Citrus College- NASA SL Preliminary Design Review

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Preliminary Design Review Report NASA Student Launch Mini-MAV Competition 2014-15 1000 W. Foothill Blvd. Glendora, CA 91741 Project Λscension November 5, 2014

Transcript of Citrus College- NASA SL Preliminary Design Review

Page 1: Citrus College- NASA SL Preliminary Design Review

Preliminary Design Review Report

NASA Student Launch

Mini-MAV Competition

2014-15

1000 W. Foothill Blvd. Glendora, CA 91741

Project Λscension

November 5, 2014

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Contents General Information .............................................................................................................................. 6

School Information ............................................................................................................................ 6

Adult Educators .................................................................................................................................. 6

Safety Officer ........................................................................................................................................ 6

Student Team Leader........................................................................................................................ 6

Team Members and Proposed Duties ........................................................................................ 6

NAR/TRA Sections ............................................................................................................................. 7

I. Summary of PDR Report .................................................................................................................. 8

Team Summary ................................................................................................................................... 8

Launch Vehicle Summary ................................................................................................................ 8

AGSE Summary ................................................................................................................................... 8

II. Changes Made Since Proposal ....................................................................................................... 9

Changes made to vehicle criteria ................................................................................................. 9

Changes made to AGSE criteria..................................................................................................... 9

Changes made to project plan ....................................................................................................... 9

III. Vehicle Criteria .............................................................................................................................. 10

Selection, Design, and Verification of the Launch Vehicle ............................................... 10

Mission Statement ...................................................................................................................... 10

Mission Requirements and Success Criteria .................................................................... 10

Design Review ............................................................................................................................. 11

Recovery Subsystem ................................................................................................................. 14

Recovery Subsystem Electrical Schematics ..................................................................... 16

Launch Vehicle Verification Plan .......................................................................................... 25

Manufacturing Plan ................................................................................................................... 32

Confidence and Maturity of Design ..................................................................................... 33

Project Risks ................................................................................................................................. 33

Mission Performance Predictions ............................................................................................. 36

Mission performance criteria ................................................................................................ 36

Vehicle Stability........................................................................................................................... 36

Simulation Results ..................................................................................................................... 37

Kinetic Energy ............................................................................................................................. 41

Drift from Launch Pad .............................................................................................................. 42

Interfaces and Integration ........................................................................................................... 43

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Payload Integration ................................................................................................................... 43

Internal Interfaces...................................................................................................................... 45

Launch Vehicle and Ground Interfaces .............................................................................. 47

Safety .................................................................................................................................................... 48

Launch Checklist ......................................................................................................................... 48

Safety Officer ................................................................................................................................ 50

Preliminary Hazard Analysis ................................................................................................. 51

Environmental Concerns ......................................................................................................... 55

IV. AGSE Criteria ................................................................................................................................... 56

Selection, Design Review, and Verification ........................................................................... 56

Design Review ............................................................................................................................. 56

Subsystems Required to Accomplish the AGSE objectives ........................................ 58

AGSE Electrical Schematics .................................................................................................... 70

Performance Characteristics for the Systems and Subsystems ............................... 71

Verification Plan and Status ................................................................................................... 76

Preliminary Integration Plan ................................................................................................. 80

Precision of Instrumentation and Repeatability ............................................................ 82

Key Components of AGSE ........................................................................................................ 83

AGSE Concept Features and Definition ................................................................................... 85

Science Value .................................................................................................................................... 86

V. Project Plan ...................................................................................................................................... 87

Timelines ............................................................................................................................................ 87

Budget .................................................................................................................................................. 89

Funding Plan ..................................................................................................................................... 93

Plans to Solicit Additional Support .......................................................................................... 94

Sustainability Plan .......................................................................................................................... 94

VI. Conclusion ........................................................................................................................................ 95

Appendix A: Citrus College Profile ............................................................................................... 96

Appendix B: Safety Contract ........................................................................................................... 97

Appendix C: NAR High Power Safety Code ............................................................................... 98

Table 1: Team Member Duties .......................................................................................................... 6 Table 2: Recovery Subsystem Components .............................................................................. 15 Table 3: Launch Vehicle Requirements and Verification .................................................... 25 Table 4: Recovery Requirements and Verification ................................................................ 29

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Table 5: Project Risk Quantitative Assessment ........................................................................ 34 Table 6: Project Risk Qualitative Assessment ........................................................................... 34 Table 7: Impact Level Definitions .................................................................................................. 34 Table 8: Likelihood Definitions ...................................................................................................... 34 Table 9: Project Risks and Mitigations ........................................................................................ 35 Table 10: Rocket Weight, Altitude, and Rail Velocity ............................................................ 36 Table 11: Simulation Results .......................................................................................................... 38 Table 12: Kinetic Energy of each Rocket Section ................................................................... 41 Table 13: Drift from Launch Pad (all sections) ....................................................................... 42 Table 14: Internal Interfaces of the Vehicle............................................................................... 46 Table 15: Preliminary Checklist ..................................................................................................... 48 Table 16: Vehicle Failure Modes .................................................................................................... 51 Table 17: Propulsion Failure Modes ............................................................................................. 52 Table 18: Recovery Failure Modes ................................................................................................ 53 Table 19: Environmental Hazards ................................................................................................. 55 Table 20: AGSE Functional Requirements ................................................................................ 57 Table 21: AGSE Functional Requirements ................................................................................ 62 Table 22: AGSE Functional Requirements ................................................................................ 63 Table 23: AGSE Functional Requirements ................................................................................ 65 Table 24: AGSE Functional Requirements ................................................................................ 68 Table 25: Subsystem Performance, Evaluation, and Verification .................................... 72 Table 26: AGSE Body Components ............................................................................................... 73 Table 27: Camera Subsystem ......................................................................................................... 73 Table 28: Laser Ranging Subsystem ............................................................................................ 74 Table 29: Payload Retrieval System ............................................................................................ 75 Table 30: Main Computer ................................................................................................................ 76 Table 31: Power Supply.................................................................................................................... 76 Table 32: AGSE Verification ............................................................................................................. 77 Table 33: Camera Subsystem Instrumentation Performance ............................................ 82 Table 34: LRS subsystem Instrumentation Performance .................................................... 82 Table 35: Payload Retrieval subsystem Instrumentation Performance ........................ 82 Table 36: Internal Interfaces of AGSE .......................................................................................... 83 Table 37: AGSE Objectives and Success Criteria ...................................................................... 86 Table 38: Mini-MAV Budget ............................................................................................................ 89 Table 39: Funding Plan ...................................................................................................................... 93 Figure 1: Organizational Flowchart ................................................................................................ 7 Figure 2: Booster Section ................................................................................................................. 11 Figure 3: AeroPack Retainer ........................................................................................................... 12 Figure 4: Structural and Aerodynamic Stability Subsystem .............................................. 13 Figure 5: Fin Dimensions ................................................................................................................. 14 Figure 6: Recovery Deployment .................................................................................................... 15 Figure 7: Electrical schematics for the main recovery system. .......................................... 16 Figure 8: Electrical schematics for the payload recovery system. .................................... 17 Figure 9: TeleGPS Schematic .......................................................................................................... 17 Figure 10: EM-506 GPS Tracking System Schematic ............................................................ 18

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Figure 11: Rocket Middle Section .................................................................................................. 19 Figure 12: Main Parachute Piston ................................................................................................. 20 Figure 13: Payload Containment System ................................................................................... 22 Figure 14: Payload Containment System ................................................................................... 23 Figure 15: Payload Containment Device .................................................................................... 24 Figure 16: Stability Diagram ........................................................................................................... 37 Figure 17: Thrust Curves of AeroTech K1100T and K1275R ............................................ 40 Figure 18: Thrust Curve of Cesaroni K590-DT ........................................................................ 41 Figure 19: Dimensional drawing, isometric view, and side view of the payload

containment bay. ........................................................................................................................ 43 Figure 20: Drawing of payload containment device integration into the main body.

........................................................................................................................................................... 44 Figure 21: Overall AGSE Dimensions ........................................................................................... 59 Figure 22: AGSE Exploded View ..................................................................................................... 60 Figure 23: Chassis Isometric and Exploded View ................................................................... 61 Figure 24: The Mechanical Arm ..................................................................................................... 67 Figure 25: Electrical Schematics for the AGSE.......................................................................... 70 Figure 26: Motor Driver Schematics............................................................................................. 71 Figure 27: NASA Student Launch Timeline ................................................................................ 87 Figure 28: AGSE and Rocket Construction Timeline .............................................................. 88 Figure 29: Outreach Timeline ......................................................................................................... 88 Figure 30: Budget Distribution ...................................................................................................... 92

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General Information School Information More information about Citrus College can be found in Appendix A. Adult Educators Lucia Riderer Rick Maschek Physics Faculty/Team Advisor Director, Sugar Shot to Space/Team Mentor [email protected] [email protected] (626) 914-8763 (760) 953-0011 Safety Officer Alex [email protected] (626) 643-0014 Student Team Leader Aaron [email protected] (509) 592-3328 Team Members and Proposed Duties The 2014-15 Citrus College NASA Student Launch team, the ‘Rocket Owls’, consists of five students, one faculty team advisor, and a team mentor. The student members’ proposed duties are listed in Table 1 below.

Table 1: Team Member Duties

Team Member Title Proposed Duties

Aaron Team Leader Oversight, coordination, and planning

Assistance with all team member duties Lead rocket design and construction

Alex

Safety Officer

Implementation of Safety Plan

Brian

Robotics Specialist

Lead AGSE design and construction

John

Payload Specialist

Oversight and coordination of payload acquisition, retention, and ejection

systems

Joseph

Outreach Officer

Educational Engagement Social Media, Website maintenance

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Figure 1: Organizational Flowchart

NAR/TRA Sections For launch assistance, mentoring, and review, the Rocket Owls will associate with the Rocketry Organization of California (ROC) (NAR Section #538, Tripoli Prefecture #48) and the Mojave Desert Advanced Rocket Society (MDARS) (Tripoli Prefecture #37).

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I. Summary of PDR Report Team Summary Citrus College Rocket Owls Mailing address: Team Mentor:

Lucia Riderer Rick Maschek Physics Department TRA #11388, Cert. Level 2 Citrus College 1000 W. Foothill Blvd. Glendora, CA 91741

Launch Vehicle Summary

Length: 112.5 in Diameter: 6 in Mass (without motor): 8.9 kg Weight (without motor): 87.2 N/19.6 lb Motor: AeroTech K1100T

Recovery system: Redundant Missile Works RRC3 altimeters will deploy a

30” elliptical drogue parachute at apogee, and a 72” elliptical main parachute at 800 ft (AGL). A separate pair of RRC3 altimeters will eject the nosecone and attached payload bay at 1000 ft (AGL), which will descend untethered under its own 42” elliptical parachute.

The milestone review flysheet is a separate document

AGSE Summary Title: Project scension A six-wheeled rover with rocker-bogie suspension will autonomously

identify and navigate as needed to a payload lying on the ground pick up the payload with a robotic arm identify and navigate as needed to the horizontally positioned rocket insert the payload into the rocket

The team or other personnel will manually:

move the rocket to a vertical launch position install the igniter launch the rocket

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II. Changes Made Since Proposal Changes made to vehicle criteria A Separate Pair of Altimeters Eject the Payload Bay In the new design, a separate pair of altimeters located in the nosecone ejects the nosecone and attached payload bay at 1000 ft AGL. In the previous design, the altimeters that deploy the main and drogue parachutes also eject the payload bay. The team was concerned that the ejection charge and igniter wire running from the main avionics bay to the payload parachute bay would interfere with the deployment of the main parachute piston. The new design eliminates this potential difficulty. Parachute Sizes Changed The payload parachute was increased from 36” to 42”. New weight estimates for the payload bay indicate that a larger parachute is required to ensure that the payload descends with the required kinetic energy. The main parachute is decreased to 72” from 96”. The proposal did not take into account the fact that the mass of the nosecone and payload is ejected before the main parachute deploys. Without this mass, a smaller main parachute is required for the tethered vehicle sections. Changes made to AGSE criteria Fewer Motors for AGSE Mobility It was determined that the amount of motors used to drive the AGSE was too high. The motors for the middle wheels have been removed and the amount of motors used to drive the AGSE has been reduced to four. Different Microcontroller for Image Processing Originally, the BeagleBone Black was picked in order to carry out image processing. However, the Arduino Uno has replaced it, since the component that is capturing images was designed to be used with the Arduino. Alternate Distance Measuring Device The Fabscan 3D Laser Scanner has been replaced with the Lightware SF02/F Laser Ranging Module. This is because the commercial LiDAR system originally considered was not precise enough. Specifically, the payload was not big enough to be detected by the laser scanner. The new component is capable of detecting the payload and determining its distance. Changes made to project plan No changes were made to the project plan since the proposal.

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III. Vehicle Criteria Selection, Design, and Verification of the Launch Vehicle

Mission Statement Project Λscension will use autonomous ground support equipment (AGSE) to retrieve a 4 oz. payload from the ground and secure it within a launch vehicle. The launch vehicle will carry the payload to an altitude of 3000 ft AGL. Upon descending to 1000 ft AGL, the payload bay will be ejected from the launch vehicle, and descend under its own parachute to the ground to be recovered.

Mission Requirements and Success Criteria In addition to meeting all NASA mission requirements (addressed below), mission success requires that the AGSE

identify the payload on the ground retrieve the payload insert the payload into the launch vehicle

The launch vehicle must

be aerodynamically stable reach apogee as close as possible to 3000 ft AGL deploy the drogue parachute at apogee eject the payload bay at 1000 ft AGL deploy the main parachute at 800 ft AGL land safely and undamaged transmit its location so that it can be retrieved

The payload bay must

secure the payload deploy its parachute when it is ejected at 1000 ft AGL land safely and undamaged transmit its location so that it can be retrieved

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Design Review Project Λscension incorporates the following subsystems to achieve its mission objectives.

Propulsion Subsystem The propulsion subsystem consists of the motor, the motor mount, and the motor retainer. This subsystem must

achieve a rail exit velocity sufficient for stable flight generate sufficient total impulse to reach apogee at 3000 ft AGL be mounted to the vehicle in a structurally sound way

The propulsion subsystem is shown in Figure 2.

Figure 2: Booster Section

The propulsion subsystem occupies the booster section of the launch vehicle. The motor mount is a 54 mm diameter, 23” length of BlueTube. The 54 mm diameter accommodates the greatest selection of motors that will reach 3000 ft AGL. Larger diameter motors typically propelled the rocket too high. Smaller diameter motors will not launch it high enough. Motor information is provided below in the Mass Statement and in the Mission Performance Predictions section.

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The motor mount is epoxied to three 1/2” plywood centering rings, and to the three fin tabs inserted through the airframe. The centering rings and fins are epoxied to the airframe. The motor is retained in the motor mount by an AeroPack retainer, pictured below. The two parts are threaded. The part on the right is epoxied to the aft end of the motor mount. After the motor casing is inserted into the motor mount, the left part screws on by hand, and secures the motor in the motor mount.

Figure 3: AeroPack Retainer

Structural and Aerodynamic Stability Subsystem The structural and aerodynamic stability subsystem consists of the airframe, nosecone, and fins. This subsystem must

withstand the thrust, weight, and aerodynamic forces bearing on the vehicle at lift-off and during flight

be aerodynamically stable This subsystem is shown in Figure 4 below.

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Figure 4: Structural and Aerodynamic Stability Subsystem

The airframe consists of three sections of 6” diameter BlueTube 2.0. BlueTube 2.0 is a proprietary material manufactured by Always Ready Rocketry. BlueTube is extremely rigid, but not brittle, and is easier and safer to work with than fiberglass. According to the manufacturer, BlueTube requires no reinforcement for subsonic speeds. The three airframe sections are fit together with 12” sections of BlueTube coupler tube, and secured by nylon shear pins, or by metal screws where the rocket should not separate. The coupler and airframe overlap by 6” (1 airframe diameter) at the joints to ensure that the airframe remains straight and rigid during flight. A fiberglass ogive nosecone is attached with metal screws to the payload bay. A plastic nosecone would be sufficient for this mission, but plastic nosecones in this exact size are difficult to find. Three trapezoidal fins are made from 3/16” 10-ply aircraft plywood. Plywood is heavy, but it is easy to cut, bond, and finish. The leading and trailing edges of each fin will be sanded into an airfoil shape to reduce drag. The trailing edges sweep forward, which reduces the chances that the rocket will land on a fin tip and break

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it. Fin tabs extend through slots in the airframe, and are epoxied to the motor mount. Fin dimensions are given in the following figure.

Figure 5: Fin Dimensions

The aerodynamic stability of the vehicle is discussed in the Mission Performance Predictions section below.

Recovery Subsystem The recovery subsystem consists of parachute deployment electronics and mechanisms, three parachutes and their attachment hardware, and two GPS tracking devices. This subsystem must

accurately detect apogee, 1000 ft AGL, and 800 ft AGL reliably deploy parachutes at these altitudes reduce the kinetic energy of each vehicle section to less than 75 ft-lbf at

landing transmit the location of each section to a ground station

The recovery subsystem components are summarized in the following table:

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Table 2: Recovery Subsystem Components

section descent weight of section (lb)

drogue parachute

main parachute

attachment scheme

deployment process

untethered payload

4.9

30" elliptical

42" elliptical 5/8" tubular nylon harness, sewn loops, attached to 1/4” U-bolts with 3/16” quick-links. U-bolts are mounted to 1/2"plywood bulkheads.

Redundant Missile Works RRC3 altimeters fire black powder charges.

middle 7.2

72" elliptical booster 8.8

Order of Deployment

1. The booster section separates at apogee to deploy the drogue chute. 2. The nosecone and attached payload capsule are ejected at 1000 ft, and

descend under their own parachute. 3. The main parachute is deployed at 800 ft out the forward end of the middle

section.

Figure 6: Recovery Deployment

Deployment Altimeters Missile Works RRC3 altimeters have the requisite functionality, are reliable, easy to use, and inexpensive. The RRC3 is a barometric altimeter and flight data recorder with three outputs (to initiate three separate flight events, such as deploying a parachute). After each flight, the peak altitude is reported by a series of beeps. A standard 9V battery powers each altimeter.

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GPS Tracking Devices The tethered sections of the vehicle will be tracked by an Altus Metrum TeleGPS mounted in its own compartment separate from the deployment altimeters. The TeleGPS is a compact, stand-alone GPS transmitter powered by its own lithium polymer battery. It uses APRS on the 70 cm band. The payload containment section of the vehicle will be tracked using an EM-506 GPS unit. An Arduino Uno R3 with a GPS shield will interface with the GPS and track the vehicle. The data from the GPS will be sent through an XBee Pro 900 wireless transceiver to the ground for analysis.

Recovery Subsystem Electrical Schematics Electrical schematics for the recovery system are shown below. The main vehicle has a recovery subsystem consisting of a main and drogue parachute, and has two sets of E-matches in order to deploy either one. The payload containment device has a similar setup, however, it contains only one parachute and needs only one set of E-matches.

Figure 7: Electrical schematics for the main recovery system.

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Figure 8: Electrical schematics for the payload recovery system.

Figure 9: TeleGPS Schematic

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Figure 10: EM-506 GPS Tracking System Schematic

Main Parachute Deployment A piston deploys the main parachute. This is necessary because the main parachute deploys out the forward end of the rocket after the nosecone/payload bay has been ejected. The main chute cannot reliably be blown out of the airframe by an ejection charge; the hot gases simply go around the parachute. Therefore, the parachute is attached to a piston inside the airframe. The piston is pushed out by the ejection charge, and pulls the main chute out with it. The images on the following pages show the design of the piston and its location in the middle section of the rocket.

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Figure 11: Rocket Middle Section

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Figure 12: Main Parachute Piston

Ground Testing Black powder charges will eject the nosecone and attached payload bay, and deploy the drogue and main parachutes. The team mentor will assist with ground testing these ejection charges to determine the required amount of black powder, and to ensure that deployment mechanisms are functioning properly. Only the team mentor will handle the black powder.

Drogue Parachute Fruity Chute 30” elliptical parachute. Materials: 1.1 oz. rip-stop nylon, 330 lb braided nylon shroud lines, 3/8” nylon bridle, 1000 lb swivel. RockSim estimates a descent rate of 50 ft/s under this parachute.

Main Parachute Fruity Chute 72” elliptical parachute. Materials: 550 lb nylon, 11/16” nylon bridle, 3000 lb swivel. According to Fruity Chutes, 17 lb will descend at 20 ft/s under this parachute. Our tethered booster and middle sections weigh 16 lb.

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Payload Parachute Fruity Chute 42” elliptical parachute. Materials: 1.1 oz. rip-stop nylon, 400 lb braided nylon shroud lines, 5/8” nylon bridle, 1500 lb swivel. According to Fruity Chutes, 6 lb will descend at 20 ft/s under this parachute. The nosecone and attached payload bay weigh approximately 4.9 lb, so we expect a descent rate somewhat less than 20 ft/s.

Harnesses, Attachment Hardware, and Bulkheads The drogue and main parachute swivels will be attached with 3/16” stainless steel quick links to a sewn loop in 20 ft long, 5/8” tubular nylon shock cords. Sewn loops at the ends of the shock cords will be attached with quick links to 1/4” steel U-bolts mounted on 1/2” thick plywood bulkheads. The bulkheads will be epoxied into the airframe. The nosecone and attached payload bay are untethered to the other sections of the rocket. The payload parachute swivel will be attached with a 1/8” stainless steel quick link to the sewn loop of a 3 ft, 3/8” tubular nylon shock cord. The other end of the shock cord will be attached with a quick-link to a 1/4” U-bolt mounted on a 1/2” plywood bulkhead. The bulkhead will be epoxied into the payload bay airframe.

Robustness of the recovery subsystem All recovery subsystem materials and hardware are in accord with the recommendations of the parachute manufacturer (Fruity Chutes). For rockets up to 30 lbs., Fruity Chutes recommends:

5/8” tubular nylon shock cord 3/16” stainless steel quick links

For the 5 lb., untethered payload section, smaller hardware is permitted:

3/8” tubular nylon shock cord 1/8” quick links

1/4” steel U-bolts mounted on 1/2” thick bulkheads epoxied into the airframe should be sufficient to withstand the forces of parachute deployment.

Payload Containment Subsystem The payload containment subsystem must

accept the payload from the AGSE securely contain the payload within the launch vehicle during flight eject from the launch vehicle upon descending to 1000 ft AGL contain its own recovery subsystem that satisfies the requirements listed

above in the Recovery System section. Figure 13 shows the payload containment system attached to the base of the nosecone.

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Figure 13: Payload Containment System

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Spring-loaded, rectangular doors in the airframe allow the payload to be inserted by the AGSE. The doors spring shut, locking the payload inside. The next figure shows how the containment system unbolts and can be removed from the airframe and nosecone.

Figure 14: Payload Containment System

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The detailed view below shows how the payload bay is a birch plywood box with a square slot in the aft end. Initially, the rocket is positioned horizontally, and the payload bay doors are facing up. The payload drops through the doors into the box. When the rocket is positioned vertically for launch, the payload slides down into this slot, which holds the payload in the center of the rocket during ascent.

Figure 15: Payload Containment Device

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The payload bay remains attached to the nosecone at apogee, and descends with the rest of the vehicle under the drogue parachute. At 1000 ft AGL, redundant Missile Works RRC3 altimeters in the payload containment section eject the nosecone and payload bay, which then descend under their own parachute untethered to the rest of the rocket.

Launch Vehicle Verification Plan The launch vehicle meets all requirements of the Student Launch Statement of Work. The following tables list each requirement, the design feature that satisfies the requirement, and the means of verification.

Table 3: Launch Vehicle Requirements and Verification

Requirement Design feature that satisfies the

requirement Verification

1.1 The vehicle shall deliver the payload to, but not exceeding, an apogee altitude of 3,000 feet above ground level (AGL).

At current mass, an AeroTech K1100T will reach 3000 ft AGL.

Simulations and test flights will determine the appropriate motor to attain this altitude.

1.2. The vehicle shall carry one commercially available, barometric altimeter for recording the official altitude used in the competition scoring.

One of the Missile Works RRC3 altimeters will record the official altitude.

Inspection

1.2.1.The official scoring altimeter shall report the official competition altitude via a series of beeps to be checked after the competition flight.

The Missile Works RRC3 altimeter reports the altitude via a series of beeps.

Inspection

1.2.2.3. At the launch field, to aid in determination of the vehicle’s apogee, all audible electronics, except for the official altitude-determining altimeter shall be capable of being turned off.

All audible electronics, except for official scoring altimeter, will be capable of being turned off.

Functional testing

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1.3. The launch vehicle shall be designed to be recoverable and reusable.

Current simulations predict that all rocket components will be recovered within 2300 ft of the launch pad, and all components are designed to be reusable.

Inspection, and functional testing

1.4. The launch vehicle shall have a maximum of four (4) independent sections.

The launch vehicle has three (3) independent sections.

Inspection

1.5. The launch vehicle shall be limited to a single stage.

The launch vehicle has only one stage.

Inspection

1.6. The launch vehicle shall be capable of being prepared for flight at the launch site within 2 hours, from the time the Federal Aviation Administration flight waiver opens.

Flight preparation will be completed in less than 2 hours. A checklist will be used to ensure that flight preparation is efficient and thorough. The team will have practiced these operations during test flights.

Functional testing

1.7. The launch vehicle shall be capable of remaining in launch-ready configuration at the pad for a minimum of 1 hour without losing the functionality of any critical on-board component.

All onboard electronics draw very little power, and can remain in launch-ready configuration for several hours.

Functional testing

1.8. The launch vehicle shall be capable of being launched by a standard 12-volt direct current firing system.

The rocket will use commercial, ammonium perchlorate motors that will ignite with 12-volt direct current.

Functional testing

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1.9. The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR), Tripoli Rocketry Association (TRA), and/or the Canadian Association of Rocketry (CAR).

The launch vehicle will use a TRA certified AeroTech K1100T motor, or a similar certified motor manufactured by AeroTech, Cesaroni, or Animal Motor Works.

Inspection

1.10. The total impulse provided by a launch vehicle shall not exceed 5,120 Newton-seconds (L-class).

The launch vehicle will use a K-class motor, which does not exceed 5,120 N-s total impulse.

Inspection

1.13. All teams shall successfully launch and recover a subscale model of their full-scale rocket prior to CDR. The subscale model should resemble and perform as similarly as possible to the full-scale model, however, the full-scale shall not be used as the subscale model.

The team will launch and recover a 2/3-scale (4” diameter) model of the full-scale rocket prior to CDR. The parts list for the subscale rocket is in the team budget. See the timeline for anticipated dates.

Inspection and functional testing

1.14. All teams shall successfully launch and recover their full-scale rocket prior to FRR in its final flight configuration. The rocket flown at FRR must be the same rocket to be flown on launch day.

The team will successfully launch and recover the full-scale (6” diameter) rocket prior to FRR in its final flight configuration. See the timeline for anticipated dates.

Inspection and functional testing

1.14.2.1. If the payload is not flown, mass simulators shall be used to simulate the payload mass.

The team plans to fly the payload in the full-scale demonstration flight.

Inspection

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1.14.2.3. If the payload changes the external surfaces of the rocket (such as with camera housings or external probes) or manages the total energy of the vehicle, those systems shall be active during the full-scale demonstration flight.

All payloads will be active during the full-scale demonstration flight.

Inspection

1.14.4. The vehicle shall be flown in its fully ballasted configuration during the full-scale test flight.

The vehicle will be flown in its fully ballasted configuration during the full-scale test flight.

Inspection

1.14.5. After successfully completing the full-scale demonstration flight, the launch vehicle or any of its components shall not be modified without the concurrence of the NASA Range Safety Officer (RSO).

The launch vehicle will not be modified after the full-scale demonstration flight without the concurrence of the NASA RSO.

Inspection

1.15. Each team will have a maximum budget they may spend on the rocket and the Autonomous Ground Support Equipment (AGSE). Teams who are participating in the Maxi-MAV competition are limited to a $10,000 budget while teams participating in Mini-MAV are limited to $5,000. The cost is for the competition rocket and AGSE as it sits on the pad, including all purchased components.

The team has budgeted $1500 for the competition rocket, and $3500 for the AGSE. Throughout development and construction of the rocket and AGSE, the team will be looking for ways to cut costs and stay within the $5000 total budget.

Inspection

1.16.1. The launch vehicle shall not utilize forward canards.

The launch vehicle will not use forward canards.

Inspection

1.16.2. The launch vehicle shall not utilize forward firing motors.

The launch vehicle will not use forward firing motors.

Inspection

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1.16.3. The launch vehicle shall not utilize motors that expel titanium sponges.

The launch vehicle will not use motors that expel titanium sponges.

Inspection

1.16.4. The launch vehicle shall not utilize hybrid motors.

The launch vehicle uses commercially available solid APCP motors.

Inspection

1.16.5. The launch vehicle shall not utilize a cluster of motors.

The launch vehicle uses only a single motor.

Inspection

Table 4 below shows the recovery subsystem requirements and the methods for verifying that those requirements have been met.

Table 4: Recovery Requirements and Verification

Requirement Design feature that satisfies the

requirement Verification

2.1. The launch vehicle shall stage the deployment of its recovery devices, where a drogue parachute is deployed at apogee and a main parachute is deployed at a much lower altitude.

Redundant Missile Works RRC3 altimeters will eject a drogue parachute at apogee, the payload bay at 1000 ft, and a main parachute at 800 ft.

Inspection

2.2. Teams must perform a successful ground ejection test for both the drogue and main parachutes. This must be done prior to the initial subscale and full scale launches.

Successful ground ejection tests will be performed prior to initial subscale and full scale launches.

Inspection

2.3. At landing, each independent section of the launch vehicle shall have a maximum kinetic energy of 75 ft-lbf.

Current simulations predict that all vehicle sections will land with less than 75 ft-lbf of kinetic energy.

The team will use simulation and test-flight data to calculate the kinetic energy of each vehicle section at landing.

2.4. The recovery system electrical circuits shall be completely independent of any payload electrical circuits.

There are no payload electrical circuits.

Inspection

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2.5. The recovery system shall contain redundant, commercially available altimeters. The term “altimeters” includes both simple altimeters and more sophisticated flight computers. One of these altimeters may be chosen as the competition altimeter.

The recovery system will contain redundant Missile Works RRC3 altimeters to deploy the parachutes. One of the RRC3 altimeters will be used as the competition altimeter.

Inspection

2.6. A dedicated arming switch shall arm each altimeter, which is accessible from the exterior of the rocket airframe when the rocket is in the launch configuration on the launch pad.

Both RRC3 altimeters will have separate external arming switches accessible when the rocket is in launch position.

Inspection

2.7. Each altimeter shall have a dedicated power supply.

Each altimeter will have a dedicated 9V power supply.

Inspection

2.8. Each arming switch shall be capable of being locked in the ON position for launch.

The arming switches will require a straight-edged screwdriver to lock them in the ON position.

Inspection

2.9. Removable shear pins shall be used for both the main parachute compartment and the drogue parachute compartment.

All parachute compartments are attached with #2 nylon shear pins.

Inspection

2.10. An electronic tracking device shall be installed in the launch vehicle and shall transmit the position of the tethered vehicle or any independent section to a ground receiver.

An Altus Metrum TeleGPS tracking device will be installed in the launch vehicle.

Inspection

2.10.1. Any rocket section, or payload component, which lands untethered to the launch vehicle shall also carry an active electronic tracking device.

The untethered payload compartment will have its own GPS tracking device.

Inspection

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2.10.2. The electronic tracking device shall be fully functional during the official flight at the competition launch site.

The GPS tracking devices will be fully functional at the competition launch site.

Functional testing

2.11.1. The recovery system altimeters shall be physically located in a separate compartment within the vehicle from any other radio frequency transmitting device and/or magnetic wave producing device.

The recovery system altimeters are installed in their own avionics bay, which is physically separated from the GPS transmitters. One GPS transmitter is installed in the payload compartment, and the other is in a separate compartment in the booster section.

Inspection

2.11.2. The recovery system electronics shall be shielded from all onboard transmitting devices, to avoid inadvertent excitation of the recovery system electronics.

The recovery system electronics will be shielded from the transmissions of the GPS transmitters, and from any other onboard devices that may adversely affect their proper operation.

Inspection

2.11.3. The recovery system electronics shall be shielded from all onboard devices which may generate magnetic waves (such as generators, solenoid valves, and Tesla coils) to avoid inadvertent excitation of the recovery system. 2.11.4. The recovery system electronics shall be shielded from any other onboard devices which may adversely affect the proper operation of the recovery system electronics.

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Manufacturing Plan

Critical to a successful mission is the detailed planning of the project. As such, the team has established a manufacturing schedule in the form of a GANTT chart. Included in the schedule is dates for construction of the vehicle, the AGSE, and the containment device, dates for static, component, and functional testing, as well as test flight dates and integration.

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Confidence and Maturity of Design

Propulsion System The design of the propulsion system is commonly used in amateur high-power rocketry. The selected motors are well tested and certified by the Tripoli Rocketry Association. Simulations predict more than sufficient thrust at lift-off to achieve stable flight. A wide selection of motors can reach the 3000 ft AGL target altitude depending on the final pad weight of the vehicle. Confidence in this system is high.

Structural and Aerodynamic Stability System This is another common design. BlueTube is a common, commercially available airframe material for amateur high-power rocketry. Its rigidity is sufficient for sub-Mach flights without reinforcement. The simulated static stability margin of 3.9 assures a stable flight. If test-flights reveal excessive weathercocking, mitigations can be implemented. Confidence in this system is high.

Recovery System The separation of the vehicle at apogee and deployment of the drogue parachute is another common design. The Missile Works RRC3 is a well known and reliable, commercially available deployment altimeter commonly used in amateur high-power rocketry. The simplicity of the design increases its chances of success. The ejection of the payload bay and deployment of the main parachute is innovative. We are uncertain how the ejection of the nosecone and attached payload bay at 1000 ft will affect the other sections of the descending vehicle, and how this may interfere with the deployment of the main parachute at 800 ft AGL. The piston system for deployment of the main parachute also requires testing. We are uncertain how the piston will be ejected smoothly from the airframe (without hitching or binding). Our confidence in this system is modest.

Project Risks

There are a variety of possibilities that pose a risk to the completion of the project. These risks are defined below and are given a risk assessment. Table 5 is a quantitative definition for the risks. This quantitative assessment table uses a scale from 1-25, where 1 is considered to be the lowest risk definition and 25 is the highest.

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Table 5: Project Risk Quantitative Assessment Likelihood Impact Level

5 Dire

4 Major

3 Medium

2 Minor

1 Trivial

5 Certain

25 20 15 10 5

4 Likely

20 16 12 8 4

3 Possible

15 12 9 6 3

2 Unlikely

10 8 6 4 2

1 Remote

5 4 3 2 1

Table 6 shows the qualitative assessment chart. The colors in the table correspond with a number range in Table 5, shown above.

Table 6: Project Risk Qualitative Assessment

Likelihood Impact Level

1 High

2 Medium

3 Low

A-High 1A 2A 3A B-Medium 1B 2B 3B C- Low 1C 2C 3C Table 7 defines the risk description for impact levels used in Table 6, shown above.

Table 7: Impact Level Definitions Description Definition 1-High Sever effect on the overall continuation of the project. May

require extensive mending if possible. 2-Medium Major effect on procession of project. Reversible with

significant effort. 3-Low Minor effect on overall project. Can be easily mended. Table 8 defines the likelihood descriptions used in Table B.

Table 8: Likelihood Definitions Description Definition A-High Chances of occurring are quite likely. B-Medium Possible likelihood of occurring. C-Low Chances of occurring are quite unlikely.

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Table 9 identifies and defines the risks involved in the project as well as their mitigations. Risks were given a quantitative and qualitative assessment pre- and post- mitigation.

Table 9: Project Risks and Mitigations

Risk Definition Pre-RAC

Mitigation Post- RAC

Building Time

Insufficient time for building adequate numbers of test models.

1A-20

Construction will begin as early as possible. Duplicate parts will be constructed simultaneously to increase efficiency.

1A-16

Functionality of Programming

AGSE design programming errors.

1B-16

Programming will begin early to allow for troubleshooting.

2B-12

Writing Time

Little time to compile all project information in a concise manner.

1B-15

The team will manage their time wisely.

1C-10

Low Budget Running out of money to fund project.

1C-6

The budget plan will be evaluated to identify sources of great expense. If possible, more cost effective alternatives for components will be found and used. If this is not possible, then further funding activities will take place.

2C-5

Low Resources

Running out of project materials during construction of scale, test, and final models.

2B-6

Materials will be purchased in sets of two.

3B-5

Functionality of Design

Project designs prove to be inefficient or incapable of completing objective.

2B-12

The design will be reevaluated periodically. Any complications will be analyzed and the design will be modified accordingly.

3B-4

Mass Statement The current mass estimate is based on RockSim and SolidWorks models, and on component spec sheets. The estimate is preliminary and likely to increase. Mass

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increases of up to 50% can be accommodated by using larger motors. The motors listed in the following table will all fit the vehicle’s planned 54 mm motor mount.

Table 10: Rocket Weight, Altitude, and Rail Velocity

vehicle weight w/out motor (lb)

motor simulated altitude (ft) 6 ft rail exit velocity

(ft/s)

19.63 AeroTech K1100T

3011 68

+25% AeroTech K1275R

3355 65

+33% AeroTech K1275R

3090 63

+50% Cesaroni K590-DT

3068 58

If the vehicle mass increases by more than 50%, the booster section can be redesigned to accommodate larger motors. Mission Performance Predictions

Mission performance criteria The primary mission performance criteria for the launch vehicle are:

stable flight 3000 ft. AGL apogee payload ejection at 1000 ft. AGL kinetic energy at landing for each section <75 ft-lbf

Vehicle Stability With the motor installed, RockSim gives the following estimates:

Center of Gravity (in from nose): 66.25 Center of Pressure (in from nose): 89.75 Stability Margin (diameters): 3.9

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Figure 16: Stability Diagram

Overstability RockSim estimates that the vehicle is overstable. Overstability can exacerbate weathercocking. Weathercocking can be a problem if it leads to increased horizontal velocity at apogee. A greater velocity at apogee makes the deployment of the drogue parachute more violent, and increases the risk of failure. However, RockSim may overestimate the stability of the rocket. When RockSim calculates the center of pressure, it ignores the forces acting on the body tube. But these forces can become significant on larger, longer airframes, and at angles of attack greater than 10 degrees. (See Apogee’s Peak of Flight Newsletter #239.) If severe weathercocking is observed during test flights, the team will consider reducing the vehicle’s stability margin by

reducing the length of the vehicle, so far as possible, or adding ballast to the booster section, or experimenting with a different fin design

Simulation Results The simulation results using RockSim are given in the following table. The simulation does not take into account the ejection of the nosecone and attached payload bay at 1000 ft. Therefore, the simulation

overestimates the descent mass below 1000 ft overestimates the velocity and kinetic energy at landing underestimates the drift of the rocket

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Table 11: Simulation Results

Ascension - Simulation results

Engine selection

[K1100T-None]

Simulation control parameters

Flight resolution: 800.000000 samples/second

Descent resolution: 1.000000 samples/second

Method: Explicit Euler

End the simulation when the rocket reaches the ground.

Launch conditions Altitude: 2000.00000 Ft.

Relative humidity: 50.000 %

Temperature: 64.990 Deg. F

Pressure: 29.9139 In.

Wind speed model: Slightly breezy (8-14 MPH) Low wind speed: 8.0000 MPH

High wind speed: 14.9000 MPH

Wind turbulence: Fairly constant speed (0.01) Frequency: 0.010000 rad/second

Wind starts at altitude: 0.00000 Ft.

Launch guide angle: 5.000 Deg.

Latitude: 34.000 Degrees

Launch guide data: Launch guide length: 72.0000 In.

Velocity at launch guide departure: 67.6563 ft/s

The launch guide was cleared at : 0.169 Seconds

User specified minimum velocity for stable flight: 43.9993 ft/s

Minimum velocity for stable flight reached at: 29.6623 In.

Max data values: Maximum acceleration: Vertical (y): 482.409 ft./s/s, Horizontal (x): 42.207 ft./s/s, Magnitude: 484.251 ft./s/s

Maximum velocity: Vertical (y): 470.1486 ft/s, Horizontal (x): 18.7792 ft/s, Magnitude: 470.2255 ft/s

Maximum range from launch site: 1559.30130 ft.

Maximum altitude: 3011.67292 ft.

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Recovery system data P: Main Deployed at : 57.821 Seconds

Velocity at deployment: 55.4566 ft/s

Altitude at deployment: 699.95153 ft.

Range at deployment: 816.01093 ft.

P: Drogue Deployed at : 13.688 seconds

Velocity at deployment: 6.5659 ft/s

Altitude at deployment: 3011.67288 ft.

Range at deployment: 34.62467 ft.

Time data Time to burnout: 1.601 Sec.

Time to apogee: 13.688 Sec.

Optimal ejection delay: 12.086 Sec.

Landing data Successful landing

Time to landing: 106.047 Sec.

Range at landing: 1559.30130

Velocity at landing: Vertical: -18.5245 ft/s, Horizontal: 14.2779 ft/s, Magnitude: 23.3884 ft/s

Altitude and Rail Exit Velocity Simulations predict that three different motors can achieve sufficient rail exit velocity (>44 ft/s) and altitude (~3000 ft AGL), depending on the final mass of the vehicle. These estimates are summarized in Table 10 above. Figure 17 compares the thrust curves for the AeroTech K1100T and K1275R motors (www.rocketreviews.com).

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Figure 17: Thrust Curves of AeroTech K1100T and K1275R

Figure 18 gives the thrust curve for the dual-thrust Cesaroni K590-DT, which would be used only if the rocket mass increases between 33% and 50%. This motor combines a fast- and a slow-burning propellant, which accounts for its unusual thrust curve.

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Figure 18: Thrust Curve of Cesaroni K590-DT

Kinetic Energy The following table summarizes the kinetic energy of each independent and tethered section of the launch vehicle. The kinetic energy of each section is well below the maximum 75 ft-lb at landing.

Table 12: Kinetic Energy of each Rocket Section

section descent weight of section (lb)

speed under

drogue (ft/s)

kinetic energy under drogue

(ft-lb)

speed at landing (ft/s)

kinetic energy at landing (ft-lb)

untethered payload

4.9 50 193 <20 <30

middle 7.2 50 279 <20 <43

booster 8.8 50 343 <20 <53

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Drift from Launch Pad Only rough estimates of the vehicle’s drift from the launch pad are possible at this time. The RockSim flight simulations assume that

1. the rocket is launched at a 5-degree angle 2. all parts of the rocket descend and drift together under the drogue parachute 3. all parts of the rocket descend and drift together under the main parachute 4. the vehicle is not buoyed by a thermal column

The third and fourth points introduce some error into the estimates. At 1000 ft AGL, the payload bay is ejected, and it descends untethered under its own parachute. The tethered sections of the rocket descend under a main parachute that is deployed at 800 ft AGL. These facts are not accounted for in the simulation. However, this error should not be very great, since the distance to the ground is small (<1000 ft), and both parts of the rocket should descend at roughly equal speeds (even under separate parachutes). Thermal columns of air at low altitudes can buoy a vehicle under parachute and extend its drift. In simulations that included random thermal columns, drift was increased by up to 1000 ft. Because thermal columns are random and infrequent, they are not accounted for here. The following table gives rough baseline drift estimates, which assume vehicle drift is not affected by thermal columns.

Table 13: Drift from Launch Pad (all sections)

wind speed (mph) drift at 1000 ft

AGL (ft) total drift at landing (ft)

0 525 525

5 662 878

10 760 1132

15 970 1560

20 1070 2287

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Interfaces and Integration

Payload Integration

Figure 19: Dimensional drawing, isometric view, and side view of the payload containment bay.

The payload will be integrated into the vehicle through the payload containment bay shown in Figure 19. The payload will fall into the payload containment area when it is inserted into the rocket. The opening that the payload falls into will be chamfered on the long sides to ensure that the payload falls into place correctly. However, the payload is not secured until the vehicle is lifted upright. When the vehicle is lifted upright, gravity causes the payload to fall into the payload slot, which will hold the

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payload through the flight. The payload slot can be seen to the left of the payload containment area in Figure 19. This containment bay will be secured to a wood sled that slides onto all-thread rods via four all-thread attachment blocks, which are pointed out in the dimensional drawing in Figure 19. This assembly will be between two bulkheads that will prevent it from sliding around on the all-threads during flight. The tracking device will be located on an electronics sled that will be above the payload containment area. It is shown to the right of the payload containment area in Figure 19. Most of the assembly will be constructed using birch plywood, although some aluminum will be used. The components are all rated to withstand the stresses that will be present during boost. Preliminary analysis indicates that with the payload, 30 lbs of stress is added to the bulkhead. This bulkhead is currently planned at 0.5” thick and will withstand this force. The walls of the payload containment area are reinforced with aluminum, so they will be able to withstand the stress exerted on them by the payload. The figure also shows the dimensions for the components that comprise the payload containment device. These dimensions have been set to fit the current design of the vehicle between the payload containment bay and the nosecone.

Figure 20: Drawing of payload containment device integration into the main body.

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The payload containment device will be made to fit into the payload section (forward section) of the vehicle. The containment device will be made to slide into the payload section from the fore end. The bottom bulkhead of the containment device will have a section flattened so that it can slide past the spring-loaded doors in the airframe. The payload containment area must line up with the doors or the payload cannot be inserted. The doors are spring-loaded so that no electronics are necessary to operate it. The payload will be pushed through the door by the AGSE in order to contain it since the doors will not open under the weight of the payload by itself. Once the payload containment device has been slid into place, the assembly will be secured to the airframe of the vehicle through a series of locking bolts. These locking bolts will slide through the airframe of the payload section and into the bulkheads of the payload containment device. The locking bolts can be seen in Figure 20 below and above the payload doors. Once the assembly is locked inside the vehicle, the nosecone can slide into the body of the vehicle. The nosecone will also be locked with bolts that extend through the airframe so that it does not come off at any point during the flight. The bottom of the payload section will be attached to a parachute and the recovery electronics will ensure that the sections separate and the parachute deploys at the proper times.

Payload Containment Preliminary Checklist Mount GPS electronics Mount recovery electronics Check to make sure that all electronics have proper connections. Secure all loose electronics with zip ties. Attach electronics to switches on the airframe Double check functionality of payload doors. Slide payload containment device into body tube Secure at bottom with locking bolts Slide nosecone into body tube Secure payload containment device and nosecone with top locking bolts

Internal Interfaces Throughout the vehicle, components will need to be in contact in order to carry out flight procedures. A list of the internal interfaces of the vehicle is shown in the table below. This table shows the components that are interfacing and describes how they interface with each other.

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Table 14: Internal Interfaces of the Vehicle

Internal Interfaces Components Interface

Booster and middle section

These components connect through the shock cord that attaches to the drogue parachute. The shock cord will be attached to a bulkhead from each section so that the sections cannot separate in flight.

Avionics and middle section

The avionics will connect to the middle section through an electronics sled. This sled will be placed in its own bay at the bottom of the middle section and will rest between two bulkheads.

Recovery electronics and parachutes

The recovery electronics will interface with the recovery parachutes through black powder charges. E-matches will be connected to the recovery electronics and these will ignite the black powder charges that separate the sections and eject the parachutes.

Middle section and main parachute piston

The main parachute will not be connected to the forward section. It will be connected to a piston in the middle section through a shock cord. Ejecting the piston will cause the parachute to eject.

Payload recovery electronics and payload section

The payload recovery electronics will be placed on the backside of the payload containment area. The electronics will be secured to a wood piece as shown in Figure 19.

Payload recovery electronics and parachute

The payload recovery electronics will be connected to an E-match. This E-match will be connected to a black powder charge. This charge will cause the complete separation of the payload section with the rest of the main vehicle and will cause the parachute to deploy.

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Payload GPS electronics and payload section

An electronics sled has been designed to extend partly into the nosecone. This electronics sled will hold the GPS electronics. The GPS electronics will be connected to the payload section containment device above the payload containment area.

Payload doors and payload section airframe

The payload doors will interface with the airframe of the vehicle through spring loaded hinges. These hinges will allow the inside of the airframe to be accessed by the AGSE.

Launch Vehicle and Ground Interfaces The launch vehicle will interface with the ground through wireless communication. The rocket will be made to communicate to the ground in two ways. One interface with the ground will be between the TeleGPS and the ground station. The TeleGPS will track the vehicle for recovery. The TeleGPS uses APRS to send latitude and longitude coordinates to the ground station. The other interface with the ground will be between the EM-506 GPS and the ground station. The EM-506 GPS will connect with the ground station through an XBee Pro 900 transceiver. The XBee will open a serial communication line with another XBee that is connected to the ground station and the latitude and longitude coordinates will be transmitted wirelessly. Launch Vehicle and Ground Launch System Interfaces The vehicle will have two large 1.5 inch rail buttons so that it can slide onto a launch rail. The simulations dictate that the launch rail should be between 4 and 6 feet so that the minimum stable speed can be achieved when exiting the launch rail. The igniter will be manually installed inside the motor and the leads of the igniter will be connected to a power source that will ignite the motor when ready.

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Safety

Launch Checklist Table 15 below shows the preliminary checklist for final assembly and launch procedures.

Table 15: Preliminary Checklist Step Verified

By Verified By

Date Time of Verification

Final Verification by Safety Officer

Location Setup

1 Unload Rocket and Equipment

2 Establish base of operations

3 Set up work tables

4 Layout rocket sections for set-up

Nose Cone and Payload Setup

1 Assemble electronics

2 Check the charge on batteries

3

Check wiring to ensure that components are connected properly.

4 Install deployment charges

5 Check continuity of igniters

6 Insert entire assembly into the forward section

7 Ensure bottom locking bolts are secure

8 Slide nosecone into airframe

9 Ensure top locking bolts are secure

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N/A This section may be expanded once construction begins.

Drogue Bay

1 Insert black powder charge in separation chamber

2 Ensure coupler firmly attaches drogue bay and main bay

3 Check for possible snags in bay

4 Correctly pack drogue parachute

5 Ensure shock cord is securely harnessed onto U-bolts

Avionics Bay

1 Check for charged battery

2 Ensure avionics systems are wired correctly

3 Check that the avionics are securely fastened

4 Ensure that the arming switches engage all subsystems

5 Secure bulkheads

N/A This section may be expanded once construction begins.

Main Bay Setup

1 Insert black powder charge in separation chamber

2 Ensure coupler firmly attaches main bay to booster section

3 Check for any possible snags in bay

4 Correctly pack the main parachute

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5

Ensure shock cords are securely harnessed to the U-bolts

Booster Section Setup

1 Load motor (refer to manufacturer procedure)

Fins

1 Check to see if fins are secured to the rocket

2 Ensure fins are not damaged or cracked

Launch Pad

1 Bring rocket out to launch

2 Set up rocket on the guide rail

3 Install igniters

4 Connect launch equipment to the igniters

5 Check for continuity

6 Clear the launch area to safe distance

Safety Officer Alex will serve as the team’s safety officer. Alex is TRA Level 1 certified and will be First Aid certified in the near future. The safety officer’s responsibilities in regards to the vehicle include safety analysis, risk mitigation, creating launch procedure checklists, and communication on safety awareness.

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Preliminary Hazard Analysis Table 16 below shows the possible failure modes of the vehicle and the mitigations for those failures.

Table 16: Vehicle Failure Modes

Risk Consequence Pre-RAC

Mitigation Post- RAC

Center of gravity is too far aft

Unstable flight 2B-12

Add mass to the nose cone

2B-9

Center of pressure is too far forward

Unstable flight 2B-12

Increase the size of the fins to lower the center of pressure

2B-9

Fin failure Unstable flight, further damage to the rocket

1C-12

Careful construction to ensure proper fin attachment

1C-8

Shearing of airframe

Loss of rocket 1C-12

A material with high shearing strength will be used

1C-8

Premature rocket separation

Failure to reach target altitude, failure of recovery system

3A-8

Check the shear pins before launch, test the timers in test launches, calculate the required mass for black powder charges

3A-6

Centering ring failure

Loss of rocket 1A-15

Check construction of centering rings for a good fit, check for damage to centering rings pre-launch and post recovery.

2B-6

Bulkhead failure

Damage to payload, avionics, failure of recovery

2C-5

Proper construction, extensive ground testing of removable bulkheads

2C-4

Nose cone failure

Flight instability, damage to payload bay, unable to re-launch rocket

2C-5

Strong nose cone constructed from fiberglass

2C-4

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Table 17 below shows the failure modes of the propulsion subsystem as well as the mitigations for those failure modes.

Table 17: Propulsion Failure Modes

Risk Consequence Pre-RAC

Mitigation Post- RAC

Motor ignition failure

Failure to launch, unstable flight, change in trajectory.

1B-16

Redundant igniters, proper motor assembly

2B-12

Motor failure Unstable flight, failure to reach target altitude, loss of motor casing

1B-15

Assembly of motors by certified members only

2B-10

Exploding of the motor during ignition

Loss of motor casing, loss of rocket

1B-15

Static testing of motors 2B-10

Motor igniter not reaching the end of the motor

Prevention of complete motor burnout or lack of motor ignition influencing the trajectory of the rocket

1C-12

The igniters will be marked with a permanent marker at the length of the motor

2C- 3

Motor mount failure

Motor launches into the body of the rocket, damage to payloads, loss of rocket

1C-12

Proper construction of the motor mount

2C-4

Premature burnout

Failure to reach target altitude

2B-6

Static testing of motors 2C-4

Improper transportation or mishandling

Unusable motor, failure to launch

1C-6

Motors to be handled by certified members only, motors to be stored properly

2C-3

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Table 18 below shows the recovery failure modes and the mitigations for those failures.

Table 18: Recovery Failure Modes

Risk Consequence Pre-RAC

Mitigation Post- RAC

Rapid Descent Damage to airframe and payloads, loss of rocket

1B-16

Redundant altimeters, verification testing of the recovery system, simulation to determine appropriate parachute size

1C-12

Parachute deployment failure

Loss of rocket, extreme damage to rocket and all components

1B-16

Ground test of parachute deployment

1C-12

Parachute separation

Loss of parachute, loss of rocket, extreme damage to rocket and all components

2A-15

Strong retention system, load testing

2B-12

Parachute tear

Damage to rocket, loss of parachute, rapid descent resulting in an increased kinetic energy

2B-12

Safety check the parachute for damage, clear parachute bays of any possible defects, properly pack the parachutes

2C- 4

Parachute melt

Damage to rocket, loss of parachute, rapid descent resulting in an increased kinetic energy

1C-10

Proper protection from ejection charges, ground testing of recovery system

2C-5

Slow Descent Rocket drifts out of intended landing zone, loss of rocket

2B-9

Verification testing of recovery system, simulation to determine appropriate parachute size

2C-5

Listing of Personnel Hazards and Safety Hazard Data A thorough evaluation of the possible hazards associated with the vehicle has been made with respect to the user as well as the environment. The hazards associated with the vehicle are included in the material safety data sheets, which pertain to every aspect of the vehicle and its operation. This data is introduced in Appendix D of the document. Precautionary measures are being taken to ensure that no harmful or explosive substances will be misplaced or misused. A listing of personnel hazards and evidence of understanding of safety is provided in the sections below.

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Launch Site Safety Before launch day, the team will receive training in hazard recognition and accident avoidance; on the day of the launch, the safety officer will perform a safety check on the motor, payload, and recovery subsystems. The team will conduct a safety briefing both before and after each launch where the recognized hazards will be discussed as well as methods for mitigation.

Certification An individual must be certified by either the NAR or TRA to purchase and use high-power rocket motors. The team leader, Aaron, and the team’s mentor, Rick Maschek, are TRA Certified Level II. The certified members of the team are aware of the risks of high-power rocketry and will help the safety officer ensure a safe launch environment.

Motor Handling and Storage High-power rocket motors contain highly flammable substances such as black powder or ammonium perchlorate. Therefore, they are considered to be hazardous materials or explosives for shipment purposes by the US Department of Transportation (DOT). The team is aware of and will follow all DOT regulations concerning shipment of hazardous materials. These regulations are contained in the Code of Federal Regulations (CFR) Title 49, Parts 170-179 and specify that it is illegal to send rocket motors by commercial carriers or to carry them onto an airliner. NFPA 1127 Section 4.19 contains the storage requirements of motors over 62.5 grams. The team will store all high-power rocket motors, motor reloading kits, and pyrotechnic modules at least 7.6 meters (25 feet) from smoking, open flames, and other sources of heat. The Tripoli Rocketry Association and the National Association of Rocketry have adopted the National Fire Protection Association (NFPA) 1127 as their safety code for all rocket operations. A general knowledge of these codes will be required of all team members. All members of the team will demonstrate competence and knowledge in handling, storing, and using high-powered motors. These include all reloadable motors, regardless of power class, motors above the F-class, and those which use metallic casings.

Adhesive Safety Much of the construction of the vehicle and payloads require the use of epoxy. Any use of epoxy will be done on construction or lab tables in a well-ventilated area and all team members present are required to wear dust masks and gloves. Acetone or isopropyl alcohol will be available along with a fully equipped first aid kit in the event that there is any contact of adhesive to skin.

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Environmental Concerns Table 19 below shows the environmental hazards that are present during the launch of the vehicle.

Table 19: Environmental Hazards Hazards to the

Rocket Description

Rocket Landing in Wheat Field

On descent, the rocket may land in a nearby wheat field. This will make locating the rocket difficult.

Wind Blowing Parachute

On descent, the winds may catch the rocket and blow it in an undesired direction or location.

Rocket Lands in nearby road

On descent, the rocket may land in the middle of a road. This would both disrupt traffic and put the rocket in danger.

Heavy Winds Interfere with Launch

The wind in the area may begin to pick up and put the launch process at risk. In this case, the launch may be delayed or canceled altogether.

Force of wind opens payload doors

At any point during the rocket’s flight the payload doors may be at risk of opening due to forces caused by strong winds. If the payload compartment opens, the stability of the vehicle may be compromised.

Electronics landing in water

On descent, the rocket may land in water. If submerged, the electronics within the rocket would be at risk.

Hazards to the Environment

Description

Rocket booster section lands in water

On descent, the rocket may land in a location with water. If the booster section of the rocket is submerged, chemicals from the motor can pollute the water.

Rocket hits a bird During the launch process, a flock of birds may be flying overhead in such a manner that the rocket blows through them. The rocket may harm or cause loss of life among the wild life.

Bird hits the parachute

On descent, a flock of birds may be flying by and interact with the parachute in a way that could compromise the functionality of the parachute.

Falls into air vent On descent, if there are any nearby structures, the rocket may land into or on top of an air vent. This may cause damage to the rocket or cause a polluted environment from booster section chemicals.

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IV. AGSE Criteria Selection, Design Review, and Verification

Design Review

The AGSE consists of a single Autonomous Rover system. The goal of the system is

to capture and contain the provided payload inside the payload bay of the launch

vehicle. The fundamental order of operations for the system will follow the

subsequent events.

AGSE activation

Payload location

Navigation to Payload

Payload retrieval

Rocket location

Navigation to Rocket

Payload containment

The functional requirements for the successful completion of the AGSE mission

objectives are outlined in Table 20, along with the corresponding subsystem that

addresses those functional requirements.

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Table 20: AGSE Functional Requirements

Subsystem Functional

Requirements Selection Rationale

Selected Concept

Characteristics

Body

The AGSE system must be self-contained and mobile over Martian-like terrain

In order to obtain the payload, the body will need to transport the AGSE subsystems to the payload

An aluminum structure to support all AGSE subsystems in a mobile unit

A lightweight durable structure (chassis)

Camera Subsystem

Determine the angle of rotation necessary to orient the AGSE towards or away from the payload, rocket or a hazard.

The AGSE must have a means of finding the payload, as well as differentiating it from hazards.

Use image analysis to determine the orientation of an object relative to the AGSE.

Collect and analyze image data.

Ranging Subsystem

Using the emission and detection of laser pulses, the distance between objects (hazards, payload, launch vehicle) and the AGSE will be determined

The distance from the AGSE to the payload will be necessary to determine how far the system needs to travel. This distance must be determined without ultrasonic sensors.

To determine the distance from the AGSE to the payload using a laser rangefinder

Emit and detect laser pulses

Payload Retrieval

Subsystem

To physically retrieve and transport the payload from its location to the launch vehicle

Retrieving the payload requires mechanical processes (gripping and lifting) with precise movement

A robotic arm Physically grip and lift the payload.

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Main Computer

To manage all AGSE subsystem operations and data acquisition, and run the custom designed on board software.

Each subsystem will acquire large volumes of data that will need to be communicated to other subsystems. Therefore a high-powered processor must be utilized

To acquire digitized data and operate all AGSE subsystems

Collect, analyze and transmit data.

Power Supply

To provide power to all AGSE subsystems

All subsystems will require a supply of power.

A network of power banks

Lithium Polymer Power Banks

Subsystems Required to Accomplish the AGSE objectives

Body

The body must support and transport all AGSE systems and subsystems over terrain

comparable to Martian terrain. The body will be comprised of three major

components, an aluminum chassis, a “rocker bogie” Suspension, and a six wheel

servo controlled drive. The different components of the body are displayed in Figure

xx and Figure xx

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Figure 21: Overall AGSE Dimensions

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Figure 22: AGSE Exploded View

Chassis The chassis is the structural framework for the entire AGSE. All subsystems and their components will be mounted onto it. The chassis will be machined from 1/8” 6061-T6 aluminum sheet metal, which was selected for its light weight, strength, availability and low cost. The chassis assembled using a combination of welding and aluminum L-brackets and bolts to form the design in Figure xx.

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Figure 23: Chassis Isometric and Exploded View

Suspension The suspension is the system that connects the chassis to the wheels and allows motion between the two. For the AGSE, the suspension is also designed to allow for the vehicle to traverse uneven terrain, similar to Martian terrain. For this reason, a Rocker Bogie design has been selected. The rocker bogie suspension is modeled after NASA’s design used on the Mars Pathfinder and the Mars Exploration Rover. The design uses two rocker arms on each side of the chassis, which are able to move up and down independent of one another. Refer to figure xx. This allows the AGSE to move over uneven terrain and even crawl over objects without tipping over. At the ends of the “rocker arms” are a total of six “bogie arms”, which refer to the links which the drive and steering motors are attached to. The rocker bogie suspension will be constructed from rectangular 2” x 2” x 1/8” hollow 6061-T6 aluminum alloy tubing. The material was selected for its light weight, strength, availability and low cost. The parts will be cut using a saw or end mill, and welded or bolted together to form the structure of the suspension. The bogie arms will each support the motor/servo drives.

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Servo Drive Wheels and wheel-related equipment will be attached to the end of each bogie arm. Hollow tubing was chosen specifically for wiring purposes; wiring will be required to attach the drive motors to their respective microcontrollers, and this wiring can easily be stored inside the suspension framing with this type of material. The drive and steering motors will be attached to the rods at the end of each bogie arm. There will be two fully driven wheels on each side of the AGSE, for a total of four powered wheels. The center wheel on each side will simply be an idle wheel. The front and back wheels will contain one motor and one servo; the motor for forward and backward motion, and the servo to allow the wheels to pivot up to 45 degrees in either direction. The pivoting of the wheels will allow the AGSE to turn in place, up to a full 360 degrees. Table 21 details the components of the AGSE body, their functional requirements, the selection rational taken into consideration for the selected concepts, and their characteristics.

Table 21: AGSE Functional Requirements

Component Functional

Requirements Selection Rational

Selected Concept

Characteristics

Chassis

To physically support all AGSE subsystems

In order to transport the AGSE subsystems, they must all be mounted to a single unit

An aluminum chassis to physically support all AGSE subsystems

Lightweight and durable

6 Wheel Servo/Motor Drive

To provide mobility to the AGSE

Servos provide a system which allows for high control in steering and drive.

6 Wheel servo/motor drive

Controlled by the central computer and various motor controllers

Rocker Bogie Suspension

To allow for the AGSE to drive on uneven terrain, similar to a Martian environment.

The Rocker Bogie concept is already a well-established design in use on current Mars rover missions.

Rocker bogie suspension

Pivoting of the bogie arms allows for the AGSE to traverse uneven terrain.

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Camera Subsystem The camera subsystem is designed to track the position of specific objects relative to

the AGSE. Such objects include the payload, the launch vehicles and hazard. The

camera system will use the angular position of the camera module to determine the

necessary compensation in the angular orientation of the AGSE to avoid or approach

an object. The system incorporates a PixyCMUCam5 camera module, a Mini Pan-Tilt

head, an Arduino Uno microcontroller, an Adafruit Data logging shield and an SD

card. The Pixy camera will be mounted onto the Mini Pan-Tilt head, which will be

secured to a cylindrical camera mast constructed of 6061-T6 aluminum. Both the

Pixy cam and the Mini Pan-Tilt head will be connected to the Arduino for data

acquisition and processing.

Table 22 details the components of the camera subsystem, their functional requirements, the selection rational taken into consideration for the selected concepts, and their characteristics.

Table 22: AGSE Functional Requirements

Component Functional Requirements

Selection Rationale

Selected Concept Characteristics

PixyCamCMU5

To take images of the environment so they can be scanned for the payload and hazards using image analysis

The AGSE design must be able to be implemented on a Martian environment. Digital image capture would work in such an environment.

To use image capture and analysis to determine the AGSE’s orientation with respect to specific objects.

Color image capture.

Camera Pan-Tilt Head

To determine the angular compensation in the AGSE’s trajectory in order to pursue or avoid specific objects

The Pan-Tilt head is able to measure the angle of rotation in its pan and tilt servos.

To use measurements of the degree of “pan” and “tilt” in the Pan-Tilt Head to calculate the angle of rotation necessary for the AGSE to avoid or approach the hazard/payload

Measurably pan and tilt the camera head.

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Arduino Uno Rev3

To run the custom designed on board software for each subsystem. Data acquisition, processing and transmission will be essential.

The Camera subsystem will acquire large volumes of data and will have to run edge detection and color detection algorithms on many images. Therefore it will require a high power processor.

To digitize data for processing and use in other systems and subsystems

Collect, analyze and transmit data.

Transcend 8gb SG Class 10 SD card

Record and store the images from the Pixy cam for analysis.

Large amounts of image data will be collected. Hence a device must be used to store them.

An SD card to store image files.

Read/write rates of up to 20/18 MB/s. Suitable for images and video.

Adafruit Data Logging Shield

Interface between the SD card and the Arduino.

The Arduino Uno does not come with an SD card interface out of the box, therefore one must be added.

A preassembled Arduino compatible shield for SD card interfacing

Works with FAT16 or FAT32 formatted cards.

Upon system activation the Camera subsystem’s pan tilt head will pan the camera

module 360 degrees. During the panning sequence the Pixy camera will collect a

video sample of the environment. The images from the video can be analyzed for the

payload and hazards using edge detection and color algorithms. Using the time of

the frame and the angular velocity of the pan tilt head, the angle of rotation from the

front face of the AGSE will be calculated by the Arduino microcontroller. This data

will be transmitted to the main computer so that the AGSE can rotate itself to avoid

the objects detected as hazards, and face the payload for distance measurements.

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Laser Ranging Subsystem The laser ranging subsystem (LRS) is designed to determine the distance from the

AGSE to the payload and the rocket. The LRS works by using the emission and

detection of laser pulses to calculate the distance from the unit to the object it is

ranging. The module uses measurements of the time it takes to detect the incoming

laser and the speed of light to perform the distance calculations. The LRS will use a

Lightware SF02/F laser ranging module, a Polulu 200 step bipolar Stepper motor, a

Polulu A4988 stepper motor driver and an Arduino Uno R3 microcontroller. The

SF02/F module interfaces directly with the Arduino Uno and is mounted onto the

stepper motor. The stepper motor rotates the entire laser module in steps allowing

for multiple distance points. The stepper motor is connected to the A4988 motor

driver, which is a breakout board that allows for current limiting and over

current/over temperature protection. The A4988 is connected directly to the

Arduino.

Table 23 details the components of the laser ranging subsystem, their functional

requirements, the selection rational taken into consideration for the selected

concepts, and their characteristics.

Table 23: AGSE Functional Requirements

Component Functional Requirements

Selection Rational

Selected Concept

Characteristics

Lightware SF02/F Laser Ranging Module

To emit and detect laser pulses and measure the emission-detection time for ranging calculations

The AGSE design must be viable in a Martian environment. Lasers are electromagnetic waves that could be used in a Martian environment

To precisely scan the surface of the payload in order to determine its distance.

Emit and detect laser pulses

Polulu 200 Step Bipolar Stepper motor

Rotate the laser in incremental “steps”

Using a stepper motor allows for a “sweep” of data points increasing the reliability of the LRS

To rotate the laser module in 0.45 degree steps on a stepper motor.

Position multiple laser pulses.

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Polulu A4988 stepper motor driver

Interface between the stepper motor and the microcontroller

Motors and servos have the potential to draw more current than the Arduino is meant to output. A driver is used to protect the Arduino board.

To use a breakout board which regulates the current. It also serves as an interface between the motor and the Arduino.

Provides the Arduino with over-current and over-temperature protection.

Arduino Uno R3

To run the custom designed on-board software. Data acquisition, calculations and transmission will be essential.

The raw data provided by the laser module requires calculations to obtain distance, which must be transmitted to other subsystems.

To digitize data, perform calculations and relay distance data to the other subsystems

Collect, analyze and transmit data.

After the AGSE has been centered to face the payload by the camera system, the LSR

subsystem will activate. On activation, the SF02/F laser module will emit a laser

pulse. The pulse will be detected by the by the modules laser receiver. The module

will also measure the time it takes for the laser to travel to and from the payload.

This data will be transmitted to the Arduino Uno as an ASCII encoded string through

the modules serial port. In order to ensure that the LSR is measuring the distance to

the payload, the LSR will emit a sustained laser beam for 2 seconds. The camera

subsystem’s edge and color detection will be used to determine if the white payload

is being hit by the LSR’s red laser. If the laser does not hit the payload, the stepper

motor will rotate the laser one step at a time until the camera module verifies that

the laser has made contact with the payload. Once the correct data is obtained, the

subsystems microcontroller will transmit the data to the main computer.

Payload Retrieval Subsystem The payload retrieval subsystem is designed to physically pick up the payload, secure it for transit to the rocket, and contain the payload in the rockets payload bay. The payload retrieval subsystem will use a Lynxmotion AL5D robotic arm, a Flexiforce pressure sensor, and an Arduino Uno microcontroller. The arm has 20 inches of reach and consists of four servos giving it four degrees of freedom. The first servo is the HS-805BB shoulder servo. It will be modified to sweep the first section of the arm from 0 to 360 degrees. The second servo is an HS-755HB servo. It will be attached to the end of the first section of the arm, and it will sweep the

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second half from 0 to 180 degrees. Finally, the wrist grip at the end of the arm has two servos of its own. The wrist can rotate a full 360 degrees axial to arm it is mounted on. The grip servo will only rotate enough to securely close the grip. The relationships between the servos are better illustrated in Figure 24.

Figure 24: The Mechanical Arm

The camera module will be used to keep the grip centered over the payload and will be fixed to the gripper. The Flexiforce sensor is used to determine when the grip is securely closed around the payload. The Arduino Uno is the microcontroller for the subsystem and is responsible for running the on board custom software for image analysis and servo control. Table 24 details the components of the payload retrieval subsystem, their functional requirements, the selection rational taken into consideration for the selected concepts, and their characteristics.

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Table 24: AGSE Functional Requirements Component Functional

Requirement Selection Rational

Selected Concept

Characteristics

HS-805BB Shoulder Servo

Rotate the shoulder of the arm centered from the base as necessary, with a range of 270 degrees.

To rotate the shoulder of the arm while retaining precision and programmatic control.

HS-805BB servo included in the Lynxmotion kit. It will be modified to have a range of 0 to 270 degrees.

343 oz-in torque, 270 degree range.

HS-755HB Elbow Servos

Connect to the aluminum brackets to rotate the arms with a range of 180 degrees

To rotate the shoulder of the arm while retaining precision and programmatic control.

HS-755HB servo included in the Lynxmotion kit.

183 oz-in torque. 180 degree range

HS-645MG Wrist Servo

Rotate the grip circularly with a range of 180 degrees

To rotate the wrist of the arm while retaining precision and programmatic control.

HS-645MG Servo included in the Lynxmotion kit.

133 oz-in torque. 180 degree range.

HS-422 Gripper

Securely close a grip around the payload for

To close the grip of the arm while retaining precision and programmatic control

HS-422 Gripper included in the Lynxmotion kit

57 oz-in.

OV7670 Camera Module

Collect images for edge detection to determine if the payload is off center from the HS-422 gripper.

Edge detection is the most efficient means of searching for an object without relying on ultrasonic sensors.

Arduino compatible camera module to collect images for analysis

Collects image data.

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Flexiforce Pressure Sensor

Determine whether the gripper is sufficiently secure around the payload

A pressure sensor was selected because of the ease of measuring the pressure force in the grip

Flexiforce pressure sensor which produces change in resistance with respect to the change in pressure applied to the sensor.

Measures pressure between payload and gripper.

Arduino Uno R3

Collect and analyze data from the camera module using custom made software. Transmit data for any necessary adjustments in the arms

The servos need to be controlled and precisely adjusted to retrieve the payload.

To digitize, analyze and transmit data to guide and control the Lynxmotion arm.

Collect, analyze and transmit data.

Once the AGSE has reached the payload the payload retrieval subsystem will activate. On activation the camera module will begin collecting images. Since the camera module is fixed to the gripper, the gripper will always occupy the same pixel space in the image. Edge detection will be used to determine the position of the payload in the gripper as the gripper is moved towards the payload. If the arm is offset, the images can be used to determine which direction the arm is offset. The arm will be programed to compensate for this error until the image is centered in the frame. This process will continue until the gripper is around the payload. This will also be verified using image analysis. Once the gripper is around the payload, the arm will be programed to close the gripper. As the gripper closes, it will apply pressure to the Flexiforce sensor. When enough pressure has been applied so that the gripper will not drop the payload, the arm will lift the payload. This specific pressure value will be determined experimentally. The AGSE is then clear to proceed to the rocket. The pressure sensor will continue to be read by the Arduino to ensure that the pressure does not change in transit. When the AGSE reaches the rocket the same process of image analysis and servo control will be used to insert the payload into the launch vehicles payload containment bay.

Main Computer The main computer is designed to slave all of the system and subsystem

microcontrollers, and perform any intensive calculations. The controller used for

this system will be an Arduino Mega which was selected for its high processing

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power. It will receive different kinds of encoded data from each subsystem which

must be analyzed to determine what functions must be carried out in the different

subsystems. With 54 digital input/output pins and 16 analog input/output pins, the

Arduino Mega has sufficient interfacing capabilities. It will be programmed in a

modified, open-source version of C/C++ The main computer will also feature a

master switch and a pause switch. The master switch will powers on the entire

system. The pause switch will temporarily disable all autonomous AGSE

subroutines.

Power Supply A system of lithium polymer power banks will provide power. An AllPower 50,000

mAh power bank will power all of the microcontrollers on the AGSE. Two Anker

Astro Pro Multi-Voltage power banks, rated for 20,000 mAh each, will power the

motors. These power banks will be used because of their large power capacity and

their ability to be toggled between 5V, 12V, 16V and 19V. This voltage selection

allows for a wider range of motors to drive the AGSE.

AGSE Electrical Schematics

Figure 25: Electrical Schematics for the AGSE.

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Figure 26: Motor Driver Schematics

The AGSE procedures will be controlled by the Arduino Mega. This component receives data from the two Arduino Uno’s that operate the sensors. The Arduino will use this data to determine what to do. It will send commands to the motor drivers shown in figure 26, or to the robotic arm shown in figure 25.

Performance Characteristics for the Systems and Subsystems

Laboratory testing will be the central approach and provide the most favorable

circumstances to test, modify, improve and authenticate the design and

effectiveness of the system and its subsystems. All components of each subsystem

must be tested and verified to guaranty a successful mission outcome.

Table 25 details the performance characteristics and evaluation and verification

metrics for each subsystem of the AGSE.

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Table 25: Subsystem Performance, Evaluation, and Verification

Subsystem Performance Characteristics

Evaluation and Verification Metrics

Body Supports and transports all AGSE subsystems

The body will be simulated in Solid Works for mechanical soundness and will be field tested using equivalent masses.

Camera Subsystem

Capture and analyze images

The camera subsystem must be able to determine if a hazard or the payload is in its view and must differentiate between the two. It will be tested in laboratory conditions for accuracy and consistency using different shaped and colored objects. The camera system will then be field tested as well.

Ranging Subsystem

Emit and detect laser pulses

The laser pulses must be reflected from the payload surface to the laser module. Laboratory testing on different materials will be conducted with an emphasis on PVC.

Payload Retrieval Subsystem

Locate, grip and pick up payload for transport.

The robotic arm will be tested for its ability to locate, and successfully lift the payload, using a payload prototype of the same mass and material.

Main Computer

Collect, analyze and transmit data between all subsystems

Testing will be carried out to ensure proper collection, analysis and transmission of data from and to all subsystems.

Power Supply

Provide power to all AGSE subsystems

Each component of the power supply subsystem will be tested for its ability to consistently supply power to its corresponding components in the AGSE using voltage and current measurements.

Component Level Table 26 details the performance characteristics and evaluation and verification

metrics for the individual components of the body.

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Table 26: AGSE Body Components

Component Performance Characteristics

Evaluation and Verification Metrics

Chassis Structurally supports the mass of all AGSE subsystems

This will be tested using solid works simulations and then experimentally verified using equivalent masses.

Rocker Bogie suspension

Pivots up and down to compensate for uneven terrain

Solid Works simulations will be used to test the design, as well as field testing on different kinds of terrain.

6 Wheel Servo Drive

Drives the AGSE

The exact torque required to move the mass of the AGSE will be calculated, and then experimentally verified on different terrains using equivalent masses.

Table 27 details the performance characteristics and evaluation and verification

metrics for the individual components of the camera subsystem.

Table 27: Camera Subsystem

Component Performance Characteristics

Evaluation and Verification Metrics

PixyCamCMU5 Takes color images

The camera will be tested alongside the data logging shield to ensure that the camera is successfully capturing and storing images for analysis.

Camera Pan-Tilt Head

Rotates the camera head using a servo

The pan and tilt will be tested prior to use in the AGSE for consistency in angular velocity

Arduino Uno Rev3

Controls other components and uses custom designed on board software for edge detection and color analysis

Laboratory testing using different shapes and colors to test for the accuracy and consistency of the software and algorithms.

Transcend 8gb Class 10 SD card

Stores Images for data analysis

This will be tested with the Pixy cam and data logging shield to determine if all components are operating correctly with one another. Debugging will be used as necessary to fix any errors in the code.

Adafruit Data Logging Shield

Interfaces between the SD card and the Arduino.

This will be tested with the Pixy cam and SD card to determine if all components are operating correctly with one another. Debugging will be used as necessary to fix any errors in the code.

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Table 28 details the performance characteristics and evaluation and verification

metrics for the individual components of the Laser Ranging subsystem

Table 28: Laser Ranging Subsystem

Component Performance Characteristics

Evaluation and Verification Metrics

Lightware SF02/F Laser Ranging Module

Emit and detect laser pulses to measure the time it takes for a laser reflect back to the module.

The laser will be tested to ensure that it successfully fires and detects the laser pulses.

Polulu 200 Step Bipolar Stepper motor

Rotate the motor in a 180 degree sweep in intervals of 0.45 degrees.

This will be tested for accuracy and consistency. Testing will also be used to determine the angular velocity of the stepper motor.

Pololu A4988 Stepper Motor Driver

Interface between the motor and the Arduino board and provide current limiting and over-current protection.

The stepper motor driver will undergo lab testing with the stepper motor to ensure that it can accurately rotate the laser module in consistent 0.45 degree steps.

Arduino Uno R3 Collect analyze and transmit data

Testing will be carried out to ensure proper collection, analysis and transmission of data from the LiDAR subsystem.

Table 29 details the performance characteristics and evaluation and verification

metrics for the individual components of the payload retrieval subsystem.

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Table 29: Payload Retrieval System

Component Performance Characteristics

Evaluation and Verification Metrics

HS-805BB Shoulder Servo

Rotate the shoulder of the arm centered from the base as necessary, with a range of 270 degrees.

All servos will be tested independently for functionality and together to ensure that the robotic arm is functional as a whole and can lift the payload.

HS-755HB Elbow Servos

Connect to the aluminum brackets to rotate the arms with a range of 180 degrees

All servos will be tested independently for functionality and together to ensure that the robotic arm is functional as a whole and can lift the payload.

HS-645MG Wrist Servo

Rotate the grip circularly with a range of 180 degrees

All servos will be tested independently for functionality and together to ensure that the robotic arm is functional as a whole and can lift the payload.

HS-422 Gripper

Securely close a grip around the payload for

The Gripper will be tested for its ability to grip and lift the payload using a payload prototype of the same dimensions and mass.

OV7670 Camera Module

Use edge detection to determine if the payload is off center from the HS-422 gripper.

The OV7670 will be lab tested for functionality and for the accuracy and consistency of the custom made edge detection algorithm it will be using. Debugging will be used as necessary to errors in the algorithm.

Arduino Uno R3

Collect and analyze data from the camera module using custom made software. Transmit data for any necessary adjustments in the arms

Testing will be carried out to ensure proper interfacing and data collection/transmission between the different components of the subsystem.

Table 30 details the performance characteristics and evaluation and verification

metrics for the individual components of the main computer subsystem.

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Table 30: Main Computer Component Performance

Characteristics Evaluation and Verification Metrics

Arduino Mega

Collect, analyze and transmit data between all subsystems

Testing will be carried out to ensure proper collection, analysis and transmission of data from and to all subsystems.

Master Switch Activate all AGSE subsystems

Pause Switch Temporarily disable all AGSE subsystems.

Table 31 details the performance characteristics and evaluation and verification

metrics for the individual components of the power supply subsystem.

Table 31: Power Supply

Component Performance Characteristic

Evaluation and Verification Metrics

AllPower 50,000 mAh Power Bank

Provide power to all AGSE microcontrollers

The power bank will be tested for its ability to supply power to the different AGSE microcontrollers using voltage and current measurements.

Component level Lithium polymer power banks

Provide power to specific subsystem components.

Each power banks will be tested individually with their corresponding component using voltage and current measurements.

Astro Pro Multi Voltage 20,000 mAh Power Banks

Provide power to the motors

The power banks will each be tested with the motors using voltage and current measurements.

Verification Plan and Status To address the requirements set in the NASA student launch handbook, the

requirements and verification plans have been set and detailed in table xx.

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Table 32: AGSE Verification

Payload Requirement

Design Feature Verification Plan Status

Teams will position their launch vehicle horizontally or vertically on the launch pad.

The AGSE’s robotic arm only has 20 inches of reach. Therefore the launch vehicle will be positioned horizontally.

N/A N/A

A master switch will be activated to power on all autonomous procedures and subroutines.

The Master Switch on the main computer will be used to power on all AGSE autonomous procedures.

Test the functionality of the switch using a prototype. Testing will also be done on the final AGSE system.

In progress. Materials are being acquired for construction.

A pause switch will be activated temporarily halting all AGSE procedures and subroutines.

The pause switch on the main computer will be used to pause all AGSE procedures and subroutines.

Test the functionality of the switch using a prototype. Testing will also be done on the final AGSE system.

In progress.

During autonomous procedures, the team is not permitted to interact with their AGSE

The main computer will be programed to manage all AGSE subsystems without human intervention.

The main computer and all AGSE subsystems will be tested using prototyping as well as field testing to ensure that the system can complete it

In progress.

Once the pause switch is deactivated, the AGSE will capture and contain the payload within the launch vehicle.

The payload retrieval system will physically lift and carry the payload for transport. It will also place the payload into the payload bay of the launch rocket.

The payload retrieval system will be tested for its ability to locate and lift the arm using a prototype payload of equivalent dimensions and mass

In progress.

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After the erection of the launch vehicle, a team member will arm recovery electronics.

The recovery electronics will be activated by a pressure sensor once the payload is inside the payload bay.

The recovery electronics will be tested using a prototype and test launches on the subscale rocket.

In progress. Materials are being acquired for construction.

The igniter is manually installed and the area is evacuated.

N/A N/A N/A

Once the launch services official has inspected the launch vehicle and declares that the system is eligible for launch, he/she will activate a master arming switch to enable ignition procedures

N/A N/A N/A

The Launch Control Officer (LCO) will activate a hard switch, and then provide a 5-second countdown.

N/A N/A N/A

At the end of the countdown, the LCO will push the final launch button, initiating launch

N/A N/A N/A

. The rocket will launch as designed and jettison the payload at 1,000 feet AGL during descent.

An altimeter will be used to initiate payload jettison at 1,000 feet AGL

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Sensors that rely on Earth’s magnetic field are prohibited.

The AGSE will not use any sensors which are dependent on Earth’s magnetic field.

N/A N/A

Ultrasonic or other sound-based sensors are prohibited.

The AGSE will use an LRS system (Electromagnetic waves) for ranging rather than Ultrasonic Sensors.

The LRS subsystem will be tested for functionality, precision and consistency by prototyping and field testing.

In progress.

Earth-based or Earth orbit-based radio aids (e.g. GPS, VOR, cell phone) are prohibited.

The AGSE will be navigated by the camera subsystem rather than Earth orbit based radio aids.

The camera subsystem will be tested for functionality, precision and consistency by prototyping and field testing.

In progress.

Open circuit pneumatics are prohibited.

The AGSE will only use servos and motors for mechanical work. rather than any pneumatic systems.

The servos will be tested for functionality by prototyping and field testing to ensure they can provide the necessary power in various environments and terrains.

In progress.

Air breathing systems are prohibited.

The AGSE will only use servos and motors for mechanical work, and the wheels will be machined out of aluminum without pressurized tires. Hence no Air breathing systems will be needed.

The servos will be tested for functionality by prototyping and field testing to ensure they can provide the necessary power in various environments and terrains.

In progress.

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Preliminary Integration Plan

The AGSE integration plan consists of three phases: tactical, operational and

analytical. Tactical refers to the manufacturing and construction of the AGSE.

Operational refers to the time during which the AGSE operates autonomously.

Analytical refers the any data analysis that takes place after the AGSE had competed

its designated tasks.

Tactical (construction and manufacturing)

Body

All subsystems of the AGSE must be contained within or attached to the body I some

manner. It is therefore important that the body is structurally solid, but also

lightweight as to allow for adequately efficient energy efficiency. For those reasons,

the team has selected to construct the body and its components from varying

thicknesses of 6061-T6 aluminum sheet metal. Aluminum was also selected for cost

effectiveness and weldability. The sheet metal will be cut to size using a saw, end

kill, or plasma cutter. Each separate aluminum piece will then be welded or bolted

together as necessary to form the chassis of the AGSE.

The AGSE suspension will be cut to size from the same aluminum sheet metal as the

chassis using a saw, end mill, or plasma cutter and welded or bolted together as

necessary. Holes will be drilled and bearings attached to aluminum rods will be

inserted to provide the pivot points. The upper bogie is attached to the chassis with

bolts and thus unable to pivot freely. The bottom two bogie arms will have

rotational freedom, which allows the wheels to remain on the ground when the

AGSE encounters uneven terrain. Wheels and wheel-related equipment will be

attached to the end of each of the bogie arms. Wheels will be fabricated from

aluminum using a CNC machine, and tread will be fabricated and attached

afterwards using sheet metal screws.

Camera Subsystem

The camera subsystem will be directly mounted onto an aluminum mast. The mast

itself will be formed from hollow aluminum round tubing and bolted to the chassis.

The pan-tilt head will be mounted directly onto the top of the mast on a fabricated

mounting plate that attaches directly to the top of the mast. The pan-tilt head will

support the pixy camera itself, which will connect to the Arduino mounted onto the

chassis through wires running inside of the aluminum tubing that comprises the

mast.

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LSR

The LSR components will be directly mounted to the front of the chassis. These

components will view through an opening in the chassis, which allows for room

(inside the chassis) to place the required pan and tilt hardware to allow the laser to

target appropriately.

Payload Retrieval

The robotic arm comes as a kit and only needs to be assembled according to the

manufacturer's instructions. Once constructed, the arm will be mounted onto the

front top section of the AGSE chassis (see fig xx)..

Main Computer /Power Supply

The main computer and power supply subsystem components will be mounted onto

the chassis in their respective locations (yet to be determined).

Operational All AGSE subsystems will be connected to a master switch which will be connected

to the master microcontroller. The master microcontroller will also be connected to

a pause switch which will function to halt all autonomous subroutines when/if

toggled

Body

Through the operational phase the suspension will be used to compensate for

unevenness in the terrain. The two rocker arms on the side of the chassis will pivot

independent of one another allowing for a height difference on both sides without

an effect on the tilt of the chassis. This will prevent the chassis from tipping over and

allow for the rover to travel on uneven terrain as would be expected in a Martian

environment

Camera Subsystem

Throughout the operational phase the camera subsystem will repeat its subroutines

to recalibrate the orientation of the AGSE with respect to the payload to ensure the

proper trajectory.

LSR

Through the operation phase, the LSR will repeat its subroutines to recalibrate the

distance from the front face of the AGSE to the payload to maintain the proper

trajectory.

Payload Retrieval

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The payload retrieval subsystem will activate when the AGSE is within a 20 inch

reach of the robotic arm. The data required will be retrieved from the LSR system

via the master microcontroller.

Analytical A qualitative analysis of the entire AGSE system and its subsystems will be carried

out to determine the success of the mission. No quantitative data will be available

for analysis.

Precision of Instrumentation and Repeatability The only subsystems utilization any instrumentation equipment are the Camera

Subsystem, the Laser Rangefinder Subsystem (LRS), and the Payload Retrieval

Subsystem. Each subsystem and its measurement instrumentation is addressed in

the following tables.

Table 33: Camera Subsystem Instrumentation Performance Instrumentation Accuracy Repeatability Recovery Camera Pan-Tilt Head

Experimentally determined

Can be repeated with every launch

N/A

Table 34: LRS subsystem Instrumentation Performance Instrumentation Accuracy Repeatability Recovery Lightware SF02/F Laser Ranging Module

±(0.1 + 1%) m Can be repeated with every launch

N/A

Polulu Bipolar Stepper

±0.1 deg Can be repeated with every launch

N/A

Table 35: Payload Retrieval subsystem Instrumentation Performance Instrumentation Accuracy Repeatability Recovery Servos ±0.1 deg Can be repeated

with every launch

N/A

Camera Module Edge Detection

Experimentally determined

Can be repeated with every launch

N/A

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Key Components of AGSE The key components of the AGSE and how they will work together to accomplish the overall mission objective are all listed in Table 36.

Table 36: Internal Interfaces of AGSE Components How they Interface

Arduino Mega / Robotic Arm Arduino Mega will interface with the Robotic Arm and provide the arm with all of its controls.

Arduino Mega / LIDAR Arduino Uno

The Arduino Mega will preside as the master controller over the LIDAR’s slave Arduino Uno, which will control the LIDAR system. This Arduino Uno will receive commands from the Arduino Mega, and will process data from the LIDAR system and send that data back to the Arduino Mega for further commands.

Arduino Mega / Pixy Arduino Uno

The Arduino Mega will preside as the master controller over the edge and color detection system’s Arduino Uno. This Arduino Uno will receive commands from the Arduino Mega, and will process data from the edge and color detection system and send that data back to the Arduino Mega for further commands.

LIDAR Arduino Uno / LIDAR system

The LIDAR Arduino Uno will control the LIDAR system, processing any data received and relaying that data back to the master controller.

Pixy Camera / Arduino Uno

The Pixy Camera will perform the color and edge detection for the AGSE, and will process any data received from the Pixy Camera and relay the data back to the master controller.

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Arduino Mega / Motor Controllers

The Arduino Mega will interface with the motor controllers and issue commands to the controllers pertaining to the movement of the AGSE. These commands will be determined from the data received from the slaved Arduino Uno units for the Pixy Camera and the LIDAR system. All four motor controllers will work together simultaneously to provide the steering and movement capabilities of the AGSE.

Front Left Motor Driver / Motors

The Front Left Motor Driver will connect to the front left drive motor and steering motor / servo. This motor driver will power the motors and be responsible for the front left wheel’s steering and navigation.

Front Right Motor Driver / Motors

The Front Right Motor Driver will connect to the front right drive motor and steering motor / servo. This motor driver will power the motors and be responsible for the front right wheel’s steering and navigation.

Back Left Motor Driver / Motors

The Back Left Motor Driver will connect to the back left drive motor and steering motor / servo. This motor driver will power the motors and be responsible for the back left wheel’s steering and navigation.

Back Right Motor Driver / Motors

The Back Right Motor Driver will connect to the back right drive motor and steering motor / servo. This motor driver will power the motors and be responsible for the back right wheel’s steering and navigation.

Main Power Supply / Arduino Mega

The Main Power Supply will interface with the Arduino Mega and provide power to the Arduino Mega and all of its slaved Arduino Uno microcontrollers, as well as provide power to the Robotic Arm subsystem.

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Power Supply / Motor Controllers

Each motor controller will have a dedicated power supply allocated to it, which will provide power to both the controller and the attached steering and drive motors.

Suspension / Motors

The motors will be mounted into the rocker-bogie suspension system. The drive motors will be connected to both the suspension system and the wheels, while the steering servos will be attached to a rotating portion of the wheel frame that will allow the AGSE to turn on a dime.

Motors / Wheels

The drive motors will be mounted into the framework of the suspension and attached to the wheels at the center of each wheel. This will provide the forward / backward motion of the wheels.

Robotic Arm / AGSE Chassis

The Robotic Arm will be bolted directly to the chassis top, and allowed to move freely via the servo motors that will control the arm.

Chassis / LIDAR System The LIDAR system will be mounted to the front right face of the Chassis.

Chassis / Pixy Camera The Pixy Camera will be mounted onto a camera mast, which will be bolted to the chassis’ top.

Suspension / Chassis

The rocker-bogie suspension system will be attached to the chassis via a single rod, and the back half of the rocker bogie suspension will be permanently attached to the chassis. The front bogie will be attached to the rear bogie (and not attached to the chassis), which will allow the suspension system to function as intended.

AGSE Concept Features and Definition The design of the AGSE chassis and framework is completely designed by the team from scratch. The only idea that will be taken from previous designs already in use is the rocker-bogie style of suspension system. However, the method by which the

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rocker-bogie suspension system will work will be designed from scratch by the team. The overall research goal of the competition is to research and develop innovative methods by which a payload may be recovered and delivered from another planet and back to Earth for analysis. This rover design is aimed to satisfy many of the concerns involved with designing such a device, including autonomously locating a payload, navigating to the payload, and delivering that payload back to the rocket for its trip back to Earth (or another location). The designed AGSE has a sufficient level of challenge for a number of reasons. First, the programming required to make the entire AGSE function is vast, and provides a large amount of the challenge associated with the entire project. Second, the manufacturing of the AGSE contributes to the overall challenge, as members of the team much learn to machine and weld to the proper standards as would be required for the AGSE function properly. Science Value The objective of the project is to research innovative methods by which an object might be recovered and loaded into a rocket autonomously. This research will prove useful when planning an interplanetary mission that requires a robotic device that will retrieve a payload (or set of payloads) and send them back to Earth for analysis. As such, the AGSE will be designed to accomplish a set of several scientific objectives that will generate data relevant to such a mission. These objectives and relevant success criteria are listed in Table 37.

Table 37: AGSE Objectives and Success Criteria Objectives Success Criteria

Construct Autonomous Ground Support Equipment (AGSE) that can navigate autonomously to a payload and to the rocket.

The AGSE navigates to the payload and rocket in such a way that allows for a successful retrieval of the payload and insertion of the payload into the rocket payload bay.

Program a purchased robotic arm to locate and acquire the payload and consequentially insert that payload into the rocket payload bay.

The robotic arm successfully retrieves the payload and inserts it into the rocket through the payload bay doors.

Design and build a payload bay that autonomously seals and houses the payload during all stages of flight (ascent, descent, landing, etc).

The payload doors seal autonomously after the payload is inserted, and the payload remains in the rocket safely, without damage, during flight and is found in such a way when the payload containment bay is retrieved by team members or other personnel.

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Deploy the payload containment bay at approximately 1000 feet AGL.

The payload containment bay is successfully deployed within 50 feet of 1000 feet AGL without damage to the rocket or the payload containment bay.

V. Project Plan Timelines

Figure 27: NASA Student Launch Timeline

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Figure 28: AGSE and Rocket Construction Timeline

Figure 29: Outreach Timeline

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Budget

Table 38: Mini-MAV Budget

Item Quantity Amount Total Shipping

6" rocket 6" AF bulkhead 2 $8.95 $17.90 $0.00 6" to 54mm centering ring 3 $9.25 $27.75 $0.00 6" x 12" avionics bay 1 $69.95 $69.95 $0.00 6" x 48" Blue Tube 3 $66.95 $200.85 $0.00 6" fiberglass nosecone (Model: FNC-6.0)

1 $99.95 $99.95 $13.95

Nylon Shock Cord: 5/8", 5 yards, presewn endloops

2 $19.00 $38.00 $0.00

54mm x 48" MMT Airframe Blue Tube

1 $23.95 $23.95 $0.00

CNC fin slots (service fee) 6 $4.00 $24.00 $0.00 96" elliptical parachute 1 $275.00 $275.00 $0.00

Total: 20 $777.35 $13.95 Motor hardware/reloads Aerotech K1100T-L reload Kit 1 $118.64 $118.64 $0.00 54/1706 Motor Hardware Set (w/ Forward Seal Disc)

1 $196.88 $196.88 $0.00

Total: 2 $315.52 $0.00 Fins 2' x 4' 1/4" Finnish birch aircraft plywood for fins

1 $56.38 $56.38 $9.14

Total: 1 $56.38 $9.14 Total Rocket: $1,576.24 4" rocket (subscale) 4" x 48" Blue Tube 2 $38.95 $77.90 4" x 8" avionics bay 1 $41.95 $41.95 38mm x 48" MMT Airframe Blue Tube

1 $16.49 $16.49

3.9" bulkhead w/ eyebolt 2 $4.29 $8.58 3.9" to 38mm centering ring 3 $4.25 $12.75 $24.95 3.9" plastic nosecone 1 $21.95 $21.95

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shock cord, 3 yd, 1/2" nylon tubular, presewn endloops

2 $14.00 $28.00

30" elliptical parachute 2 $64.00 $128.00 18" elliptical drogue 3 $51.00 $153.00 $17.00 Aerotech I161W-M reload 1 $37.79 $37.79 38/360 Motor Hardware Set 1 $114.61 $114.61 $60.83 Total: 19 $641.02 $102.78 Avionics Missile Works RRC3 Sport Altimeter

2 $69.95 $139.90 $6.10

Altus Metrum TeleGPS 1 $214.00 $214.00 $8.37 Nuts / Bolts / Hardware 1 $50.00 Total: 4 $403.90 $14.47 Robotic Arm AL5D Robotic Arm Combo Kit (BotBoarduino)

1 $309.81 $309.81 $0.00

Beaglebone Black Microcontroller

1 $79.95 $79.95 $0.00

Aluminum Multi-Purpose Servo Bracket

1 $11.95 $11.95 $0.00

Aluminum "C" Servo Bracket 1 $12.90 $12.90 $0.00

HS-805BB (343 oz. in.) Mega Servo

1 $39.99 $39.99 $0.00

Heavy-Duty Wrist rotator 1 $45.84 $45.84 $0.00 Total: 6 $500.44 $0.00 Payload Compartment Arduino Uno 1 $24.95 $24.95 $0.00 EM506 GPS 1 $39.95 $39.95 $0.00 GPS Shield 1 $14.95 $14.95 $0.00 Xbee Pro 900 2 $54.95 $109.90 $0.00 Antenna 2 $7.95 $15.90 $0.00 Power Source 1 $19.99 $19.99 $0.00 Servo Motor 1 $13.95 $13.95 $0.00 Rack/ Pinion Set 1 $7.99 $7.99 $0.00

Spring-Loaded Hinge 2 $3.38 $6.76 $0.00

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Total: 12 $254.34 $0.00 Navigation Package Arduino Uno 2 $24.95 $49.90 $9.47

Pixy Camera Module 1 $69.00 $69.00 $0.00

Pan-Tilt Head 1 $39.00 $39.00 $0.00

FabScan 3D Scanner Shield 1 $12.84 $12.84 $40.33

Pololu A4988 Stepper Motor Driver Carrier

1 $12.22 $12.22 $0.00

200 Step, Unipolar/Bipolar Stepper Motor

1 $19.92 $19.92 $0.00

5mW Laser Module emmiter (Red Line)

1 $7.57 $7.57 $0.00

Adafruit Data logging shield 1 $19.95 $19.95 $9.47

Tanscend 8gb SD card 1 $11.95 $11.95 $0.00

Total: 10 $242.35 $59.27 AGSE Structural Components Pack of 30 ball bearings 1 $17.82 $17.82 $0.00

2" x 3" x 1/8", 8-ft long aluminum rectangular Tubing (6061-T6)

1 $52.26 $52.26 $52.80

Alluminum Sheet Metal 6061-T6 (36" x 48" sheet)

2 $133.35 $266.70 $0.00

Nuts, bolts, and washers 1 $50.00 $0.00 1" Inner Diameter 6061-T6 Alluminum Rod

1 $22.42 $22.42

12V 200 RPM DC Motor 10 $11.95 $119.50 $20.00

Helical Couplers 6 $20.62 $123.72 $15.00 Anker Astro Pro 20Ah Lithium Battery Pack

3 $99.99 $299.97 $0.00

Total: 25 $952.39 $87.80 AGSE Controller Components Arduino Mega 1 $49.95 $49.95 $0.00

T-Rex Motor Driver 5 $74.95 $374.75 $0.00

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Total: 6 $424.70 $0.00 Launch Competition Airfare (6 x 400) $2,400.00 Hotel (5 x 2 x 90) $900.00 Rental Van (1 x 200) $400.00 Food and Entertainment $500.00 Freight $400.00 Total Launch Competition: $4,600.00 Outreach $3,500.00 Launch Pad Total: $4,314.05 Project Total: $12,955.80

Figure 30: Budget Distribution

$1,576.24

$743.80

$500.44

$254.34

$301.62

$1,040.19

$424.70

$4,600.00

$3,500.00

Budget Distribution

Rocket

Subscale Rocket

Robotic Arm

Payload Compartment

Navigation Package

AGSE Structural Components

AGSE Controller Components

Competition

Outreach

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Funding Plan

Table 39: Funding Plan

Activity Funded by Amount Funds used for

Junior Rocket Owls Program

Citrus College Foundation / Private Donors

$8,000.00

$6, 000.00 to sponsor the Citrus Rocket Owls’ participation in the NASA Student launch and $2,000.00 to purchase supplies for the Junior Rocket Owls activities

Azusa STEM Pathways

Canyon City Foundation

$6,500.00

$5,000.00 to sponsor the Citrus Rocket Owls’ participation in the NASA Student launch and $1,500.00 to purchase supplies for the STEM Pathways activities

Science and Technology Fundraiser Event

Citrus College in collaboration with local businesses

$2,000.00 Sponsor the Citrus Rocket Owls’ participation in the NASA Student launch

Night on the Plaza Fundraising Event

Glendora Public Library Foundation

$200.00 Sponsor the Citrus Rocket Owls’ participation in the NASA Student launch

Presentation to the RACE to STEM committee members

RACE to STEM Title V Grant

$500.00 Sponsor the Citrus Rocket Owls’ participation in the NASA Student launch

KIWANIS Club Presentation

KIWANIS Club $500.00 Sponsor the Citrus Rocket Owls’ participation in the NASA Student launch

Solicitations to local businesses

Private donations $2,000.00 Sponsor the Citrus Rocket Owls’ participation in the NASA Student launch

Total:

$19,700.00

Sponsor the Citrus Rocket Owls’ participation in the NASA Student launch and their educational engagement activities

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Plans to Solicit Additional Support The team received a license to use the simulation program STAR-CCM+ for 5,000 hours to help with simulations of the rocket's aerodynamic and structural requirements. The team has access to Pasadena City College's Engineering Lab, and to the expertise of the USC Rocket Propulsion Laboratory (RPL) team. Sustainability Plan

In order for the rocketry project to be sustained at Citrus College and in the community, the team plans to do the following:

Maintain collaboration with the Rocket Team at Cal Poly Pomona. This collaboration will extend the existing relationship between the Citrus College Physics Department and the Aerospace Department at Cal Poly that started two years ago as a result of a joint High Altitude Balloon Project between the two colleges, funded by NASA through the California Grant Space Consortium.

Submit abstracts of team research for presentations to the annual HTCC Student Research Conference for California Community Colleges hosted by the University of California at Irvine. This conference is an opportunity for the Rocket Owls to highlight their research, gain experience in a research conference setting and potentially receive recognition, scholarships and have their research published.

Maintain our relationship with the Glendora Public Library by conducting rocketry workshops open to the public every year.

As part of the newly created Physics lab curriculum with a peer-led component, the Rocket Owls will facilitate a rocketry lab every semester for the PHYS 201 (General Physics–Mechanics) students, under the supervision of the instructor in charge, Lucia Riderer, who is also the faculty advisor of the Rocket Owls.

Present team research, methods and results at the Citrus College Physics Festival to approximately 200 Physics students every December and June.

Present team research, methods and results at the Natural and Physical Science Division faculty and staff meeting at least twice every year.

Prepare and present a poster at the yearly Citrus College STEM Symposium, open to all Citrus College students.

Recruit new members for next year’s competition and mentor the new team for at least one month. The new Rocket Owls will be recruited by the team at the end of a rocketry workshop open to all interested students at the beginning of June, 2015.

Maintain existing Facebook page (https://www.facebook.com/CitrusCollegeRocketOwls) and create other social media web pages to constantly update their progress on the NASA SLP.

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VI. Conclusion Project scension’s mission is to retrieve a 4 oz. cylindrical payload from the ground, launch it to an altitude of 3000 ft AGL, and eject the payload at 1000 ft AGL to be recovered separately from the rest of the launch vehicle. The payload will be identified and retrieved autonomously by a six-wheeled rover using a camera and laser navigation system and a robotic arm. The rover will navigate autonomously to the launch vehicle and insert the payload through spring-loaded doors into the payload bay of the vehicle. Team personnel will manually move the launch vehicle to a vertical launch position, install the igniter, and clear the area for launch. The 20 lb, 6” diameter, 112” long launch vehicle will be powered by an AeroTech K1100T motor to an altitude of 3000 ft AGL. Upon descent, the payload bay will be ejected at 1000 ft AGL and descend under its own parachute. GPS tracking units will facilitate recovery of the launch vehicle and payload.

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Appendix A: Citrus College Profile Since 1967, Citrus College has been providing a quality educational experience for the communities of Azusa, Glendora, Duarte, Claremont and Monrovia. It is currently home to over 12,000 students, the majority of whom are considered ethnic minorities, and is dedicated to creating a diverse and welcoming learning environment that supports educational achievement for all its students. Citrus College offers many programs that promote community awareness in numerous STEM (Science, Technology, Engineering, and Mathematics) related fields. Biological and Physical Sciences is the second most common major in the school. There are also numerous extracurricular programs aimed at increasing interest in STEM subjects within the community, such as the SIGMA (Support and Inspire to Gain Motivation and Achievement) peer mentor program; the PAGE (Pre-Algebra, Algebra, Geometry Enrichment) summer K-12 mathematics enrichment program; and the Secrets of Science Summer Camp that provides K-12 students with practical experience in biology, chemistry, astronomy and physics laboratories. Students at Citrus College are active participants in many STEM-related activities. In past years, students have participated in NASA’s Reduced Gravity Education Flight Program (RGEFP), have launched a near-space sounding balloon, and have also traveled to Huntsville, Alabama and to Salt Lake City, Utah as participants in the 2013 and 2014 USLI SLP (University Student Launch Initiative Student Launch Projects). We hope to duplicate, if not supersede, the previous years’ successes with our own 2015 Student Launch project.

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Appendix B: Safety Contract

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Appendix C: NAR High Power Safety Code

1. Certification. I will only fly high power rockets or possess high power rocket motors that are within the scope of my user certification and required licensing.

2. Materials. I will use only lightweight materials such as paper, wood, rubber, plastic, fiberglass, or when necessary ductile metal, for the construction of my rocket.

3. Motors. I will use only certified, commercially made rocket motors, and will not tamper with these motors or use them for any purposes except those recommended by the manufacturer. I will not allow smoking, open flames, nor heat sources within 25 feet of these motors.

4. Ignition System. I will launch my rockets with an electrical launch system, and with electrical motor igniters that are installed in the motor only after my rocket is at the launch pad or in a designated prepping area. My launch system will have a safety interlock that is in series with the launch switch that is not installed until my rocket is ready for launch, and will use a launch switch that returns to the “off” position when released. The function of onboard energetics and firing circuits will be inhibited except when my rocket is in the launching position.

5. Misfires. If my rocket does not launch when I press the button of my electrical launch system, I will remove the launcher’s safety interlock or disconnect its battery, and will wait 60 seconds after the last launch attempt before allowing anyone to approach the rocket.

6. Launch Safety. I will use a 5-second countdown before launch. I will ensure that a means is available to warn participants and spectators in the event of a problem. I will ensure that no person is closer to the launch pad than allowed by the accompanying Minimum Distance Table. When arming onboard energetics and firing circuits I will ensure that no person is at the pad except safety personnel and those required for arming and disarming operations. I will check the stability of my rocket before flight and will not fly it if it cannot be determined to be stable. When conducting a simultaneous launch of more than one high power rocket I will observe the additional requirements of NFPA 1127.

7. Launcher. I will launch my rocket from a stable device that provides rigid guidance until the rocket has attained a speed that ensures a stable flight, and that is pointed to within 20 degrees of vertical. If the wind speed exceeds 5 miles per hour I will use a launcher length that permits the rocket to attain a safe velocity before separation from the launcher. I will use a blast deflector to prevent the motor’s exhaust from hitting the ground. I will ensure that dry grass is cleared around each launch pad in accordance with the accompanying Minimum Distance table, and will increase this distance by a factor of 1.5 and clear that area of all combustible material if the rocket motor being launched uses titanium sponge in the propellant.

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8. Size. My rocket will not contain any combination of motors that total more than 40,960 N-sec (9208 pound-seconds) of total impulse. My rocket will not weigh more at liftoff than one-third of the certified average thrust of the high power rocket motor(s) intended to be ignited at launch.

9. Flight Safety. I will not launch my rocket at targets, into clouds, near airplanes, nor on trajectories that take it directly over the heads of spectators or beyond the boundaries of the launch site, and will not put any flammable or explosive payload in my rocket. I will not launch my rockets if wind speeds exceed 20 miles per hour. I will comply with Federal Aviation Administration airspace regulations when flying, and will ensure that my rocket will not exceed any applicable altitude limit in effect at that launch site.

10. Launch Site. I will launch my rocket outdoors, in an open area where trees, power lines, occupied buildings, and persons not involved in the launch do not present a hazard, and that is at least as large on its smallest dimension as one-half of the maximum altitude to which rockets are allowed to be flown at that site or 1500 feet, whichever is greater, or 1000 feet for rockets with a combined total impulse of less than 160 N-sec, a total liftoff weight of less than 1500 grams, and a maximum expected altitude of less than 610 meters (2000 feet).

11. Launcher Location. My launcher will be 1500 feet from any occupied building or from any public highway on which traffic flow exceeds 10 vehicles per hour, not including traffic flow related to the launch. It will also be no closer than the appropriate Minimum Personnel Distance from the accompanying table from any boundary of the launch site.

12. Recovery System. I will use a recovery system such as a parachute in my rocket so that all parts of my rocket return safely and undamaged and can be flown again, and I will use only flame-resistant or fireproof recovery system wadding in my rocket.

13. Recovery Safety. I will not attempt to recover my rocket from power lines, tall trees, or other dangerous places, fly it under conditions where it is likely to recover in spectator areas or outside the launch site, nor attempt to catch it as it approaches the ground.

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MINIMUM DISTANCE TABLE

Installed Total Impulse

(Newton-Seconds)

Equivalent High Power Motor Type

Minimum Diameter of

Cleared Area (ft.)

Minimum Personnel

Distance (ft.)

Minimum Personnel Distance (Complex

Rocket) (ft.)

0 — 320.00 H or smaller 50 100 200

320.01 — 640.00 I 50 100 200

640.01 — 1,280.00

J 50 100 200

1,280.01 — 2,560.00

K 75 200 300

2,560.01 — 5,120.00

L 100 300 500

5,120.01 — 10,240.00

M 125 500 1000

10,240.01 — 20,480.00

N 125 1000 1500

20,480.01 — 40,960.00

O 125 1500 2000

Note: A Complex rocket is one that is multi-staged or that is propelled by two or more rocket motors

This document is effective as of August 2012 (http://www.nar.org/safety-information/high-power-rocket-safety-code/)