop on fire. Causes the motor to spin too fast causing damage to surrounding personnel or structures, or causes the motor to short causing an electrical fire. Causes excessive amounts of resistance resulting into an electrical short, which in turn results in an electrical fire. Causes for potential shorts in the electrical circuits, which could result in an electrical fire. structural pieces. 4 3 Personnel machining aluminum will be required to wear proper personal protective equipment. Also work area will be required to be cleaned up before and after the machining process. 4 2 Motor will only be tested in a mount that will not allow extraneous movement, and safeties will be put in place to ensure only safe voltages will be sent to the motor. 4 2 Proper calculations and research will help to better determine the correct size of the copper wire to be used. 3 2 Properly training personnel will help in keeping soldering materials within acceptable ranges. 28

130 Section 5.3.3: Recovery Safety is taken into consideration for every part of building the rocket. There are steps that will be taken by the recovery team to ensure the safety of the members while they construct the recovery system for the rocket. There are three different areas that we will look at while considering failure modes for safety protocols for recovery: operations, materials, and construction. Operations: During the flight. Potential Failure The parachute(s) is not packed properly. () Parachute tears (2) Parachute fails to deploy (3) Flight Recovery Operations Potential Effect Impact Risk Risk Mitigation The parachute 5 4 Strict packing does not fully instructions will deploy causing be followed by rocket to fall in the team an uncontrolled members to manner. ensure a proper packing of the parachutes. A checklist will be filled out during the packing process and signed off by proper The parachute fabric material is torn causing the rocket to fall in an uncontrolled manner Parachute fails to deploy causing the rocket to fall supervisors. 5 3 Fabric material of the parachute will be strength tested before actual use. Container in which the parachute is kept will not contain any sharp edges. 5 4 Multiple tests will be taken with the parachute to Risk2 2 29

131 The shock cords break after deployment of parachutes. (4) Tensile strength test of shock chord back lashes (5) Winds blow rocket off course. (6) in an uncontrolled manner Uncontrolled descent of the rocket with potential crowd endangerment. Damage to body parts of the workers involved Rocket could become lost, damaged, or ensure the parachute will deploy. On the day of launch systems will be checked to ensure the parachute will deploy at the proper time. Nose cone will be checked multiple times to ensure proper fitting so parachute will not be blocked by the nose cone when deploying. 5 3 Shock cords will be thoroughly tested to ensure the strength capabilities of the chosen cord material. 4 3 All workers when performing the tensile test must stay an appropriate distance away from the testing area. People performing the tensile test must also be wearing safety glasses and appropriate lab clothing. 5 3 The rocket will not be launched in improper 30

132 The parachute deploys at the incorrect time. (7) The altimeter fails. (8) The drogue parachute fails to deploy. (9) Too much or too little charge is used in the could endanger observers. Structural damage to rocket causing unsafe descent or location of descent potentially endangering observers. The parachute deploys at incorrect time or not at all resulting in structural damage or uncontrolled descent. Potentially endangering observers. Uncontrolled descent until main parachute opening then resulting in structural damage with potential endangerment of observers. If too much is used the rocket and systems weather conditions. All parts of the rocket will have a GPS locater device securely attached. 5 4 Recovery systems will be thoroughly tested prior to flight operations, and checklists will be completed and signed off by the correct supervisors. 5 3 Extensive testing will be performed on flight computer and associated electronics ensuring proper functioning. During testing and prior to launch checklists will be filled out and signed by proper supervisors. 5 4 Drogue deployment systems will be thoroughly tested, checked off, and signed off on prior to launch operations. 5 4 The amount of charge used will be based on test

133 nose cone expulsion. (0) Altimeter switch fails () Unexploded black powder charges (2) Shear pins not failing in the recovery stage of the rocket flight(3) inside could be damaged resulting in an unstable platform or a potential ignition of flammable parts. If too little is used the nose cone will not pop off, which will result in a failure in parachute deployment. Both failures could result in launch observer endangerment. One or both altimeters does not power on, thus delaying launch and or scrubbing launch On recovery of landed rocket, possible unexploded charges are in the rocket, having the potential to explode while handling rocket Shear pins do not break causing rocket to not deploy parachutes, potentially and calculations to ensure proper expulsion. The charges will be measured extremely carefully and a checklist will be completed and signed off by the proper supervisors. 4 3 Extra working switches that are the same diameter will be available for replacement should the altimeter switch fail. 5 2 Rocket will be observed throughout flight to see if the charges go off. On recovery proper personal protective equipment will be worn in case of unexploded ordinance 5 4 Ground testing will be done using static test bases with proper safety equipment and 32

134 Nose cone coupler having an incorrect fit to the inside of the rocket body(4) resulting in an uncontrolled crash landing Because of improper fit, nose cone may not separate from body causing parachutes not to deploy proper safety zone for testing. Proper number of shear pins will be established from tests. 5 4 Ground testing will be done using static test bases with proper safety equipment and proper safety zone for testing. Proper fit of nose cone will be established from tests. Operations: During testing. Risk Mitigation Tables: Wind Tunnel Potential Risk Potential Effect Impact Risk Mitigation Risk2 Debris in the wind tunnel () Open test section (2) Inexperienced personnel (3) Damage to wind tunnel or object being tested Incorrect results calculated from the wind tunnel that can have potentially damaging effects on the rocket in the future Wind tunnel and project can be damaged through setting up and running the wind 5 3 Ensure all material on tested object is attached firmly. Wind tunnel is cleaned before testing. 5 2 Ensure the wind tunnel is closed when performing tests 5 3 Lab with wind tunnels will be locked to prevent any unauthorized use from occurring 33

135 Running the wind tunnel too high (4) Overusing Motor (5) tunnel the wrong way. Can cause structural damage within the wind tunnel, hurt the intended test object, and hurt the engine running the wind tunnel Engine becomes damaged and would cost large amounts of money to repair or replace 5 3 Limits will be sent on how high the wind tunnel can be run at. Authorized personnel will be when running the wind tunnel 5 3 Scheduling for use of the wind tunnel will be necessary. Periodic checks of the system will be performed to keep engine running properly Risk Mitigation Tables: Tensile Test Rig Potential Risk Potential Effect Impact Risk Mitigation Risk2 Object being tested is improperly aligned () Fractured particles during test (2) Heavy weights and high forces generated (3) Results acquired from tests are incorrect and have a damaging effect on the rocket in the future Damage to eyes and body extremities when the item being tested fractures Body damage, specifically crushed body extremities, from misuse of 4 4 Object is carefully measured by an authorized worker and double checked by a second authorized worker to ensure proper alignment 4 4 All personnel must stay a safe distance away from tensile test rig while performing test. Safety eyewear must also be worn along with proper clothing covering body extremities 5 2 Don t touch object being tested when machine is active and stay a safe distance away 34

136 Unauthorized use (4) Improper testing material (5) machine while testing Damage to machine, personnel, and tested object Unneeded use of machine, possible damage to machine, and waste of material 5 2 Have machine locked up by an authorized worker and keep power off 3 3 All workers must check with authorized personnel to make sure the material they are testing is ok to be tested with the machine Risk Mitigation Tables: Shear pin test rig Potential Risk Potential Effect Impact Risk Mitigation Risk2 Shear pin being tested is improperly aligned () Fractured particles during test (2) Heavy weights and high forces generated (3) Unauthorized use (4) Results acquired from tests are incorrect and have a damaging effect on the rocket in the future Damage to eyes and body extremities when the item being tested fractures Body damage, specifically crushed body extremities, from misuse of machine while testing Damage to machine, 4 4 Shear pin is carefully measured by an authorized worker and double checked by a second authorized worker to ensure proper alignment 4 4 All personnel must stay a safe distance away from tensile test rig while performing test. Safety eyewear must also be worn along with proper clothing covering body extremities 5 2 Don t touch object being tested when machine is active and stay a safe distance away 5 2 Have machine locked up by an authorized 35

137 Improper testing material (5) personnel, and shear pin Unneeded use of machine, possible damage to machine, and waste of material worker and keep power off 3 3 All workers must check with authorized personnel to make sure they have the authorization to test a shear pin Materials: Risk Mitigation Tables: Kevlar Potential Risk Potential Effect Impact Risk Mitigation Risk2 Breathing in Fiber Dust () Fiber dust in eyes and on skin (2) Touching moving Kevlar fiber (3) Fire Hazard (4) Leaving in direct sunlight (5) Respiratory Problems Can cause irritation to both eyes and skin 5 5 Always wear respirators when working with Kevlar 3 4 Wear protective eye gear when working with Kevlar. If dust gets in eyes wash out immediately with water Damage to limbs 5 4 Do not touch moving Kevlar with fingers or get near moving Kevlar. Treat any cuts with first aid. Any serious lacerations call 9 Potential to catch on fire given the wrong conditions Discoloration of Kevlar 5 3 Keep fire extinguisher in the same room when working with Kevlar Keep stored in closed containers Risk Mitigation Tables: Nylon Potential Risk Potential Effect Impact Risk Mitigation Risk2 Breathing in Fiber Dust () Respiratory Problems 5 5 Always wear respirators when working with Nylon 2 36

138 Fiber dust in eyes and on skin (2) Melted Nylon (3) Can cause irritation to both eyes and skin Potential to catch on fire given the wrong conditions, melted cast nylon will cause thermal burns 3 4 Wear protective eye gear when working with Nylon. If dust gets in eyes wash out immediately with water 5 3 Keep Nylon away from sparks and open flames. Keep fire extinguisher in the same room when working with Nylon. Any melted Nylon on skin Do Not attempt to peel off 2 Risk Mitigation Tables: Carbon Dioxide Potential Risk Potential Effect Impact Risk Mitigation Risk2 Improper Ventilation () Explosion of canisters containing CO2 (2) Broken O-Ring (3) Over pressurizing rocket (4) CO2 gas can cause headache, nausea, and hurt the respiratory system Canister shrapnel can cause serious damage to the body CO2 can leak into the surrounding air and be uncontrollable when tapped Over pressurization can cause problems with parachute deployment and damage the rocket when released 5 5 Keep lab ventilated at all times when working with CO2. Also wear ventilation masks when working with CO Cylinders should be stored upright in a well-ventilated, secure area, protected from the weather. Storage area temperatures should not exceed 25 F 4 4 Periodic check of O rings on Canisters will be implemented. All faulty O-rings will be replaced immediately 4 4 Authorized individuals on the recovery team will determine appropriate amount of CO2 pressure. 2 37

139 Under pressurizing rocket (5) Under pressurization can cause problems with parachute deployment, worst possible scenario the parachute doesn t come out at all Multiple tests will be done before full scale use. 4 4 Authorize individuals on the recovery team will determine the appropriate amount of CO2 pressure. Multiple tests will be done before full scale use Risk Mitigation Tables: Black Powder Potential Risk Potential Effect Impact Risk Mitigation Risk2 Improper Ventilation () Body Contact (2) Highly Reactive Substance (3) Improper storage (4) Black powder is hazardous to respiratory system when inhaled. Also particles may form explosive mixtures in air. Can irritate skin and eyes Can catch workers and rocket on fire, and in large amounts cause explosions that can cause burns, shrapnel cuts, and toxic gases Degrades material and increases likelihood of black 5 5 Keep lab ventilated at all times when working with Black Powder. Also wear ventilation masks when working with Black Powder 5 4 Wear proper lab clothing and protective eye wear when working with black powder. Wash skin if contact is made and flush large amounts of water into eyes if contact is made there. 5 4 Black powder must be handle extremely carefully. Keep away from heat, sparks and open flames. Avoid impact or friction. Have fire extinguisher ready at all times. 5 4 Storage of black powder must be kept between 40 F to 2 38

140 Improper measuring of black powder for rocket use (5) powder catching fire Can cause problems with parachute deployment and can damage rocket if measured amount is too much 20 F in a cool dry place and tightly sealed. Must not be stored with any other flammables 5 4 Authorized individuals on the recovery team will measure the correct amount and do extensive testing to ensure proper amount is used Risk Mitigation Tables: Fiber glass Potential Risk Potential Effect Impact Risk Mitigation Risk2 Ventilation issues () Eye and Skin contact (2) Can cause Respiratory problems Can cause irritation with skin and eyes 5 5 Proper ventilation in lab area and wear respirators when working with Fiber glass 3 5 Wear proper lab clothing and wear eye protection when working with Fiber glass 2 Construction: Risk Mitigation Table: Orbital Sander Potential Risk Potential Effect Impact Risk Mitigation Risk2 Hands and fingers damage from moving parts () Damage to or loss of extremities 5 4 Using thick gloves to operate the sander. Turning the sander off when Eye Damage (2) Wood chip, metal particles and other debris hitting eyes not in use. 5 5 Wear safety glasses or other eye protector Electric Shock (3) Electrocution 5 3 Prevent contact with grounded surfaces and never 39 2

141 Unintentional Starting (4) Improper Tool Storage (5) Hazardous Work Environment (6) Improper Work Attire (7) Dust, carbon fiber and metal shards, and air quality (8) Insecure project (9) Over-reaching (0) Improper Tool Maintenance () Over Exerting Tool (2) Tool damage to body, work area, or project due to surprise start Misuse of tool by unauthorized personnel Damage to body, work area, or project from debris in work operate in damp or wet locations. Look for problems with wiring before use 4 4 Don t carry the tool while connected to power source and make sure the tool is in off position before plugging in 5 3 Store tool in a dry place and in a locked storage area 5 4 Clean all work areas before use of orbital sander and after use of orbital area sander Damage to body 5 5 Always wear long pants, closed toe shoes, and long sleeved shirts when operating tool. Damage to throat and lungs Damage to project and damage to hands Improper use of tool and possible damage to body Ineffective tool that causes unsafe handling and damage to body or project Causing damage to project due to excessive force applied to tool 5 5 Always use respirator or mask when operating orbital sander 4 3 Properly secure project with clamps or other immobilizing tools 3 2 Ensure proper footing and balance while operating sander 4 3 Sander must remain clean, sand paper replaced periodically, and inspections made on wires 3 3 Let the tool due the work, don t apply too much force 2 40

142 Improper Tool Use (3) Improper Tool Replacement Parts (4) Misuse of tool on a project it was not intended for Tool becomes unusable while using the sander 4 3 Do not use the tool for jobs other than proper sanding 3 3 Only use replacement parts intended for the orbital sander Risk Mitigation Table: Sewing machine Potential Risk Potential Effect Impact Risk Mitigation Risk2 Sewing over fingers () Pin misuse (2) Improper machine use (3) Cord can fray (4) Hurting fingers and causing irreparable damage Damage to body from the pins and damage to project Inexperience personnel can damage material and damage self This can cause trip hazards and fire hazards 5 3 Proper gloves will be worn while operating machine. Training to use the machine must be done before using sewing machine 3 3 Proper gloves will be worn while operating machine. Training to use the machine must be done before using sewing machine 5 3 Proper training must be done before any use of machine can commence 5 3 Regular maintenance of machine will occur before and after use. Chord will be close to wall when in use to prevent tripping. Machine s chords will be looked over regularly. Risk Mitigation Table: Hand Tools Potential Risk Potential Effect Impact Risk Mitigation Risk2 4

143 Improper use () Body damage from tools (2) Improper tool maintenance (3) Flying Debris (4) Insecure workbench or project (5) Improper tool storage (6) Irreparable body damage can occur. Damage to project will also occur Infection can occur on untreated wounds and tetanus can infect wound Damage to project or body from tools breaking or not working as designed Debris may cause specifically eye and/or body damage Damage to project or body can occur if project is not secure when using hand tools Tools can become damaged if stored improperly. Loss of tools can occur. Potential for unauthorized use of tools can occur 5 5 Authorized personnel will be the only one s operating the hand tools. Hand tools will be checked with team leads to make sure the hand tool in use is appropriate for the specific project job 5 4 Proper clothing will be worn at all times to prevent damage to body. If damage does occur clean wound and provide first aid. Visit a doctor if wound doesn t heal properly and infection is seen 5 5 Regular scheduled maintenance will occur for every tool. Tools beyond repair will be thrown out to prevent more use. 5 3 Proper eye ware will be worn at all times to prevent injury. Proper clothing will also be worn 5 4 Project will be secured properly by straps, clamps, or through help by a work partner before any hand tool use. 5 4 All hand tools will have a designated place to be stored. All tools will be kept under lock

144 Section 5.3.4: Outreach Safety is the primary concern in every aspect of the Auburn USLI rocket program, especially when young children are involved. There are steps that will be taken during the outreach program to ensure safety to the children in the community and will allow them the most amount of enjoyment while learning about rockets. The three primary safety concerns are: Operations, Construction, and Materials. Operations o Transportation to outreach site Car accident o Introduction/help students design their rockets Children jam fingers Children hurt by tools o Multiple rocket launchings Rocket stands fall Rockets have mid-air collisions Rockets land in the woods Construction o Tools for rocket kits Children incapable of using tools o Toy rocket motor Children accidentally ignite motor during time other than directed Materials o Toy rocket kits Children break rocket model 43

145 Hard pieces may hurt children Risk Mitigation Table Example: Outreach Operations Potential Risk Potential Effect Impact Risk Mitigation Risk2 Car Accident () Minor injuries to death 5 3 All participants will wear seatbelts and only certified drivers will operate Children jam fingers (2) Children accidentally hurt by tools (3) Mid Air rocket collisions (4) Rocket stands fall (5) Rockets fall in the woods (6) Children s fingers would experience minor pain Children could experience trauma to numerous body areas. Rockets would not reach highest altitude due to mid-air collision Failure of rocket launch Slight environmental contamination. motor vehicles. 2 2 USLI team will demonstrate how to perform all tasks for rocket completion and help the children when needed. Students will always be supervised. 3 2 All tools that could prove dangerous to children will be operated by USLI team members while wearing necessary protective equipment. 2 Students rockets will be launched from significant distances from each other. Rockets will be launched one at a time 2 2 All equipment will be examined prior to departing for the outreach event. Any non-functioning equipment will be fixed or replaced. 2 2 All rockets will not be designed to achieve significant 44

146 distance and all will be recovered. Risk Mitigation Table Example: Outreach Construction Potential Risk Potential Effect Impact Risk Mitigation Risk2 Children ignite motor at time other than directed () Trauma to hands, eyes, ears, nose, 5 2 Children will be under constant supervision and any potentially dangerous materials will be handled by the USLI outreach Children incapable of using tools (2) Danger to child, and other children s face, hands, and body team 3 2 Children will be under constant supervision and any potentially dangerous use of tools will result a removal of the tool. The task will then be completed by the USLI outreach team for the child Risk Mitigation Table Example: Outreach Materials Potential Risk Potential Effect Impact Risk Mitigation Risk2 Children Break Rocket model () Hard pieces may hurt children (2) Student will not be able to launch a rocket or participate in the primary outreach activity Trauma to children hands, eyes, nose, mouth, ears 2 2 Students will be under constant supervision and any misbehavior will be handled appropriately 2 2 Students will be under constant supervision and any misbehavior will be handled appropriately 45

147 Section 5.4: Preliminary Environmental Effects Section 5.4.: Vehicle Effects on Environment The rocket has many different effects on the environment with the different substances it is made of to the exhaust it outputs. Some of the main affecting substances would include the epoxy, carbon fiber, and carbon dioxide. The epoxy releases volatile organic compounds along with other unhealthy gases and chemicals during the curing process, and then there is always left over cured epoxy that is just thrown away. When the cured epoxy cups leave the facilities they are taken to landfills where they help add to the mountains of trash and leach hazardous chemicals into the ground below. The carbon fiber, when machined, releases tiny dust particles into the air that are so small they are hard to filter out of the air. That leads to people and animals breathing in the dust, which could lead to lung, eye and skin irritation. Then the carbon dioxide is hazardous for humans and animals to breathe in, while in the condensed form that it will be after expulsion from the rocket. Then when the rocket motor is ignited it will burn whatever is in the path of the motor exhaust, which could potentially set fire to the fields where we will launch or the vegetation where it will land. Section 5.4.2: Environmental Effects on the Vehicle The environment also has many ways that it can affect the rocket such as through humidity, wind current, and thermal fluctuations. The humidity can cause corrosion in the different metals and materials used in the systems, but also the humidity can really mess with the electronics on-board the rocket as well as the electronics on the launch-pad. Then the wind currents can bother the rocket while it is being transported from place to place, or while it is sitting on the launch-pad preparing to launch. Although the most significant affect the wind has on the rocket would be during flight because it could make the flight path extremely unexpected and very unstable. Then the thermal fluctuations can cause different materials to behave very differently as well as causing the electronics to have issues as well. 46

148 Section 5.5: Updated Environments Effects Section 5.5.: AGSE When considering the different possibilities for the environmental condition on a day used for testing subsystems or for operating the entire structure, there are different affects that could occur. If there is a high enough wind speed, but is still low enough for rocket launch, the payload could be blown out of the robot hand and would require the payload to be found again and be re-secured. High humidity or rain could cause electrical shorts or cause potential oxidation on potential metal parts. These could be mitigated by concealing electrical components in a weather proof casing and either coating metal parts or consider fabricating parts out of non-reactive materials. As well the system might have a short span of use that oxidation or other issues will not have time to accumulate on the components. High amounts of solar activity could lead to interference with the control systems. For this testing will be done to ensure regular to slightly increased solar activity will not interfere. Section 5.5.2: Recovery On day of launch, any assembly needed prior to rocket launch will take place in a designated area. All recovery pieces will be carefully watched and accounted for to ensure debris isn t lost in the environment. Chemicals and pyrodex will be in special containers and handled by the lead members of the recovery team to prevent spillage in the environment. After completion of assembly, the area the assembly was done in will be searched thoroughly for any pieces or debris. This will be done by the recovery team standing in a police line and walking slowly searching for anything that shouldn t belong in the environment. The area that was worked in should look better than before the assembly began. When preparing to launch the rocket, the area around the rocket should be cleared to ensure if any explosion happens on the launch pad the surrounding area won t catch fire. Launch area will also be checked for debris prior to launch. After the rocket has landed, the recovery team will go out along with the rest of the team to clean up the rocket. If pyrodex was used with the parachute deployment, the rocket will be checked to make sure they ve all fired. If any of the parachutes have caught fire, the recovery team or safety 47

149 officer will immediately put them out. Any debris that comes off the rocket during landing will be picked up and properly stored or disposed of. Section 5.5.3: Airframe One of the environmental effects that would affect the airframe is high winds on the launch. Upon launch if there are high winds then the rocket will be launched at an angle to counter act the high wind. Other environmental factors that could affect the airframe upon launch are things such as rain or lightning. If either of these occurs on the launch date the launch will be postponed until the rain or lightning has stopped and it has been authorized that it is safe to launch. Section 5.5.4: Outreach The main environmental effect that can occur with outreach is that wind speeds higher than about five miles an hour can push the small model rockets off their projected paths. With the small amounts of materials in the rockets and motors there are minimal effects on the surrounding environments unless one of the rockets set the surrounding grass on fire, but that will be mitigated beforehand. 48

150 Section 6.: Timeline Project Plan Moving forward into the final construction phase of the project, the timelines become ever tighter. Therefore, proper management of time is absolutely essential as the margin for errors get slimmer. Delays in the project at this juncture prove even more costly as the full-scale testing and AGSE prototyping come to a close. Thus, the following table and figures demonstrate Auburn University s plan for carrying out the fabrication and testing of the full-scale vehicle and AGSE through the FRR phase of the project. Table 6.: Task List for FRR Phase Task Name Duration Start Finish FRR Phase 43 days Fri /6/5 Mon 3/6/5 Full-Scale Vehicle Production 38 days Fri /6/5 Sun 3/8/5 Body Tube Layup days Fri /6/5 Thu /29/5 Body Tube Inspection 2 days Fri /30/5 Sat /3/5 Fin Slot Cutting 3 days Sun 2//5 Tue 2/3/5 Flat Plate Manufacturing 3 days Fri /6/5 Tue /20/5 Fin Cutting 2 days Wed /2/5 Thu /22/5 Fin Inspection day Thu /22/5 Thu /22/5 49

151 Bulk Plate/ Centering Ring Cutting day Thu /22/5 Thu /22/5 Bulk Plate Inspection day Fri /23/5 Fri /23/5 Coupler Manufacturing days Fri /6/5 Thu /29/5 Coupler Inspection day Fri /30/5 Fri /30/5 Switch Ring Manufacturing 4 days Fri /30/5 Wed 2/4/5 Altimeter Bay Manufacturing 4 days Wed 2/4/5 Sat 2/7/5 Altimeter Bay Testing day Sat 2/7/5 Sat 2/7/5 Nose Cone Mold Manufacturing days Fri /6/5 Thu /29/5 Nose Cone Manufacturing 6 days Sun 2//5 Fri 2/20/5 50

152 Payload Bay Door Manufacturing 3 days Fri 2/20/5 Tue 2/24/5 Payload Bay Manufacturing 2 days Fri 2/20/5 Sat 2/2/5 Centering Ring Mounting 2 days Thu /22/5 Fri /23/5 Motor Mount Tube Mounting 2 days Fri /23/5 Sat /24/5 Bulkplate Mounting 2 days Sat /24/5 Sun /25/5 Parachute Construction 5 days Sun /8/5 Thu /22/5 Pack Testing 2 days Thu /22/5 Fri /23/5 Ejection Ground Tests 7 days Sat /24/5 Sat /3/5 Full-Scale Test 2 days Sat 3/7/5 Sun 3/8/5 AGSE Production 43 days Fri /6/5 Mon 3/6/5 5

153 Robotic Arm Testing day Sat /24/5 Sat /24/5 Robotic Arm Modification 6 days Sat /24/5 Fri /30/5 Modified Robotic Arm Testing 6 days Fri /30/5 Fri 2/20/5 Igniter System Manufacturing 27 days Fri /6/5 Fri 2/20/5 Igniter System Testing day Fri 2/27/5 Fri 2/27/5 Launch Box Construction 6 days Fri /6/5 Fri /23/5 Launch Rail Construction 37 days Fri /6/5 Fri 3/6/5 T-Rail Purchasing 6 days Fri /6/5 Fri /23/5 52

154 Motor and Gearbox Integration 2 days Fri /6/5 Fri /30/5 Motor and Gearbox Testing 2 days Fri /30/5 Fri 2/27/5 Launch Rail Testing 6 days Fri 2/27/5 Fri 3/6/5 Systems Integration Testing 5 days Fri 3/6/5 Thu 3/2/5 53

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158 Section 6.2: Critical Path As the program progresses, it becomes evident that the AGSE production is the critical path. Since each individual system must function codependent with each other section, any delay in the production of a separate subsystem of the AGSE system delays the period where they can be tested. In addition, it also significantly delays the testing of the integrational scheme. Therefore, special attention is being paid to ensure that the AGSE subsystems are developed on time in order to guarantee successful completion of the project. Section 6.3: Budget Plan The following represents the team s current estimations of the budget. With over half the planned expenses listed in the budget already completed, the team is highly confident that the overall program will come in well under the required budget. Recovery Ripstop Nylon 20 sq yard $7.99 $59.80 Tubular Kevlar - in diameter 00 foot $0.66 $66.00 Fireproofing Sleeves (Nomex 4 unit $0.00 $40.00 Sleeves) Black Powder lb $5.99 $5.99 Ubolts + Attaching Hardware 4 N/A $3.75 $5.00 Tubular Nylon 50 foot $0.33 $6.50 Telemetrum GPS Unit $32.00 $32.00 PerfectFlight Altimeters MAWDS 2 Unit $79.95 $59.90 (StratoLogger) 9V Batteries 8 Units $.99 $5.92 Electric Matches 50 Units $.00 $ Nylon Machine Screws 3 Packages $3.69 $.07 QuickLink Connectors 8 Units $2.25 $8.00 GPS Dog Collars 4 Units $50.00 $ Total $,

159 Subscale Carbon Fiber Plates (Fins, 2 36 X 2 inch $35.00 $70.00 Bulkplates) ABS Plastic for Coupler 0. Roll $32.99 $3.30 Carbon Fiber Tube (Airframe) 4 Units $35.00 $40.00 /3" Thick Plywood (Altimeter Bay X 24 Inch $8.99 $2.25 Caps) Sheet Motor Mount Tube Unit $7.50 $7.50 J-425 Motor Unit $59.99 $ Inch Nosecone Unit $33.00 $ mm Aeropack Unit $28.00 $28.00 Total $ Full Scale Carbon Fiber Plates (Fins, 2 36 X 2 Inch $35.00 $70.00 Bulkplates) Carbon Fiber Tube 3 Units $55.00 $65.00 Motor Mount Tube 2 Units $7.00 $34.00 Fiberglass (Nosecone) roll $8.00 $8.00 L-475 unit $75.00 $75.00 Paint 2 units $80.00 $60.00 Total $ Manufacturing Motor Mount Tube 3 unit $60.00 $80.00 Epoxy - Gallon unit $00.00 $00.00 Powder Filler-Silica 5QT $20.00 $20.00 Zip-Ties 5 Package $4.99 $

160 /4" Hex Nuts Package $4.49 $4.49 /4" Lock Nuts Package $8.37 $8.37 U-Bolt 2 Unit $.69 $20.28 /4" Washers Package $3.69 $3.69 /4" Threaded Rod 8ft rod $4.99 $4.99 Aluminum Stock 4 Unit $55.50 $ Nylon Thread Spool $2.25 $2.25 Sand Paper 5 Packages $4.75 $23.75 Total $64.77 Travel Hotel 24 Units $26.00 $3, Gas for Travel 630 miles $0.50 $35.00 Total $3, Educational Outreach Plywood 2 4 X 8 Ft. Sheets $22.00 $44.00 Alpha Rocket Kits 20 Units $.00 $2,2.00 Motorcycle Battery Units $5.00 $5.00 Lumber 2 2 X 4 $8.00 $6.00 Total $2, AGSE Plywood 3 4 X 8 ft Sheet $22.00 $66.00 Robotic Arm Unit $ $

161 Motorcycle Battery 3 Units $5.00 $53.00 Servos 2 Units (Included $0.00 $0.00 in Robot) Hinges 4 Units $3.25 $3.00 Stepper Motors Unit $, $, Pinion Assembly Unit $5.00 $5.00 Total $2,47.00 Project Management Overhead and Miscellaneous Unit $2, $2, Purchases Uniforms 6 Units $55.00 $ Total $2, Overall Total $3, Section 6.4: Funding Plan With the budget well defined, the next step is to secure the required funding. Auburn s funding is summarized in the following table: Funding Auburn Space Grant Unit $3, $3, Organizational Board Funding Unit $2, $2, Material Donations Unit $3, $3, Total $8,

162 With the funding over $5000 the required amount of the overall program, the team is very confident in the funding required to successfully complete the program outlined in the budget plan. 6

163 Section 7: Educational Engagement Section 7.: General Mission Statement The Auburn University Rocketry Association (AURA), along with the Department of Aerospace Engineering at Auburn University, are entering into an exciting new era of growth, influence, and leadership, as devotion to the future advancement of aeronautical and astronautical sciences continues to swell within the department. Although the USLI competition requires teams to plan and execute educational engagement activities as a component of their overall project, AURA does not seek to fulfill the launch requirement for the sake of the competition. Just as NASA and the USLI competition have instilled the spirit of rocketry in AURA s team members, AURA truly aspires to instill the spirit of science, technology, engineering, mathematics and rocketry in young students here on the Plains. Statistical studies show that increasingly more youngsters are losing interest in STEM careers every year. For this reason, AURA is committed to combat attrition in this field, especially in the area of aerospace engineering. On the one hand, one may recognize that a large body exists of middle school, high school, and college students who possess great talent in math and science, and who aspire to pursue STEM careers as a vocation. On the other hand, society seems to have developed a stigma that discourages students from entering into STEM disciplines by depicting careers in STEM fields as though they are reserved for the academic elite and that only few graduates have the ability to succeed in shaping humanity s future. Contrary to that stigma, AURA believes that it is urgent to curb society s attitudes toward STEM fields. A dramatic change of perspective is needed to show that even though huge gains have been made in math, science, engineering and technology, careers in these areas are still accessible and attainable for those who set their minds to a career in STEM. Naturally, the solutions to the world s incumbent problems lie in the minds of generations to come. Given the role that Auburn University students play in the Auburn community, AURA plans to leverage its influence to enrich the young minds of students at Auburn and to promote the importance of STEM careers and aerospace interests throughout the community in the state of Alabama. 62

164 Section 7.2: Drake Middle School 7 th Grade Rocket Week This year, AURA s primary plans begin with its venture in engaging young students by bringing a hands-on learning experience for the seventh grade class of J. F. Drake Middle School (DMS). The program is entitled DMS 7 th Grade Rocket Week, and the goal of the program is to instill interest in math, science, engineering, technology and rocketry through an interactive three-day teaching curriculum that will reach more than 700 middle school students. In general, many students do not know much about rocketry or any relevant interdisciplinary applications that space exploration entails. The seventh grade science curriculum at DMS focuses on life science only. Therefore, the rocketry unit curriculum will include lessons about Newton s law and g-forces, and how they affect the human body. Also, most students have certainly never built their own rockets. So additionally, the students will be divided into teams of 2-3 groups and provided a small alpha rocket to construct and launch under the supervision of AURA and certified professionals. This program was successfully implemented during the school year, and the school has requested that we return to repeat the program with the new seventh grade class (see Figure 7. and Figure 7.2). A summarized plan of action is written below, and its detail will be refined and added as more formal pending agreements are made between the school and the team. Once all formal decisions are made final for the year, a fully detailed program handbook will be printed for the both teachers and administrators. The handbook will include specific details concerning the plan of action, the launch, scheduling outlines, procedures, worksheets, teaching materials, lesson plans, feedback forms, etc. In the following section, a preliminary draft plan of action, an ideal launch plan, and the learning objectives for the outreach program are furnished. 63

165 Figure 7.: A photo taken from DMS 7th Grade Rocket Week in April 204. Section 7.3: Rocket Week Plan of Action Day : The students will participate in an engaging in-class lesson presented by AURA members. The lesson will first teach the students about g-forces through a presentation that is followed by a practical demonstration. Secondly, students will learn how the human body reacts under stress in high and low g-force environments via a presentation and a video. This part of the lesson will be both educational and highly engaging. A curriculum guide will be provided for the teacher, along with all presentation materials that are to be utilized. A worksheet will be distributed to the students for them to fill out key concepts as they follow the lesson. Day 2: The students will be split into teams of 2-3 individuals and given a small alpha rocket assembly kit and the required materials to build and decorate the rocket. The teachers will need to divide the students into teams since the teachers can more appropriately handle their students. AURA team members will lead and guide the students and faculty in every step of assembly in a very organized and well-prepared fashion. At no point will the students be given the motors for their rockets. AURA team members and certified professionals will take care of this portion at the 64

166 launch event. The students and faculty will sand, glue, assemble and paint their own rockets as AURA team members instruct them and guide them throughout the process. Figure 7.2: A photo taken from DMS 7th Grade Rocket Week in April 204. Day 3: All science classes will head to the P.E. field on DMS s campus during each period throughout the day. Students will also be informed of all safety and launch procedures for the event when they first arrive on the field. A summary of what will take place at the launch site and a launching order will be announced on that day. Section 7.4: Rocket Week Launch Day The launch day will be held on the DMS P.E. field on the third and final day of the program. Each period of the school day, four or five science classes will proceed to the launch field. There will be multiple launch rails set up in sanctioned safe zones in different parts of the field, thus meeting all NAR Safety Guidelines for the launching of model rockets. Each class will be assigned to a launch rail, and instructions will be delivered by an AURA member. In the order that they are 65