fiu_nsl_pdr

NASA University Student Launch Initiative

2015-2016 Preliminary Design Report for MAV

FIU PantherWorks Space Team

November 6, 2015

PantherWorks Space Team | 2015 – 2016 NASA USLI PDR | 2

TABLE OF CONTENTS List of Tables .................................................................................................................................. 5

List of Figures ................................................................................................................................. 7

Section 1: Summary of PDR Report ............................................................................................. 12

I. Team Summary .................................................................................................................. 12

II. Launch Vehicle Summary .............................................................................................. 12

III. AGSE/Payload Summary ............................................................................................... 12

Section 2: Changes Made Since Proposal ..................................................................................... 13

I. Vehicle Criteria Changes ................................................................................................... 13

II. AGSE Criteria Changes.................................................................................................. 14

III. Project Plan Changes ...................................................................................................... 14

Section 3: Vehicle Criteria ............................................................................................................ 15

I. Launch Vehicle Design ...................................................................................................... 15

Mission Statement ................................................................................................................. 15

Mission Verification ............................................................................................................. 15

Mission Success Criteria ....................................................................................................... 16

Design of Launch Vehicle..................................................................................................... 16

Designs At a Systems Level.................................................................................................. 21

Launch Vehicle Verification ................................................................................................. 44

Critical Mass Statement ........................................................................................................ 48

Full Launch Vehicle Assembly ............................................................................................. 50

II. Recovery Subsystem ...................................................................................................... 50

Parachute Selection Rationale............................................................................................... 50

Recovery Flight Path............................................................................................................. 51

Recovery System Components ............................................................................................. 52

Electrical Schematic.............................................................................................................. 56

Recovery System Verification .............................................................................................. 56

III. AirBrake Subsystem ....................................................................................................... 59


Subsystem Characteristics..................................................................................................... 60

System Components.............................................................................................................. 62

Aerodynamic Analysis .......................................................................................................... 73

Code Development................................................................................................................ 81

IV. Mission Performance Predictions................................................................................... 88

Mission Performance Criteria ............................................................................................... 88

Launch Vehicle Characteristics Summary ............................................................................ 88

Motor Selection ..................................................................................................................... 89

Stability Analysis .................................................................................................................. 92

Flight Simulations ................................................................................................................. 93

Landing Analysis .................................................................................................................. 99

V. Interfaces and Integration ............................................................................................. 100

Internal Vehicle Interfaces .................................................................................................. 100

External Vehicle Intergration .............................................................................................. 103

VI. Safety ............................................................................................................................ 104

Safety Officer ...................................................................................................................... 104

Preliminary Checklist.......................................................................................................... 105

Safety Procedures................................................................................................................ 106

Hazard Analysis .................................................................................................................. 107

Design Failure Modes ......................................................................................................... 115

Section 4: AGSE Criteria ............................................................................................................ 116

Ground Support Performance Criteria ................................................................................ 116

Overall AGSE Sequence of Events..................................................................................... 117

Launch Rail Design............................................................................................................. 118

Rover Design....................................................................................................................... 136

AGSE Verification .............................................................................................................. 142

Section 5: Project Plan ................................................................................................................ 146

Budget Plan ......................................................................................................................... 146


Funding Plan ....................................................................................................................... 148

Timeline .............................................................................................................................. 149

Educational Engagement..................................................................................................... 150

Section 6: Appendices................................................................................................................. 151

Appendix A – Stability Analysis Graphs ................................................................................ 151

5 – MPH Wind .................................................................................................................... 151

10 – MPH Wind .................................................................................................................. 153

15 – MPH Wind .................................................................................................................. 156

20 – MPH Wind .................................................................................................................. 158

Appendix B – Timelines ......................................................................................................... 161

General Timeline................................................................................................................. 161

Detailed PDR/CDR Timeline.............................................................................................. 163


LIST OF TABLES Table 1: Launch Vehicle Overview .............................................................................................. 12

Table 2: Overall Ground Support Dimensions ............................................................................. 12

Table 3: Basic Mission Verification Metrics ................................................................................ 16

Table 4: Carbon Fiber Material Properties ................................................................................... 25

Table 5: Phenolic Tubing Material Properties .............................................................................. 25

Table 6: Payload Components ...................................................................................................... 29

Table 7: Telemega Specifications ................................................................................................. 38

Table 8: Perfect Flight Specifications ........................................................................................... 39

Table 9: Propulsion component descriptions ................................................................................ 41

Table 10: G5000 Specifications .................................................................................................... 44

Table 11: Vehicle Statement of Work Verification ...................................................................... 47

Table 12: Overall Assembly Mass Properties ............................................................................... 49

Table 13: Individual Subsystem Masses ....................................................................................... 50

Table 14: Parachute Selection Chart ............................................................................................. 50

Table 15: Recovery Events and Descriptions ............................................................................... 52

Table 16: Recovery System Verification ...................................................................................... 58

Table 17: AirBrake Components and Function ............................................................................ 60

Table 18: Screw Critical Speed calculations................................................................................. 63

Table 19: Power Calculations ....................................................................................................... 64

Table 20: Peak linear velocity....................................................................................................... 64

Table 21: Pitch Calculation ........................................................................................................... 64

Table 22: Torque from inputted force ........................................................................................... 65

Table 23: Force from inputted torque ........................................................................................... 65

Table 24: Stepper Motor Specifications........................................................................................ 67

Table 25: Proposed First Iteration Drag Force Calculation .......................................................... 74

Table 26: CFD Drag (Z) results with AirBrake closed ................................................................. 77

Table 27: CFD Drag (Z) results with AirBrake at five degrees .................................................... 78

Table 28: CFD Drag (Z) results with AirBrake at 15 degrees ...................................................... 78

Table 29: CFD Drag (Z) results at full deployment ...................................................................... 80

Table 30: Spreadsheet calculation at full deployment .................................................................. 81

Table 31: AirBrake Computer Inputs............................................................................................ 88

Table 32: Overall Rocksim Vehicle Values.................................................................................. 89

Table 33: Motor Simulated Specifications.................................................................................... 91

Table 34: Flight Simulation Data .................................................................................................. 97


Table 35: Descent Rate Energy Calculation ............................................................................... 100

Table 36: First Draft, Procedure Checklist ................................................................................. 106

Table 37: Risk Assessment Matrix ............................................................................................. 108

Table 38: Lab/Machine Shop Risk Assessment.......................................................................... 109

Table 39: AGSE Risk Assessment Matrix .................................................................................. 111

Table 40: Stability/ Propulsion Risk Assessment ....................................................................... 113

Table 41: Launch Day Risk Assessment..................................................................................... 113

Table 42: Environmental Effects Risk Assessment .................................................................... 115

Table 43: Failure Mode Analysis ................................................................................................ 116

Table 44: AGSE Sequence.......................................................................................................... 117

Table 45: Overall Launch Rail Dimension ................................................................................. 119

Table 46: PA-04 Specifications .................................................................................................. 130

Table 47: Ground Support Component Description ................................................................... 135

Table 48: AGSE Verification...................................................................................................... 145

Table 49: Comprehensive Launch Vehicle Budget .................................................................... 147

Table 50: Comprehensive AGSE Budget ................................................................................... 148


LIST OF FIGURES Figure 1: Overall Launch Vehicle Assembly................................................................................ 16

Figure 2: Overall Vehicle Configuration ...................................................................................... 17

Figure 3: Rocksim Simulation of Launch Vehicle........................................................................ 20

Figure 4: Von Karman nosecone, current model .......................................................................... 21

Figure 5: Von Kármán nosecone, proposal CAD model .............................................................. 22

Figure 6: Rational for Von Karman nosecone .............................................................................. 22

Figure 7: Base Nosecone Dimensions .......................................................................................... 23

Figure 8: Back View of Nosecone ................................................................................................ 23

Figure 9: Nosecone Mass Properties............................................................................................. 24

Figure 10: Launch Vehicle Airframe assembly ............................................................................ 24

Figure 11: Parachute Bay Airframe .............................................................................................. 26

Figure 12: Section View of Parachute Bay ................................................................................... 26

Figure 13: Mass Properties of parachute bay ................................................................................ 27

Figure 14: Payload Bay Subsystem .............................................................................................. 27

Figure 15: Isometric View with Airframe removed...................................................................... 28

Figure 16: Mass Properties Section View in the Open Position ................................................... 28

Figure 17: Rack and Pinion Assembly.......................................................................................... 29

Figure 18: HS-645MG Servo Characteristics ............................................................................... 30

Figure 19: Micro Maestro 6-channel Servo Controller Specifications and wiring diagram. ........ 30

Figure 20: Payload Retention Clips .............................................................................................. 31

Figure 21: Payload Assembly Housing Isometric Views ............................................................. 32

Figure 22: Payload Assembly Housing with components loaded................................................. 32

Figure 23: AirBrake Section Mass Properties............................................................................... 33

Figure 24: Electronics Bay section cut ......................................................................................... 34

Figure 25: Electronics Bay mass properties.................................................................................. 34

Figure 26: Computer Mounting Bay ............................................................................................. 35

Figure 27: Component Placement Height ..................................................................................... 35

Figure 28: Wiring Path.................................................................................................................. 36

Figure 29: Electronics bay on rails ............................................................................................... 36

Figure 30: Rail attachment location .............................................................................................. 37

Figure 31: TeleMega Flight Computer ......................................................................................... 37

Figure 32: PerfectFlite flight computer......................................................................................... 38

Figure 33: BeagleBone computer ................................................................................................. 39

Figure 34: Propulsion Bay assembly dimensions ......................................................................... 40


Figure 35: Propulsion Bay ............................................................................................................ 41

Figure 36: Cesaroni Pro 75 casing ................................................................................................ 42

Figure 37: Cesaroni Pro 75 casing dimensions ............................................................................. 42

Figure 38: Cesaroni Pro 75 ........................................................................................................... 42

Figure 39: FinSim Simulation ....................................................................................................... 43

Figure 40: Fin mounting location.................................................................................................. 44

Figure 41: Overall assembly mass, without motor ....................................................................... 48

Figure 42: Overall Launch Vehicle Dimensions........................................................................... 50

Figure 43: Iris Ultra 120'' Parachute ............................................................................................. 51

Figure 44: Black Powder Location ............................................................................................... 51

Figure 45: Recovery Sequence ..................................................................................................... 52

Figure 46: Steal eyebolt with bulkhead......................................................................................... 53

Figure 47: Eyebolt data sheet ........................................................................................................ 53

Figure 48: Archetype Rocketry Cable Cutter ............................................................................... 54

Figure 49: Suggested Manufacturer Configuration ...................................................................... 54

Figure 50: Redundant Cable Cutters ............................................................................................. 55

Figure 51: Recovery System Electrical Schematic ....................................................................... 56

Figure 52: AirBrake Bay ............................................................................................................... 59

Figure 53: AirBrake System Operational Diagram....................................................................... 59

Figure 54: System Components .................................................................................................... 60

Figure 55: AirBrake Closed .......................................................................................................... 61

Figure 56: Airbrake Open ............................................................................................................. 61

Figure 57: Grade 8 Steel Fully Threaded Rod .............................................................................. 62

Figure 58: End Fixity Factor ......................................................................................................... 63

Figure 59: FEA of basic steel lead screw...................................................................................... 66

Figure 60: FEA of Aluminum lead screw ..................................................................................... 66

Figure 61: Selected Stepper Motor ............................................................................................... 67

Figure 62: Stepper Motor Dimensions .......................................................................................... 68

Figure 63: Torque vs. RPM........................................................................................................... 68

Figure 64: Stepper Motor Design Alternative............................................................................... 69

Figure 65: Left - SBACB606DD-20 1/4 I.D bearing housing. Right - 6383K214 1/4 I.D bearing

....................................................................................................................................................... 69

Figure 66: Teflon Bushings........................................................................................................... 70

Figure 67: AirBrake Flap Linkage ................................................................................................ 70

Figure 68: Flap hinge mounting .................................................................................................... 71

Figure 69: AirBrake System Flaps ................................................................................................ 72


Figure 70: Airbrake Flap section cut............................................................................................. 72

Figure 71: AirBrake Flap mass properties .................................................................................... 73

Figure 72: Flap Dimensions .......................................................................................................... 75

Figure 73: CFD results with AirBrake closed............................................................................... 76

Figure 74: Nosecone tip stagnation pressure ................................................................................ 76

Figure 75: Five Degree angle flap deployment............................................................................. 77

Figure 76: Close-up of pressure distribution around flap ............................................................. 77

Figure 77: Pressure Close-up of 15 degree flap deployment ........................................................ 78

Figure 78: Airflow over flap at 15 degree deployment................................................................. 79

Figure 79: Pressure Distribution over Flap ................................................................................... 79

Figure 80: Pressure distribution at engine cut-off......................................................................... 80

Figure 81: Rocksim Schematic of Launch Vehicle ...................................................................... 89

Figure 82: Supplier-Provided Thrust Curve.................................................................................. 90

Figure 83: Simulated Thrust Curve............................................................................................... 91

Figure 84: CG/CP vs. Time, no wind............................................................................................ 92

Figure 85: Static Margin vs. Time, no wind ................................................................................. 93

Figure 86: No Wind Flight Profile ................................................................................................ 94

Figure 87: 5-mph Wind Flight Profile .......................................................................................... 94

Figure 88: 10-mph Wind Flight Trajectory................................................................................... 95

Figure 89: 15-mph Wind Flight Profile ........................................................................................ 96

Figure 90: 20-mph Flight Profile .................................................................................................. 96

Figure 91: Altitude vs. Time, no wind .......................................................................................... 97

Figure 92: Range (Drift) vs. Time, no wind ................................................................................. 98

Figure 93: Mach Number vs. Time, no wind ................................................................................ 98

Figure 94: Nosecone Interface .................................................................................................... 101

Figure 95: Parachute Bay to AirBrake Bay Interface ................................................................. 102

Figure 96: AirBrake Bay to Propulsion Bay Interface................................................................ 102

Figure 97: AGSE Sequence Flowchart ....................................................................................... 103

Figure 98: Launch Rail Assembly in Horizontal State ............................................................... 118

Figure 99: Launch Rail Assembly in Launch Position ............................................................... 118

Figure 100: Initial Launch Rail configuration ............................................................................ 120

Figure 101: Final Launch Rail configuration.............................................................................. 120

Figure 102: Launch configuration............................................................................................... 121

Figure 103: Sliding motion mechanism ...................................................................................... 122

Figure 104: Launch Rail Angular Velocity................................................................................. 123

Figure 105: Framework Base Body ............................................................................................ 125


Figure 106: Framing Piping Cross Section ................................................................................. 125

Figure 107: Basic Piping Specifications ..................................................................................... 126

Figure 108: Rover Ramp ............................................................................................................. 126

Figure 109: Two piece joint ........................................................................................................ 127

Figure 110: Three piece joint ...................................................................................................... 127

Figure 111: Articulated joint ....................................................................................................... 128

Figure 112: Repurposed car antenna........................................................................................... 129

Figure 113: PA-04 Linear Actuator ............................................................................................ 129

Figure 114: PA-04 Linear Actuator dimensions ......................................................................... 130

Figure 115: 12V DC, Speed vs. Load ......................................................................................... 131

Figure 116: 12V DC, Current vs. Load....................................................................................... 131

Figure 117: Load on Main Supporting Beam ............................................................................. 132

Figure 118: Supporting beam buckling simulation ..................................................................... 133

Figure 119: Top View, Ground support ...................................................................................... 134

Figure 120: Side View, Ground support ..................................................................................... 134

Figure 121: Ground Support Assembly Components ................................................................. 134

Figure 122: RaspberryPi computer ............................................................................................. 136

Figure 123: Standard Webcam.................................................................................................... 137

Figure 124: Continuous Servo .................................................................................................... 138

Figure 125: Standard Servo......................................................................................................... 139

Figure 126: Launch Vehicle Cost Chart ..................................................................................... 146

Figure 128: AGSE Budget Chart ................................................................................................ 148

Figure 130: CG/CP vs. Time, 5-mph wind ................................................................................. 151

Figure 131: Static Margin vs. Time, 5-mph wind ....................................................................... 151

Figure 132: Altitude vs. Time, 5-mph wind................................................................................ 152

Figure 133: Range (Drift) vs. Time, 5-mph speed ...................................................................... 152

Figure 134: Mach Number vs. Time, 5-mph wind ..................................................................... 153

Figure 135: CG/CP vs. Time, 10-mph wind ............................................................................... 153

Figure 136: Static Margin vs. Time, 10-mph wind ..................................................................... 154

Figure 137: Altitude vs. Time, 10-mph wind.............................................................................. 154

Figure 138: Range (Drift) vs. Time, 10-mph wind ..................................................................... 155

Figure 139: Mach Number vs. Time, 10-mph ............................................................................ 155




Figure 143: Range (Drift) vs. Time, 15-mph .............................................................................. 157


Figure 144: Mach Number vs. Time, 15-mph wind ................................................................... 158




Figure 148: Range (Drift) vs. Time, 20-mph .............................................................................. 160

Figure 149: Mach Number vs. Time, 20-mph wind ................................................................... 160


SECTION 1: SUMMARY OF PDR REPORT

I. Team Summary

University Name: Florida International University

Team Name: PantherWorks Space

Mailing Address: Florida International University

College of Engineering and Computing

10555 West Flagler Street

Miami, FL 33174

Launch Vehicle Name: Ad Astra

Mentor Name: Joseph Coverston

Certification: Level 2 Tripoli Rocketry Association

Contact Information: [email protected] | 407-754-6572

II. Launch Vehicle Summary

The following table outlines a basic overview of this year’s competition launch vehicle.

The launch vehicle preliminary design followed has focused on minimizing manufacturing

complexity, while keeping structural integrity and stability in flight.

Overall Length (in.) 138.8

Diameter (in.) 6.155

Mass (lbs.) 23.83

Motor Choice Cesaroni 4263L1350 – P

Recovery System Signal Main Deployment Table 1 : Launch Vehicle Overview

III. AGSE/Payload Summary

Dimension At resting position At launch position

Length 108.00” 108.00”

Width 68.81” 68.81”

Height 31.86” 104.46”

Weight 120 lbs Table 2 : Overal l Ground S upport Dimens ions

The table above outlines a summary of the basic dimensions of the updated ground

support assembly. The autonomous rover body will be 3D printed, and controlled by a

RaspberryPi computer. Its motion will be controlled by servos. The entire AGSE design focuses

on functionality and innovation on autonomous that could be used on Mars.

mailto:[email protected]


SECTION 2: CHANGES MADE SINCE PROPOSAL

I. Vehicle Criteria Changes

The competition launch vehicle has met with a number of changes intended to optimize

the design since the initial proposal stage. The following list summarized the changes

encountered since the proposal:

The parachute bay will now come before the payload in the launch vehicle, in order from

nosecone to motor bay.

The AirBrake bay dimension has been accounted for in the total launch vehicle

dimension. In addition, mechanism has been finalized, and components have been

chosen.

Payload bay design has changed from an “open and closed” type configuration to a

rotationally opened door.

Motor selection has been changed from a Cesaroni 3300L3200 – P to a Cesaroni

4263L1350 – P.

Recovery system has been changed from a six foot drogue and a ten foot main to only a

twelve foot main parachute.

The parachute bay location has been changed in order to decrease the number of

components needed to be ejected from the launch vehicle at separation. Previously, both the

nosecone and the payload bay had to be separated in order for the parachute to be deployed. This

configuration allows only the nosecone to need ejection for parachute deployment.

The total launch vehicle length now accounts for the air – braking bay, and the

mechanism that will be used to operate the flaps has been developed.

The payload bay design has been changed to save on both weight and length. Using an

actuator to open and close the payload bay proved to add a significant amount of weight to

launch vehicle, as well as, undesirably, increase the amount of length the payload bay would

need to take up.

Although the initial weight prediction in the conceptual phase of the design was mostly

accurate, the overall length of the vehicle changed significantly. Because of this, a motor with a

longer burn time was chosen in order to reach the desired apogee height.

In order to save on both overall weight and length of the vehicle, it was determined that

only a twelve foot main parachute was needed for the given application. The ejection of the


nosecone will be used to initialing destabilize the descent of the launch vehicle, and the

deployment of the main will sufficiently slow the descent rate of the vehicle for a safe landing.

II. AGSE Criteria Changes

The AGSE design has been fully developed from an initial concept idea to a fully

developed design. The support structure has been designed using readily available components,

and the system for lifting the launch vehicle into position and inserting the igniter has been

established. The following list summarizes changes made since the proposal:

The basic overall structure has been changed.

The ignitor device changed from a linear actuator to a standard telescoping antenna.

Launch rail movement has been changed and been developed, and the lifting mechanism

is more feasible.

The overall dimensions have been defined.

A more detailed computer – aided design model has been designed.

While the overall purpose of the AGSE hasn’t changed, the way the task is performed has

changed dramatically. In general, a much more detailed and refined version has been built. From

initial concept, with a basic intuition of how the AGSE performed, the current design has a much

more complete layout in terms of the components to be used and basic movement analysis. The

base of the AGSE originally had legs that placed the rail above the ground to allow the linear

actuator to clear the ground. The current design uses the truss structure to obtain the same

results; however, the linear actuator has been replaced with a car’s telescoping radio antenna.

The movement has been changed from the originally design, from rotating around the

blast plate to a planar motion that positions the rail and launch vehicle close to the middle of the

ground support itself to create a much more stable platform during launch. Finally, the lifting

mechanics went from a motor and gear box assembly to a linear actuator that would pull the rail

towards the center. This system, while slightly more complicated than the original, uses less

custom made parts; and the ones that are purchased have all been tested with a high fidelity of

information about their properties.

III. Project Plan Changes

No significant changes have been made to the project plan.


SECTION 3: VEHICLE CRITERIA

I. Launch Vehicle Design

MISSION STATEMENT

The FIU PantherWorks Space team will design and build an Autonomous Ground

Support Equipment (AGSE) capable of performing on – pad operations to prepare a high –

powered rocket for launch. The rocket will be designed to be capable of reaching an altitude no

greater than 5,280 feet above ground level. In addition, the AGSE will recover a payload located

outside the launch vehicle’s mold line and insert it into the delegated payload compartment.

MISSION VERIFICATION

Requirement Reasoning Verification

FIU ASME will design and

build a launch vehicle in a

timely manner consistent

with guidelines specified by

NASA Student Launch

officials.

FIU ASME wishes to

comply with all competition

requirements and does not

want to be penalized or

disqualified from the

competition.

FIU ASME will develop and

maintain a schedule for the

design, construction, and

testing of the launch vehicle

such that all requirements are

met by specified NASA

Student Launch deadlines.

FIU ASME will follow and

comply with all NAR rules

when conducting any testing

and launch procedures.

FIU ASME wishes to protect

the safety of its members as

well as the public present at

testing and launch events.

FIU ASME’s Safety Officer

will ensure that all of its team

members are educated in

safety practices and will

enforce safety in all aspects

of construction, design,

testing, and launch of the

vehicle.

FIU ASME will conduct a

subscale flight test of the

launch vehicle prior to the

full-scale flight test and prior

to CDR.

FIU ASME wishes to verify

that design choices for the

vehicle are valid by testing

them on a subscale rocket

before entrusting them to the

full-scale rocket. The

subscale flight test also

serves to satisfy a

competition requirement.

FIU ASME will follow its

project schedule to ensure

that both flight tests are

conducted in a timely manner

to ensure compliance with

NASA Student Launch

competition deadlines.

FIU ASME will complete a

full-scale test flight of the

vehicle prior to FRR in order

to validate vehicle design by

ensuring all parts function as

In addition to verifying

design choices, the full-scale

launch will serve to satisfy

NASA Student Launch

competition requirements.

FIU ASME will contact

several local rocketry groups

to ensure that different

options are available for the


designed and ensure that the

vehicle can remain launch-

ready for at least one hour.

location and time of test

launches.

Table 3 : Bas ic Mis s ion Verification Metrics

MISSION SUCCESS CRITERIA

The vehicle’s mission will be considered to be a success if the following criteria are met:

1) The vehicle’s apogee does not exceed 5600 ft. above ground level.

2) AirBrake successfully guilds rocket to 5280 ft.

3) The main parachute is deployed at 800 ft.

4) The vehicle’s descent is controlled and does not result in damage to itself, property, or

people.

5) No safety violations occur.

DESIGN OF LAUNCH VEHICLE

Figure 1 : Overal l Launch Vehicle As s embly

The team is focusing on overall efficiency and reusability of the launch vehicle by

employing a modular design. The launch vehicle designed this year features revamped versions


of certain systems; these revisions were done by using prior experiences as guides to improve

upon the quality and precision of the of all components and assemblies present in launch vehicle.

Figure 2 : Overal l Vehicle Configuration

Figure 2 shows the basic layout of all sub sections of the launch vehicle: nosecone bay,

main recovery bay, cache containment bay, electronics bay, air – braking bay, and propulsion

bay. The launch vehicle is designed to be made of Kraft Phenolic wrapped in carbon fiber and

will feature an air –braking system to insure target altitude is never surpassed.

Applicable Formulas

In order to accurately ascertain the stability and success of the rocket, three important

values must be calculated: peak altitude, center of gravity, and center of pressure. The peak

altitude is found through a specific sequence of equations. The average mass is first calculated

using:

In this equation, 𝑚𝑟 is the rocket mass, 𝑚𝑒 is the motor mass, and 𝑚𝑝 is the propellant

mass. The aerodynamic drag coefficient (kg/m) is further computed by:

In the equation above, 𝜌 is the air density (1.22 kg/m3), 𝐶D is the drag coefficient, and 𝐴

is the rocket cross-sectional area (m2). Equations 1 and 2 are used to calculate the burnout

velocity coefficient (m/s) using,


Here, 𝑇 is the motor thrust, and 𝑔 is the gravitational constant (9.81 m/s2). Equations 1, 2,

and 3 are then used to compute the burnout velocity decay coefficient (1/s) with the following

formula:

Then, equations 3 and 4 are used to calculate the burnout velocity (m/s) as follows:

where t is motor burnout time (s). The altitude at burnout can then be calculated by:

With the burnout altitude having been calculated, the coasting distance can be found by

first beginning with the value of the coasting mass which is calculated as follows:

The average mass in equations 3 and 4 is replaced with the coasting mass. This

replacement results in equations 8 and 9 for the coasting velocity coefficient and coasting

velocity decay coefficient, respectively:

Equations 8 and 9 are subsequently used to calculate the coasting velocity (m/s) using:


Furthermore, the coasting distance can then be computed as follows:

From the coasting distance calculation, the peak altitude can be found as follows:

The center of gravity location is calculated using:

where W is the total weight, d is the distance between the denoted rocket section center of

gravity (nose, rocket, body, engine, and fins, respectively) and the aft end. Moreover, the center

of pressure measured from the nose tip can be found using this equation:

In this equation, the (CN)N is the nose cone center of pressure coefficient, and the 𝑋N is

computed by this formula:

where 𝐿N is the nose cone length. The (CN)F in equation 14 is the fin center of pressure

coefficient calculated using the following equation:


The variables in this equation are defined as follows, R is the radius of the body at the aft

end, S is the fin semi – span, N is the number of fins, Lf is the length of the fin mid – chord line,

CR is the fin root chord length, and 𝐶T is the fin tip chord length. The final variable in equation

14, 𝑋f, is calculated using;

where 𝑋B is the distance from the nose tip to the fin root chord leading edge and XR is the

distance between the fin root leading edge and the fin tip leading edge measured parallel to body.

Equations 14 through 17 are also known as the Barrowman Equations (The Theoretical

Prediction of the Center of Pressure, 1966).

Stability and Construction

The launch vehicle airframe will be constructed primarily of Kraft Phenolic wrapped in

carbon fiber, and the internal structure will be constructed out of fiberglass, plywood, ABS

plastic, and aluminum. The vehicle is designed to house a cache capsule payload within its

airframe. To ensure an efficient design, the launch vehicle has been designed to use as much

internal space as reasonably possible.

Figure 3 : Rock s im S imulation of Launch Vehicle

The vehicle is designed such that the payload bay will be located directly beneath the

main recovery system. This allows one of the heavier systems in the vehicle to sit high up in the

rocket, thus raising the center of gravity and increasing stability. The figure above also shows the

location of electronic bay, right below the payload compartment. The AirBrake system is housed

below the electronic bay, followed by the propulsion bay.


DESIGNS AT A SYSTEMS LEVEL

The following sections serve to analyze the design for each subsystem and reiterate the

individual system’s requirements.

Nosecone Design

The design of this year’s launch vehicle utilizes a Von Kármán nosecone.

Figure 4 : Von Karman nos econe , current model

The following equations were used to create the shape of the Von Kármán nosecone in a

3D computer – aided design modeling software to insure accurate simulation results.

(18)

(19)


The variables in the previous two equations are defined

as follows:

C = 0 for Von Kármán

L = Length

R = Radius of base for nose cone

x = Distance from tip of nose cone

Selection Rationale

The Von Kármán was the initial first choice for our team’s design, due to its optimal

performance at subsonic and transonic speeds. Unlike other nosecone shapes, the Von Kármán

nosecone is mathematically derived for the purpose of minimizing drag. The following equations

were used to create the shape of the Von Kármán nosecone in a 3D computer – aided design

modeling software to insure accurate simulation results. The geometry can then be exported to a

CNC lathe; allowing the nosecone to be manufactured with ease. Fiberglass was chosen as the

nosecone material because of its lightweight characteristics and strong material properties.

Figure 6 : Rational for Von Karman nos econe

Figure 5 : Von Kármán nos econe , propos al CAD model


The nosecone is at a 5:1 ratio, with a length of thirty inches from the shoulder to the tip

with a 6.155 inch base.

Characteristics

Figure 8 : Back View of Nos econe

The back of the nosecone is enclosed with fiberglass in order to ensure that the ejection

events push the parachute out of the upper airframe tube, allowing the main parachute to open.

Figure 7 : Bas e Nos econe Dimens ions


Figure 9 : Nos econe Mas s Properties

The above image demonstrates the mass properties given by the computer software

program used to model the launch vehicle assembly.

Airframe

Selection Rationale

By using commercially available phenolic tubing, the ease of being able to use pre –

made bulkheads and coupler sections can be taken advantage of, simplifying the manufacturing

process. The carbon fiber overlay is used to increase the rigidity and load – bearing capacity of

the airframe. The outer layer of carbon fiber provides the airframe with excellent axial material

properties. Overlaying the phenolic with a carbon fiber layer the otherwise somewhat brittle

phenolic. Pure carbon fiber airframe were considered, but the increased cost of both couplers and

the airframe causes it to be quickly eliminated from our design selection.

Figure 10: Launch Vehicle Airframe as s embly


Characteristics

Having the team manufacture the carbon fiber wrapped phenolic airframe sections in

house enables allows for a saving in cost and the ability to quality check each airframe as it is

being manufactured. This provides us with an inexpensive way to manufacture carbon fiber

reinforced airframe sections. Replacement couplers, centering rings and bulkheads are all readily

available for commercial purchase. Furthermore, carbon fiber wrapped phenolic is strong and

resistant to hard landings. The following two tables outline material properties for both carbon

fiber and phenolic tubing.

Table 4 : Carbon Fiber Material Properties

Table 5 : Phenol ic Tubing Material Properties


Parachute Bay

Figure 11: Parachute Bay Airframe

The parachute bay and the payload bay will be both be housed in a single phenolic

airframe tubing. This was done to insure space was used efficiently, making this the longest bay

of our launch vehicle. These two systems will be placed in one full length, 48’’ section of

phenolic tubing. A bulkhead will separate the parachute bay from the payload bay. Finite

element analysis (FEA) simulations will be conducted on the bulkhead and airframe structure to

insure components can handle a black powder charge to be used for parachute ejection. The

following two images show the overall dimensions and mass properties of this section.

Figure 12: S ection View of Parachute Bay


Figure 13: Mas s Properties of parachute bay

Further analysis of specifics related to the recovery subsystem may be found in section 2

of the vehicle criteria portion of this report.

Payload Bay

Figure 14: Payload Bay S ubs ys tem

The payload bay will be accessible to the AGSE by separation of the airframe into two

halves using a rack and pinion system. A housing for the electronics operating the rack and

pinion system will be 3-D printed out of ABS plastic. The encasing will be JB welded to the

lower tube and mounted to a bulkhead by a #6-32 socket head screw. A balsa wood plate is

mounted on top of the housing using four ¼ inch counter bore socket head screws as well. Two

¾ inch retention clips will also be printed from ABS and mounted above the balsa wood plate in

order to safely secure the payload cache into the assembly. A pin and spring mechanism acting

through the rack will provide a lock for the subsystem during flight and ejection of parachute.


Figure 15: Is ometric View with Airframe removed

Figure 16: Mas s Properties S ection View in the Open Pos i tion

As can be seen from the Figure above, optimizing the payload bay to be as compact and

lightweight as possible while maintaining the safety of the payload cache is our main goal for the

efficiency of this subsystem.

Payload Bay Components

Component Description

1 – Rack Assembly 32 Pitch 6.54 in. Delrin Rack and Aluminum Beam

2 – Servo Motor 133 Oz-in Continuous Rotation Motor and 32 Pitch Gear

3 – Retention Clip ABS Plastic

4 – Battery Pack 6V 1600mah NiMH Battery

1 2

3

6

4

5

7 8


5 – Locking Mechanism Metal Spring and pin rod lock

6 – Wiring Door Balsa Wood

7 – Micro Maestro 6-Channel USB Highly Compact Servo Controller

8 – Assembly Component Housing ABS Plastic Table 6 : Payload Components

Rack and Pinion Assembly

A 32 pitch rack made of Delrin was chosen due to its light weight and durable

characteristics. The rack is attached by a #6-32 socket head cap screw with a Nylon nut to an

aluminum beam to provide more support for the upper portion of the airframe as is moves out of

the channel within the assembly housing.

Figure 17: Rack and Pinion As s embly

The bolt also holds steel two right angle brackets that are mounted to the upper bulkhead

containing the eye bolt for the parachute bay. This will allow the entire upper portion of the

airframe to be pushed forward on the launch rail upon interaction with the AGSE.

The HS-645MG Ultra Torque servo was chosen according to the amount of force required to

displace the upper portion of the launch vehicle 5 inches in order to for the AGSE to gain access

to the retention clips where the payload cache will be held. This particular type of motor

provides the most efficient capabilities over gear motors and stepper motors according to the

application it is needed for. It is lightweight, 1.94 oz., in comparison to gear motors, and

provides continuous rotation with precision position control from the potentiometer, whereas

stepper motors have a tendency to lose position due to the magnetic field used to drive it. The

figure below provides further characteristics.


Figure 18: HS -645MG S ervo Characteris tics

Payload Electronics

The HS-645MG servo will be controlled by the Micro Maestro 6-channel USB servo

controller. This controller a bit bigger than a dime and provides all that is needed to control the

servo for the application at hand. With internal scripting control and its own USB program

interface for position control it is a versatile controller that can easily get the job done at low

cost, . Other controller’s common controllers such as an Arduino UNO were considered by

proved to be too big for the compact space provided. The high-resolution pulse range, 64-3280

microseconds, makes it perfect for a high performance application and reliability. This will allow

the servo’s position to be controlled over the time frame of the cache loading. The figure below

provides more specifications on the microcontroller.

Figure 19: Micro Maes tro 6 -channel S ervo Control ler S peci fications and wiring diagram.


Both the microcontroller and servo will be powered by a 6V 1600mah NiMH

rechargeable power supply. This is a very high capacity battery pack in a small size providing

better efficiency than NiCAD batteries. LiFe and LiPo batteries were considered but were not as

cost effective as the currently chosen pack providing the necessary power for the equipment

being used.

Retention Clips

In order to securely transport the payload throughout flight in the launch vehicle,

retention clips made of ABS plastic were an optimal solution. Once the payload assembly is in

the open position the payload can then be placed between the clips providing a snug compression

fit to reduce and sloshing or bouncing. The clips not only can be manufactured and tested in

house, but can be optimized according to the test data.

Figure 20: Payload Retention Cl ips

The clips not only provide great stability for the payload but are also very light weight.

The angled ends of the clips provide guidance as the cache is brought closer into the center of the

clip. The clips will be 3D printed in ABS plastic due to its flexibility as compared to commonly

used PLA plastic.

Payload Assembly Housing

The assembly housing for the payload is an integral component to allow this system to

perform properly. The following figures provide insight into the detail of the design.


Figure 21: Payload As s embly Hous ing Is ometric Views

Figure 22: Payload As s embly Hous ing with components loaded

As can be seen from above the housing will be unique to the design and 3D printed out of

PLA plastic due to the necessity for a more detailed design and higher strength and rigidity than

ABS. Each component has its own location within the housing to be mounted with #6-32 socket

head cap screws. The rack and beam have a channel that it will move within in order to prevent it

from dislodging and contacting other components within the housing. The battery pack has a

small hole next to its placement for the wiring to enter through, while there is also a hole for the

wiring of the servo as well. A small balsa wood door will block off the excess wiring from both

components and provide a common hole for them to enter through and attach to the

microcontroller. This solves the problem for excess wiring or requirement of a harness.

This housing is attached to the airframe tube by JB weld, a highly efficient epoxy capable

of withstanding the proposed load rating from our design at ejection of the parachute. The load

from ejection on the eyebolt will be transferred down the rack and aluminum beam which was

engaged by a spring pin lock on the final hole. The load will then be dispersed from the metal pin


over the assembly housing which is attached to the airframe by JB weld. The lower bulkhead is

attached to the housing by a #6-32 socket head cap screw.

AirBrake Bay

Figure 23: AirBrak e S ection Mas s Properties

The success of the launch vehicle for this competition is determined by two key factors:

an apogee height of one mile and launch vehicle stability during its ascent. In order to achieve

am apogee height of as close to one mile as possible, the PantherWorks team is designing a

continuously self-adjusting AirBrake system. The launch vehicle will overshoot the one mile

goal at launch, and approximately half – way through its trajectory deploy the proposed

AirBrake system. It will consist of flaps that use information from the altimeter to calculate

projected maximum altitude and real – time trajectory.

This system is not just designed for a target apogee of 5280ft, it can be reprogramed to

6000ft, 8000ft, or 10,000ft. This allows our launch vehicle to be capable of a wide range of

altitude marks. Further analysis into the detail of this subsystem may be found in section 3 of the

Launch Vehicle Design portion of this report.

Electronics Bay

The electronics bay will house the two flight computers and the AirBrake computer, in

addition to the batteries needed to power the three computers and the AirBrake stepper motor.


Figure 24: Electronics Bay s ection cut

Figure 25: Electronics Bay mas s properties

The flight computers will be mounted on one of the 3 surfaces of the triangle seen in

Figure 17 below. This triangle design allows 45.3 square inches of area on each surface for

mounting components with a clearance of 1.80 inches, as seen in Figure 18. The electronics bay

will also house at least three 9 volt batteries, one for each flight computer.


Figure 26: Computer Mounting Bay

Figure 27: Component Placement Height

Wires will be organized thorough the center of the triangle, seen in Figure 19. This

allows for a main wiring harness which connects all the flight computers and the stepper motor

and clears the wiring harness from any tangling obstacles.


Figure 28: Wiring Path

The electronics bay is planned to be 3D printed in ABS plastic. This allows us to design

mounting locations for all three computers into the design to be 3D printed, saving the extra time

and weight that would be needed if mounting would need to be design. In addition, no wiring

holes will have to be drilled. ABS plastic has excellent material properties and can handle heat

better than other 3D printed plastics.

Figure 29: Electronics bay on rai ls

The electronics bay is designed to be on metal rails for easy access to any flight

components. If repairs are needed, launch vehicle can be taken apart at the AirBrake section and

electronics bay pulled out to be worked on. The rails also act as reinforcement for the airframe

and will be mounted, on one end, to the bulkhead between AirBrake and electronics bay

compartments.


Figure 30: Rai l attachment location

Selection Rationale

The function of the electronics bay is to safely house flight electronics used to record

flight data during launch and landing, as well as perform ejection events and AirBrake control at

specified altitudes. Due to the need for redundancy, it will need to house two flight computers

and a GPS unit along with the required power for each.

In addition to strength requirements, the electronics bay must be easily accessible and

removable for electronics maintenance and charge reloading. The triangle design allows this as

the electronics are mounted on the surface. Once the electronics bay is pulled out the electronics

can be worked on without removing them from the bay, allowing for simple repairs.

Electronics

Figure 31: TeleMega Fl ight Computer

The TeleMega flight computer is a high – end recording, dual – deploy altimeter for high

power model rocketry with integrated GPS and telemetry link. The features included make

TeleMega the ideal choice for complex projects. In particular, pyrotechnic events are


configurable and can be based on time and various flight events and status, including angle from

vertical. The following tables outlines the specifications of this computer.

Telemega Key Features:

Recording altimeter for high power model rocketry

Supports dual deployment and 4 additional pyro events.

Pyro events are configurable and can be based on time and various flight events and

status, including angle from vertical (for safety in staging and air start flights).

70cm ham-band transceiver for telemetry downlink

Barometric pressure sensor good to 100k feet MSL

1-axis 105-g accelerometer for motor characterization

3-axis 16-g accelerometer for gyro calibration

3-axis 2000 deg/sec gyros

3-axis magnetic sensor

On-board, integrated GPS receiver

On-board non-volatile memory for flight data storage

USB for power, configuration, and data recovery

Integrated support for LiPo rechargeable batteries

User choice of pyro battery configuration, can use primary LiPo or any customer-chosen

separate pyro battery up to 12 volts nominal.

3.25 x 1.25 x 0.625 inch board designed to fit inside 38mm airframe coupler tube

Weight: 25g (0.88oz) Table 7 : Telemega S peci ficati ons

Figure 32: PerfectFl i te fl ight computer

The StratoLoggerCF collects flight data (altitude, temperature, and battery voltage) at a

rate of 20 samples per second throughout the flight and stores them for later download to a

power removed. The following table outlines the specifications of this flight computer.

StratoLoggerCF Key Features

Works to 100,000 feet MSL, audibly reports peak altitude and maximum velocity after

flight.

Stores 16 flights of 18 minutes each (altitude, temperature, and battery voltage at 20

samples per second) for download to a computer with the optional DT4U USB interface.

Hi-speed sampling and storage of battery voltage serves as a useful aid in diagnosing

intermittent problems with your battery, switch, and wiring. All data are preserved

with power off.

Deploys drogue and main chutes with audible ematch continuity check.


Outputs capable of 5A current for 1 full second to allow use with nearly any ematch or

ematch substitute. Reverse polarity protection prevents spontaneous firing if battery is

connected backwards.

Main chute deployment altitude is adjustable from 100 feet to 9,999 feet in 1 foot

increments. 9 presets allow for quick change in the field.

No Mach delay necessary for Mach+ flights: Automatic Mach Lock assures proper

operation with any flight.

Brownout protection will tolerate 2 second power loss in flight – no need for multiple

batteries.

Precision sensor & 24 bit ADC yield superb 0.1% accuracy.

Built-in voltmeter reports battery voltage on power up – no more guessing about battery

condition.

Post flight locator siren aids in locating your rocket.

Confusion-free individual terminal blocks – unreliable multiple wires per terminal are

not necessary. Dedicated switch terminal block eliminates the need for splicing switch

into battery wire.

Highly resistant to false trigger from wind gusts; tested in 100+ MPH winds!

Selectable apogee delay for dual altimeter setups prevents overpressure from

simultaneous charge firing.

Low power design runs for weeks on a standard 9V alkaline battery. Post-flight locator

siren will run for months, giving you multiple “second chances” to find a lost rocket.

Telemetry output for real-time data in flight with your RF link.

Rugged SMD construction, stringent QC testing, and internal self-diagnostics assure

uncompromising reliability.

Wide operating temperature range of -40F to +185F.

Measures just 2.0"L x 0.84"W x 0.5"H, fits 24mm tube, weighs just 0.38 oz. Table 8 : Perfect Fl ight S peci fications

Figure 33: BeagleBone computer


The BeagleBone is credit card sized Linux with plenty of I/O and processing power for

real-time analysis provided by an AM335x 720MHz ARM® processor. The BeagleBone can be

complimented with cape plug – in boards to augment functionality.

This will be the dedicated computer for the AirBrake stepper motor. It will perform all

flight calculations need to deploy the AirBrake system. The BeagleBone will receive flight data

from the flight computer for its altitude projection calculations.

Propulsion Bay

Figure 34: Propuls ion Bay as s embly dimens ions


Figure 35: Propuls ion Bay

Component Description

1 – Engine Casing 6 grain aluminum engine casing

2 – Coupler Tube 12 in. coupler

3 – Bulkhead Provides engine and fin support

4 – G10 Fins Provides stability Table 9 : Propuls ion component des criptions

The propulsion bay will serve two specific purposes: as the connection point for the fins

and the motor and motor case housing. Our propulsion bay is design to be modular and easily

interchangeable with a larger or smaller propulsion bay as the needed.

Selection Rationale

The lower airframe consists of the lower body tube, the motor mount and the fins. This

portion of the rocket will be exposed to the largest forces of any rocket section. As a result an

emphasis must be placed on obtaining the sturdiest design possible.

The rationale of making the propulsion bay interchangeable from the launch vehicle

allows the ability to fulfill multiple mission profiles. Interchanging a larger engine allows us to

achieve a higher altitude.

Characteristics

The engine casing will be flush against the airframe coupler to optimize space, as can be seen in

Figure 26. Three trapezoidal G10 fiberglass fins will be attached to the propulsion bay and will

be both surface mounted and center ring mounted. The engine casing will be held in place by

centering rings.


Engine Casing

Figure 36: Ces aroni Pro 75 cas ing

Figure 37: Ces aroni Pro 75 cas ing dimens ions

Cesaroni Pro 75 rocket motor hardware is CNC machined from 6061 – T6 aluminum

alloy and anodized for corrosion protection. Five components make up a complete Pro75 motor

hardware set: the motor casing, the forward closure, the nozzle holder (rear closure), and the one

threaded retaining rings for each end of the case. An accessory wrench is available which make

fitting and removing the retaining rings simple.

Figure 38: Ces aroni Pro 75


Fins

Three trapezoidal G10 fiberglass fins will attach near the end of the propulsion bay to provide

stability to the launch vehicle. Trapezoidal G10 fiberglass fins where selected based on past

years’ experience. They are simple and cost effective to manufacture. A thickness of a quarter

inch was selected for the thickness due to fin divergence and flutter calculations. FinSim was

used to test the fin design, as can be seen in the following image.

Figure 39: FinS im S imulation

Fin flutter can rip our propulsion bay apart. With a G10 fin thickness of a quarter inch,

the flutter speed is 2768.22 ft/s. Our maximum velocity is 765 ft/s, well under the simulated

flutter speed of the fins being used.


Figure 40: Fin mounting location

The fins will have a root cord length of 12 inches, a tip cord of 3 inches, a semi- span of 5

inches, and a front sweep of 3 inches. Fins will be surface mounted and mounted on the engine

centering rings, as can be seen in Figure 31 above. Fin will be mounted using G5000 high

strength epoxy or an equivalent epoxy. The epoxy will be placed as filets on the surface and on

the sides of the centering rings. The specifications of this epoxy can be found in the table below.

Specific gravity, Resin 1.52

Specific gravity, Hardener 1.48

Specific gravity mixed 1.5

Tensile strength 7,600 psi

Compression strength 14,800 psi

Elongation at break % 6.3% Table 10: G5000 S peci fications

LAUNCH VEHICLE VERIFICATION

The table below outlines the verification of all statements of work provided by the

competition for the launch vehicle design.

Requirement

Number

Requirement Design Feature Verification Method


1.1 The vehicle shall deliver the

payload to an apogee

altitude of 5,280 feet above

ground level (AGL).

The vehicle has been

designed to a certain length

and weight in order to reach

the given apogee height on

an L – class motor and the

air – braking system.

The apogee height analysis

will be done using software

such as Rocksim, while also

being supported by rough,

general hand calculations for

redundancy. The AirBrake

will be verified with the use

of a computer code to

determine in – flight

adjustment requirements.

Both systems will be tested

during the subscale flight

test.

1.2 The vehicle shall carry one

commercially available,

barometric altimeter for

recording the official altitude

used in the competition

scoring.

The launch vehicle

electronics bay will house

two altimeters for

redundancy purposes.

The altimeters will be tested

by inspection upon purchase,

in addition to being tested on

the vehicle subscale flight.

1.3 The launch vehicle shall be

designed to be recoverable

and reusable. Reusable is

defined as being able to

launch again on the same

day without repairs or

modifications.

The launch vehicle is being

designed to be a modular

vehicle. Any one section

may be easily replaced with

a new section in an

emergency. In addition, all

components in the vehicle

can be accessed and repaired

given any kind of

circumstance on the day of

launch.

Manufacturing processes

will be kept secure enough to

withstand the force of launch

and landing, but will be kept

simple enough to allow for

repairs. Spare parts will be

available on competition day

for any small, necessary

repairs.

1.4 The launch vehicle shall

have a maximum of four (4)

independent sections. An

independent section is

defined as a section that is

either tethered to the main

vehicle or is recovered

separately from the main

vehicle using its own

parachute.

The recovery scheme for the

launch vehicle allows for

only one main separation

event. The nosecone will

land alongside the rest of the

vehicle tethered by a shock

cord. This corresponds to

only two independent

sections.

The shock cord tethering the

nosecone will be purchased

to exceed the force needed to

withstand ejection and

landing. Furthermore, the

shock cord will be tested

upon purchase.


limited to a single stage.

The vehicle will only house

one motor in the main

propulsion bay.

The design of the vehicle

will be inherently limited to

one motor by the

construction of only one

propulsion bay.



capable of being prepared

for flight at the launch site

within 2 hours, from the

time the Federal Aviation

Administration flight waiver

opens.

The entire team will be

trained on vehicle

integration and assembly

launch day procedures.

Each member will practice

their individual tasks

multiple times before launch

day, and the entire team will

practice a full assembly

procedure until the process is

smooth and organized.


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 power sources for the on

– board flight computers will

be fully charged prior to

launch.

The ability of all batteries to

stay charged enough to be

fully operational throughout

an hour’s time will be

empirically tested before the

competition date.


capable of being launched

by a standard 12 volt direct

current firing system.

The firing system will be

provided by the NASA-

designated Range Services

Provider.

The propulsion system of the

launch vehicle is being

designed to have the ability

of being launched by a

standard 12 volt direct

current firing system.

The team will verify this

requirement during the

subscale launch, and before

the competition day launch.

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 motor being used for the

launch vehicle is

commercially available for

purchase.

Any change in motor choice

between the preliminary and

critical design reviews will

entail only certified,

commercially available

motors.

1.10 The total impulse provided

by a launch vehicle shall not

exceed 5,120 Newton-

seconds (L-class).

The total impulse on the

motor selected does not

exceed the given value of

5,120 Newton – seconds.

This is verified by the data

provided by the

manufacturer. The

specifications of the motor

will be carefully checked if

any change in motor occurs.

1.11 Pressure vessels on the

vehicle shall be approved by

the RSO and shall meet the

following criteria:

There are no pressure

vessels present or planned in

the current launch vehicle

design.

Appropriate measures will

be taken if a change in

design requires the presence

of a pressure vessel.


1.11.1. The minimum factor

of safety (Burst or Ultimate

pressure versus Max

Expected Operating

Pressure) shall be 4:1 with

supporting design

documentation included in

all milestone reviews.

1.11.2. Each pressure vessel

shall include a pressure

relief valve that sees the full

pressure of the tank.

1.12 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 has a subscale

launch planned for the first

few days of 2016.

This launch will serve as a

proof of concept for all

pertinent subsystems on the

launch vehicle.

1.13 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

A test launch for the full –

scale launch vehicle is

planned for mid – February

of 2016.

All components of each

subsystem will be tested on

the test flight for verification

of all processes.

Table 11: Vehicle S tatement of Work Verification


CRITICAL MASS STATEMENT

Figure 41: Overal l as s embly mas s , without motor

The estimated mass of the launch vehicle is 26.18 pounds. This is based on CAD

calculated mass. Rocksim gives a similar mass of 26.75 pounds.

The basis of the first mass estimation is CAD calculated mass, where the software takes

into account material volume and density to calculate mass. Based on past experiences, the CAD

software is accurate if done correctly.

Sources of error can develop from assigning wrong material value to a component, or by

incorrectly creating a material if it is not already found on the software program’s library.


This calculated mass was double checked by Rocksim, where the components have

weight given to them straight from the manufacturer – provided data. Since both mass

estimations differ by only half a pound, this mass estimate can be assumed to be correct. As

component design becomes more detailed, the 3D CAD software mass tool will become more

valuable at mass predictions, since Rocksim only has a limited amount of built in components.

Mass properties of 001 Top Down Asm. Rev1

Configuration: Default

Coordinate system: -- default --

Mass = 26.18 pounds

Volume = 1206.97 cubic inches

Surface area = 13205.86 square inches

Center of mass: ( inches )

X = 0.00

Y = -0.00

Along axial body Z = 9.65

Principal axes of inertia and principal moments of inertia: ( pounds * square inches )

Taken at the center of mass.

Ix = (0.00, -0.00, 1.00) Px = 259.42

Iy = (0.74, 0.67, 0.00) Py = 27875.14

Iz = (-0.67, 0.74, 0.00) Pz = 27878.72

Moments of inertia: ( pounds * square inches )

Taken at the center of mass and aligned with the output coordinate system.

Lxx = 27876.76 Lxy = 1.78 Lxz = -0.00

Lyx = 1.78 Lyy = 27877.10 Lyz = -0.06

Lzx = -0.00 Lzy = -0.06 Lzz = 259.42

Moments of inertia: ( pounds * square inches )

Taken at the output coordinate system.

Ixx = 30313.41 Ixy = 1.78 Ixz = 0.06

Iyx = 1.78 Iyy = 30313.75 Iyz = -0.12

Izx = 0.06 Izy = -0.12 Izz = 259.42

Table 12: Overal l As s embly Mas s Properties

The team plans on a minimal mass growth of 25%. This is due to epoxy and design

components becoming more fully defined. Weight addition can also come in unforeseen

reinforcement of critical areas.

Due to CAD mass calculations, the team can accurately track gain in mass throughout the

design process. In a worst case scenario in which we become too heavy to reach a one mile

apogee height, the engine casing is large enough to select a larger engine, allowing us to reach

the target altitude. Based on Rocksim calculations, a six pound gain in mass will make our

vehicle too heavy reach the target altitude on the chosen motor.

Launch Vehicle Section Section Length (in.) Mass (lbs.) 25% Mass Increase

Nosecone 30.0 2.47 3.09

Parachute Bay 24.0 7 8.75

Payload Bay 10.8 2 2.5


Electronics Bay 12.0 2 2.5

AirBrake Bay 9.5 4.16 5.2

Propulsion Bay (Without

motor casing

30.80 6.81 8.51

Motor Wight (With

casing)

36.89 8.88 11.1

Total Weight 33.32 41.65 Table 13: Individual S ubs ys tem Mas s es

FULL LAUNCH VEHICLE ASSEMBLY

Figure 42: Overal l Launch Vehicle Dimens ions

II. Recovery Subsystem

PARACHUTE SELECTION RATIONALE

The performance characteristics from three parachute geometries were compared to select

the optimal geometry for this year’s competition recovery system.

Parachute Drag Coefficient Descent Rate

with ~33 lbs

Size Packing

distance in 6in

diameter

Sky Angle CERT

– 3TM Series

2.59 17 ft./s 132 in. 14in

Iris Ultra 2.2 15 ft./s 120 in. 10in

Pro-Experimental

1.9 – Rocket Man

.8 17.45 ft./s 144 in. N/A

Table 14: Parachute S election Chart

The Iris Ultra 120 inch parachute was selected for the main parachute due to its greater

efficiency for the given application when compared to that of the other options. Its high drag

coefficient of 2.2 allowed for low descent rate. Another deciding factor in this choice was the low

packing height, in a 6 inch tube, that this parachute provided. Since this parachute will support entire


weight of the descending launch vehicle, a low descent rate was valued. With this type of parachute

we have the option of selecting a high power toroidal or a more compact toroidal depending on

the development of the current vehicle design. Both of these derivatives of the Iris Ultra still

maintain a drag coefficient of 2.2. The bridal of this parachute is one inch webbing with a swivel

that can support 6000 lbs, with the option of a quarter inch Kevlar bridal. The shroud lines are

Nylon llla 400# Paraline and the overall parachute has a weight of 49oz.

Figure 43: Iris Ultra 120 ' ' Parachute

The final factor in deciding on this parachute came with the kinetic energy calculations

that were done to ensure that the parachute chosen would allow the launch vehicle to land with a

kinetic energy of less than 75 ft – lbf.

RECOVERY FLIGHT PATH

At apogee, a black powder charge located above the payload bulkhead (see Figure 15

below) will eject the nosecone from the rest of the launch vehicle airframe. This will, in turn,

pull the main parachute out of the airframe. This is demonstrated in step 1 of Figure 16.

Figure 44: Black Powder Location

The main parachute will be ejected restrained by a plastic cable tie, which can be seen in

step 2 of Figure 16. This restrained main will function more like a drogue parachute by

destabilizing and slowing down the descent of the launch vehicle. Simultaneously, the AirBrake

system will open to its maximum angle to further slow the descent of the vehicle. At 800 feet, a cable


cutter will release the main parachute, allowing it to open. At this point, the AirBrake system will

retract its flaps so that they are not damaged during the impact of landing, as seen in step 3 of Figure

16.

Figure 45: Recovery S equence

Event Altitude Segment Event Description

1 5280 ft. Apogee. Nose cone ejection. Entire launch vehicle under restrained main parachute acting as drogue.

2 5280 ft. – 800 ft. Main restrained with cable ties.

AirBrake at full flap deployment.

3 800 ft. Main cable ties are cut allowing main to open. Airbrake flaps retract for landing

Table 15: Recovery Events and Des criptions

RECOVERY SYSTEM COMPONENTS

Shock Cable and Eyebolt

The parachute shock cable will be attached to a steel, nylon – coated eyebolt, which can

be seen below in Figures 17 and 18. The eyebolt is rated for a vertical load capacity of 2600 lbs.

This eyebolt will be attached to a half inch thick wooden bulkhead and half inch locking hex nut.


Figure 46: S teal eyebolt with bulk head

Figure 47: Eyebolt data s heet


Parachute Cable Cutter

Figure 48: Archetype Rock etry Cable Cutter

The cable cutter takes a different perspective on a tether by shearing a plastic cable tie,

which can be used to hold a large number of recovery components. One of the most beneficial

aspects of the cable cutter is by providing the ability to use a single parachute as both the drogue

and the main parachute. Its small size of approximately 1.8 inch length and 0.37 inch diameter

allows it to be placed anywhere in the parachute bay.

Figure 49: S ugges ted Manufacturer Configuration

The cable cutter uses a black powder charge to fire a shearing pin towards the cable tie,

consequentially shearing it. This process will occur in the enclosed airframe area, and the pin is

fired inward so no damage can befall the user or the launch vehicle.


Figure 50: Redundant Cable Cutters

In case of an altimeter failure, a secondary altimeter and cable cutter combination will be

available to successfully fire and deploy the recovery equipment. Two cable cutters will be

placed on the same cable tie. This ensures that if one fails, the other can cut the cable.


ELECTRICAL SCHEMATIC

Figure 51: Recovery S ys tem Electrical S chematic

RECOVERY SYSTEM VERIFICATION

The table which follows outlines the verification guidelines for the competition

requirements pertinent to the recovery system of the launch vehicle.

Requirement

Number


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.

The vehicle recovery has been

designed for a tumble recovery

at apogee and a main

deployment at 800 feet. The

tumble recovery will consist of

the nosecone being ejected off

the upper airframe, pulling the

The recovery system will be

tested empirically both on

the subscale and final full

scale flight.


Tumble recovery or streamer

recovery from apogee to

main parachute deployment

is also permissible, provided

the kinetic energy during

drogue-stage descent is

reasonable, as deemed by

the Range Safety Officer.

main parachute out of the

parachute bay. At 800 feet, a

cable cutter will be used to

deploy the main parachute.

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.

The parachute electronics bay

will not be integrated into the

subscale and final, full scale

assembly until the recovery

system ejection has been tested

for functionality.

The team will perform

ejection tests before the

final vehicle assembly is

built to ensure the system

will deploy correctly.

2.3 At landing, each

independent section of the

launch vehicle shall have a

maximum kinetic energy of

75 ft-lbf.

The chosen parachute has a

coefficient of drag allowing the

rocket to slow down

sufficiently for a safe descent.

Hand calculations have been

done for landing kinetic

energy alongside

verification from Rocksim

simulation values.

2.4 The recovery system

electrical circuits shall be

completely independent of

any payload electrical

circuits.

The design of each subsystem

has been done so that the

payload electrical circuits are

not being integrated with the

flight computer electrical

circuits.

The team will be aware of

the wiring of both the

payload and the electronics

bay in order to avoid

integrating the two.

2.5 The recovery system shall

contain redundant,

commercially available

altimeters.

The launch vehicle recovery

system has been designed to

include two altimeters for

redundancy purposes.

Both altimeters will be

tested before being

integrated into the vehicle.

2.6 Motor ejection is not a

permissible form of primary

or secondary deployment.

An electronic form of

deployment must be used for

deployment purposes.

There are no plans of ejecting

the motor out of the launch

vehicle, nor using it as means

of primary or secondary

deployment. The flight

computers will be used to send

out signals for parachute

deployment.

The team will continue

developing the current

means of recovery, without

considering motor ejection

as a means of deployment.

2.7 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 flight computers present

in the launch vehicle will have

independent arming switches

accessible from the exterior of

the airframe.

The team will clearly mark

the arming switch for each

flight computer on the

exterior of the launch

vehicle airframe.


2.8 Each altimeter shall have a

dedicated power supply.

All three flight computers

present in the electronics bay

are equipped with a 9 volt

battery power source.

During manufacturing, the

team will make sure the

each flight computer is

wired correctly to its

respective 9 volt battery.

2.9 Each arming switch shall be

capable of being locked in

the ON position for launch.

The arming switch for each

altimeter is being designed to

be turned on and off by using a

simple lock and key

mechanism.

The team will integrate each

flight computer circuitry to

an independent arming

switch. Each individual key

will be dually marked for

ease of recognition.

2.10 Removable shear pins shall

be used for both the main

parachute compartment and

the drogue parachute

compartment.

2.11 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.

A GPS unit will be placed in

both the electronics bay of the

main airframe and in the

tethered nosecone.

The team will make sure

both GPS units are

integrated securely in each

independent section

separate from any other

components.

2.12 The recovery system

electronics shall not be

adversely affected by any

other on-board electronic

devices during flight (from

launch until landing).

The launch vehicle will be

designed so that each flight

computer will experience no

disruption to its individual

performance.

The team will be

conscientious of keeping the

electronics of the launch

vehicle safe from

disturbance.

Table 16: Recovery S ys tem Verificati on


III. AirBrake Subsystem

Figure 52: AirBrak e Bay

Figure 53: AirBrak e S ys tem Operational Diagram


SUBSYSTEM CHARACTERISTICS

The AirBrake section is 23.77 inches long by 6 inch diameter airframe.

Figure 54: S ys tem Components

Component Function

1 – Bulkhead Separates the airbrake system

2 – Bearings Allows the Screw to rotate with minimal

friction

3 – Flap link Transfers motion to flaps

4 – Screw nut Rides on screw allowing for transfer of

rotational motion to liner

5 – Screw Extends rotation from motor allows nut to

move on it.

6 – Carbon fiber Flap Provides drag force at different angles

7 – Flap hinge Allows flap to rotate open

8 – Airframe Main structure

9 – Stepper Motor Allows precise control of flaps Table 17: AirBrak e Components and Function


Flap Motion

Figure 55: AirBrak e Clos ed

Figure 56: Airbrak e Open


SYSTEM COMPONENTS

Bulkhead

The bulkhead closest to the engine bay will be made of G – 10 fiberglass in order to be

able to handle the thrust rating of our chosen engine. The bulkhead on which the stepper motor

will be mounted will be either plywood or G – 10 fiberglass, depending on more in depth load

calculations on the flaps.

Ball Screw

Figure 57: Grade 8 S teel Ful ly Threaded Rod

Both steel and aluminum were considered as material choice for the ball screw, but the

steel screw was chosen over an aluminum one because of steel’s high strength, commercial

availability, and inexpensive cost. This screw will be cut and machined to our specific needs.

Force Calculations

Calculated theoretical critical velocities are carried out for the case when both ends of a

screw are fixed into bearings, as it is in our specific application. However, the team’s research

has found that maximum velocity should be less than 80% of this calculated value.

Critical speed refers to the RPM at which the natural frequency of a rotating shaft will

occur. This vibration, also called resonance, will occur regardless of the orientation of the

leadscrew. The critical speed also applies to rotating nuts about a lead screw.

𝐶𝑠 =𝑁 × 4.76 × 106 × 𝑑

𝐿2


d = Minor Diameter of Screw

L = Distance between nut bearing

N = Fixity type

= 0.25 for fixed-free

= 1.00 for supported-supported

= 2.00 for fixed-supported

= 4.00 for fixed-fixed

Critical speed (RPM) 18593.75

Root diameter of screw (in) 0.25

End fixity factor 1

Length (in) 8

Table 18: S crew Critical S peed calculations

As can be seen from the table above, the maximum stable RPM of this screw is over

18,000 RPM. For the given application, the screw is never expected to go above this value.

Column Strength Calculation

Column strength refers to the maximum load a lead screw can withstand in compression

before failing. Compression loads that exceed the column strength will cause the lead screw to

buckle, or bend. Even the slightest bend can ruin a leadscrew, and threaten the AirBrake system.

The following equation was used to calculate maximum column strength:

𝐶𝑙 =𝑁 × 14.03 × 106 × 𝑑4

𝐿2

After calculation, it was determined that the column strength in the screw is 857 lbs.

Power Calculations

Calculating minimum power output to translate the load provides a starting point for

specifying the rest of the system's components. The equation which follows was used to find the

minimum power output.

𝑃 = 𝐹 × 𝑆

𝑡

where F is the load, S is distance traveled, and t is required time to get there.

Figure 58: End Fixi ty Factor


Calculating minimum power output

Time 4

Distance (in) 5

Power (W) 43.78162 Table 19: Power Calculations

Peak Velocity Calculations

The calculated peak linear velocity for this application was calculated using the following

equation:

𝑣𝑝𝑘 = 3𝑆

2𝑡

where vpk = peak linear velocity in in/s, S is distance traveled, and t is required time to get there.

peak linear velocity (in/s) 1.875

Table 20: Peak l inear veloci ty

This shows the peak linear velocity for the given power output would be 1.875 in/s. This

peak velocity will allow the flaps to be fully opened in 2.5 seconds. The team is aiming for an

ideal velocity of 1.25 in/s, which allow the flaps to fully open or close in 4 seconds.

Minimum Required Pitch

The following equation is used to calculate the minimum pitch needed to keep the

leadscrew speed at approximately 2,000 rpm. The 2,000 RMP value is based on the selected

motor, which is discussed further below.

𝑃𝑚𝑖𝑛 =𝑣𝑝𝑘 × 60

𝑀𝑎𝑥𝑖𝑚𝑢𝑚 𝑠𝑐𝑟𝑒𝑤 𝑅𝑃𝑀

where 𝑃𝑚𝑖𝑛 is minimum pitch needed and vpk is the peak linear velocity in in/sec.

Minimum pitch (in.) 0.05625

Maximum Screw rpm 2000 Table 21: Pi tch Calculation

The minimum distance that the screw would need to advance in one revolution is 0.05625

inches. Using these calculations, the team will be able to soundly choose the proper screw for the

needed requirements.

Torque Calculations

The torque required to lift or lower a load can be calculated by equation shown below.

𝑇 = 𝐹𝑇 ×𝐿

2𝜋𝑒


where 𝐹𝑇 is Total force (lbs.), L= Lead (distance that a screw, or nut, advances in one

revolution), e= efficiency (no units, use 0.9 for ball screws assemblies). The maximum amount

of load the flaps will ever experience is 310 lbs; therefore, this was used as the selected load for

the torque calculations.

Force Input 310 lbs Pitch 0.05625 in

Torque Needed 54.17897 oz*in Table 22: Torque from inputted force

If the torque of the motor is known, the equation above can be rearranged to solve for

maximum lifting load as follows:

𝐹𝑇 = 𝑇𝑚 ×2𝜋𝑒

𝐿

where 𝑇𝑚 is motor torque.

Input Torque 75 0.529616

Pitch 0.05625 1.42875

Output Force 1908.88 429.1333 Table 23: Force from inputted torque

The output force of 429 pounds is 119 pounds greater than the worst case scenario,

allowing for there to be a sufficiently large amount of play area in the case of any emergency. As

a reference, the above force and torque calculations are for a worst case scenario if the Airbrake

system were to deploy at a maximum velocity. Under nominal operating conditions, the system

is never expected to experience this type of load.

Lead Screw FEA Analysis

In order to have a successful AirBrake system, the lead screw must be able to withstand

the drag induced by the flaps and the load from the thrust of launch vehicle engine. To check that

the screw would be able to withstand these, FEA simulations were run. The thread of the screw

was ignored to save on computation time.


Figure 59: FEA of bas ic s teel lead s crew

FEA analysis of our lead screw under a worst case a scenario load resulted in a maximum

stress of 8.615 ksi. The yield strength for this alloy of steel is 51 ksi, giving the lead screw a

factor of safety value of 5.7.

FEA analysis was also run for an aluminum lead screw for comparison purposes in the

following figure.

Figure 60: FEA of Aluminum lead s crew


The aluminum lead screw also did well under a worst case a scenario load. The maximum

stress was 9.265 ksi, giving a factor of safety of 4.3. An aluminum screw was considered due to

being lighter than its steel counterpart.

Stepper Motor

A high torque stepper motor will rotate the lead screw, and was chosen due its ability to

rotate in precise angles. This selection will also allow for keeping track of position, which would

assist the AirBrake computer in knowing the location of the flaps at all instants in flight.

Figure 61: S elected S tepper Motor

This motor was selected dude to its high torque, compact size, low weight, and low

inertia, which can be seen in the following specifications table.

Table 24: S tepper Motor S peci fications


Figure 62: S tepper Motor Dimens ions

Figure 63: Torque vs . RPM

Depending on the development of the final AirBrake design, the stepper motor selection

may change with calculated torque and RPM needs. A possible design alternative is a stepper

motor with a built – in lead screw, seem in the image below.


Figure 64: S tepper Motor Des ign Alternative

Bearings

Figure 65: Left - S BACB606DD-20 1 /4 I.D bearing hous ing . Right - 6383K214 1 /4 I.D bearing

Bearings where chosen due to being inexpensive and easy to replace. In addition, they

allow low fiction rotation of the drive screw. The 6383K214 bearing has a maximum dynamic

load capacity of 356 lbs, which surpasses the amount of load this subsystem is ever expected to

encounter. The majority of the loading will be in the axial direction, allowing the transfer of

loads to the housing of the bearing.

Bearing housing for the SBACB606DD-20 is 2017 aluminum alloy. This was chosen

over steel for the purpose of saving weight while still maintaining structural integrity. 2017

aluminum also has good machinability and mechanical properties. Another selection for bearings

could be manufactured Teflon bushings, would allow for greater weight savings but increase

rotational friction.


Figure 66: Teflon Bus hings

Flap Linkages

Figure 67: AirBrak e Flap Link age

The linkages between the lead nut and the AirBrake flaps will be an aluminum link. This

link allows the transfer of linear motion to the rotational motion causing the flaps to open. Link

size was determined by computer – aided design, which allowed for optimization of the

maximum flap opening angle.

Aluminum was selected because of its low density when compared to steel. Also, it is

easy machinable, allowing for flexibility when deciding on the final design of the linkages. As

the design of this subsystem is developed, the aluminum link geometry will be more concretely

designed.


Flap Mounting and Geometry

Figure 68: Flap hinge mounting

AirBrake flaps will be mounted on one end by hinges, which allow the flaps to rotate

through their range of motion easily. The hinges to be used will be standard door/cabinet hinges,

which was decided upon in order to reduce cost and allow for the hinges to be easily replaceable

in case of any unforeseen emergency. Hinges will be secured onto the bulkhead and the flaps

using standard fastening hardware.


Figure 69: AirBrak e S ys tem Flaps

The AirBrake flaps will open onto the free stream of air, inducing drag on the launch

vehicle. This action will slow down the vehicle, allowing it to accurately reach the target altitude.

The flaps will be cylindrically shaped, in order to minimize drag when in the closed position.

The flaps will be constructed much like the construction of the airframe, with a section of

phenolic tubing wrapped in carbon fiber. This method strengthens the flaps to be able to

withstand the expected loads.

Figure 70: Airbrak e Flap s ection cut

The AirBrake flap will be attached with a pin connection, which will allow for removal

of the flap to either perform repairs or access the internal components of the bay. The following

figure demonstrates the mass properties of the flap design.


Figure 71: AirBrak e Flap mas s properties

AERODYNAMIC ANALYSIS

The key factor of this system is having the rocket be aware of its current trajectory

through a serious of recurring in – flight calculations. If the estimated apogee height of the

rocket, at any given time, is greater than the targeted altitude; the flaps will deploy themselves.

This will induce an extra drag force on the rocket, consequently slowing down its ascent and

decreasing its apogee height. If the system calculated that the rocket is on an optimal trajectory,

the flaps will stay in place and not deploy. The table below gives a general overview of the

calculations the AirBrake system will need to make to achieve its purpose.


Table 25: Propos ed Firs t Iteration Drag Force Calculation

The flight computer will gather the necessary data, such as the launch vehicle altitude,

velocity, orientation, acceleration, and location. Using this acquired data, a second computer will

process the data and take into account aerodynamic forces to simulate the rocket’s projected

altitude and trajectory.

If the highest point on the calculated trajectory is over one mile in altitude, the computer

will calculate the amount of drag force needed to hit that target altitude and deploy flaps as

necessary. As vehicle velocity decreases more drag force will be needed; therefore, the flaps will

be extended to full deployment. This will provide maximum amount of drag force.

Since the flap computer will start to process data after engine cut off we may safely

assume that there is no thrust acting on the launch vehicle. Thus we can start with:

𝐹 = −𝑑𝑟𝑎𝑔 − 𝑤𝑒𝑖𝑔ℎ𝑡

Where drag is:

And weight is


𝑊 = 𝑚𝑔 sin 𝜃

Making the appropriate substitutions will leave this derived equation for acceleration:

𝑑𝑣

𝑑𝑡=

− (12 𝐶𝑑𝜌𝐴𝑣 2) − 𝑚𝑔 sin 𝜃

𝑚

The area (A) is a based on the positon of the AirBrake flaps at the time, where 0 degrees

would be only the cross – sectional area of the rocket, and 90 degrees would be cross – sectional

area of the rocket plus the area of the flaps. The velocity would be inputted real – time from our

flight computer, which would also calculate the coefficient of drag as it varies with the positon of

the flaps.

Computational Fluid Dynamics (CFD) Analysis

The preliminary aerodynamics were teste through the use of flow simulation software.

The launch vehicle was simulated with the air speed after engine cut – off of 765 ft/s, obtained

from Rocksim simulations. Drag force and moments were the main parameters being observed.

For all the following figures, the “Z” axis is taken to be axial with the launch vehicle.

Figure 72: Flap Dimens ions

The AirBrake flap will be relatively large due to the increased ability it would provide to

generate drag at slower velocities. As the launch vehicle approaches apogee, its velocity is

greatly decreased when compared to the velocity at engine cut – off.


Figure 73: CFD res ults with AirBrak e clos ed

As expected, the pressure at the nosecone is higher than at the rear. This can be a credited

to the Von Kármán nosecone purpose of minimizing drag.

Figure 74: Nos econe tip s tagnation pres s ure

Above, you can see as flow reaches a stagnation point, the fluid velocity reaches zero and

all kinetic energy has been converted into pressure and temperature isentropically. For this

reason our Von Kármán nosecone will have an aluminum tip to withstand increased heat and

pressure of a stagnation point.


Table 26: CFD Drag (Z) res ults with AirBrak e clos ed

The drag value of 245 N is equivalent to 55 lbs of drag at maximum launch vehicle

velocity. This is only marginally different from the value of 58 lbs obtained when calculating the

drag force in the spreadsheet seen in Table 17.

Figure 75: Five Degree angle flap deployment

Figure 76: Clos e-up of pres s ure dis tribution around flap


After engine cut – off the launch vehicle will be at its maximum flight velocity. During

this time, small angle increments will have large impact on drag force. An increase of five

degrees in flap deployment angle results in 62.5 lbf of drag force, which is an increase of 7 lbf

when compared to no flap deployment.

Table 27: CFD Drag (Z) res ults with AirBrak e at five degrees

The following two images show the pressure distribution and drag at fifteen degrees of flap

deployment.

Figure 77: Pres s ure Clos e -up of 15 degree flap deployment

Table 28: CFD Drag (Z) res ults with AirBrak e at 15 deg rees

The drag force being provided by the flaps at fifteen degrees of deployment is about 33

lbf.


Figure 78: Airflow over flap at 15 degree deployment

Figure 79: Pres s ure Dis tribution over Flap

Figure 49 demonstrates airflow over the AirBrake flap with a fifteen degree deployment,

where the orange and red colors highlight areas of high local air speeds. Figure 50 shows the

total pressure distribution over the AirBrake flap. Orange and red signal a higher pressure than

the surrounding green area.


Worst Case Scenario

The worst case scenario the AirBrake will ever see is if the flaps were to fully deploy

right after engine cut – off, when the surrounding air speed is 765 ft/s. This scenario would only

occur if a computer malfunction causes the flaps to fully deploy immediately after engine cut –

off.

Figure 80: Pres s ure dis tribution at engine cut-off

Table 29: CFD Drag (Z) res ults at ful l deployment

As can be seen from the images above, 282.8 lbs of drag are provided by the AirBrake in

this scenario. In accordance with this, the AirBrake subsystem will be designed to be able to

withstand this maximum load. The images below demonstrate that the spreadsheet calculations

also confirm this maximum drag force.


Table 30: S preads heet calculation at ful l deployment

CODE DEVELOPMENT

The design of the air-braking system’s code can be divided into several steps and sub-steps.

These steps are listed below.

1. Initial Altitude Prediction

2. Onboard Altitude Prediction

3. Error Correction

4. Integration

These steps are not necessarily the order in which the system was designed. Nonetheless,

they provide a good structure for understanding the way the system works.

Initial Altitude Prediction

As stated previously, for this system to work correctly the vehicle needs to overshoot the

given one mile apogee height under any circumstances. To accurately predict the altitude in

varying conditions the team used the program Rocksim, which is known to be reliable and highly

accurate. Throughout the entire code design process Rocksim will be used as our control to test

different methods of altitude prediction against before any actual launch tests occur. Rocksim


allows the user to change many variables outside of rocket configuration including

environmental variables like wind speed, thermal occurances, humidity, and many others. For

this reason, we are assuming our Rocksim predictions to be almost completely accurate.

To determine what motor we needed we ran simulations in worst case scenarios to find

what total impulse/thrust we would need to ensure we at least reach 1 mile. This includes excess

launch angle, launching with the wind, and max allowable wind speed (25 mph). From this we

were able to determine a max altitude and therefore determine a suitable size for the brakes, as is

explained further in the sections prior.

Onboard Coefficient of Drag Prediction

Once the vehicle is launched and the motor is finished firing, the AirBrake software will

begin iterations. The purpose of the software is to actuate the flaps to induce drag. In order to

achieve this, the software will iteratively predict what drag is required to attain an apogee of one

mile. Ideally, we would use the highly accurate methods which programs like Rocksim use to

determine maximum altitude. However, this code, instead of predicting max altitude, is assigning

maximum altitude as a static variable of one mile and solving for the drag required to reach that

maximum altitude. Because Rocksim’s methods are numerical in nature, they cannot be solved

backwards and are not suitable for our application. Instead we will begin by using the kinematic

equation for projectile motion with air resistance, shown below.

This is the original equation where given values of mass, gravity, velocity, air density,

cross sectional area, and coefficient of drag; the apogee height can be determined. Solving for k

in the equation above, the equation below is obtained. This equation can be easily solved for

Cd*A (coefficient of drag times cross-sectional area) and will be the primary equation in the

code.


While these equations are fairly accurate, they would have a tendency to predict a higher

apogee than reality and therefore cause the brakes to over-extend and the rocket to undershoot.

To combat this, we will be including error correction methods which will be discussed in a later

section.

Determination of Coefficient of Drag

Once the rocket is in flight and the computer has predicted the value for Cd*A required to

reach one mile, the computer will then open the flaps until they match that value. To determine

how much the flaps need to be opened, information needs to be obtained on values of Cd*A for

varying flap positions. For this we decided to run many flow simulations on the launch vehicle

for flap positions starting at 0 degrees all the way to maximum. From this data, and the values of

air density and launch vehicle velocity, we can obtain accurate values for Cd*A at all possible

flap positions.

An additional consideration to take into account during this data acquisition process is the

effect of velocity on the Cd*A value. We will test the same flap position at several different

velocities to determine if a constant value of Cd*A can be used or if another dimension of data

needs to be included in the code.

Once all the data is obtained it can be organized in two different ways. The first would be

to include it in an array within the code. Depending on the rocket’s current state, the correct

value can be found in a table form. The other way would be to create a best fit curve for the date

and obtain the correct corresponding flap position using an equation. Both of these methods will

be evaluated to determine which is more efficient while the code is running.


Error Correction

Assumption of Projectile Motion with Drag

The assumption of projectile motion with drag allows us to drastically simplify our code

and ensure quick performance. The kinematic equations can be separated into component form

and the horizontal component can be neglected. While this makes coding easy, it is provides a

large source of error in accurately predicting the correct coefficient of drag value. Not only does

it neglect atmospheric variances, but also the effect of the relationship between the center of

pressure and center of gravity on the rocket. Having a Cg located forward of the Cp, increases

stability but introduces a moment on the rocket which will cause the actual apogee of the rocket

to be lower than the one predicted.

Our solution to correct this discrepancy is to compare the differences between apogees

predicted using both projectile motion with drag, and Rocksim, given the same initial starting

conditions. The variables to be tested are initial velocity, and initial tilt. Every combination of

those two variables will be tested and the variances in results recorded. Since we are assuming

Rocksim to be accurate, we will then use the calculated differences to introduce correction

factors into the code to force our predicted apogee using projectile motion with drag to match

that of Rocksim. There is still some debate over where the correction factor should be applied

and there are currently two options. The first is to apply it to the Cd*A value to produce an

“effective Cd*A.” This method will apply the correction factor after it has been calculated in the

code. The other option is to apply it to the fixed one mile altitude, which would in essence

produce an “effective altitude.” Both of these options will be tested to determine which is more

accurate and efficient.

The Density Problem

The value of air density to be used during flight will be directly obtained from the flight

computer in real – time. This is then used to determine a constant flap position that will

theoretically be maintained until the target apogee is achieved. In reality, density changes with

altitude causing another source of error. To solve this, an average value of air density will be

calculated and used in the code. This can be as simple as using the mean value between the

current density and the density at one mile. Analysis will be done to determine if this would be a

good assumption or not.


Possible real-time error correction using flight data

Another method that is currently being discussed is real-time error correction based on

actual flight data. In this method, the computer will not only predict the value of Cd*A needed to

reach apogee but also predict its own altitude over a small time frame. Once the rocket actually

traveled for the given time, its actual altitude can be compared to the previously predicted

altitude and a correction factor can by dynamically generated and applied to future iterations of

the code.

This idea has some potential issues however. Firstly, it is a lot more for the computer to

calculate; and during the time period where data is gathered, the flaps must stay constant.

Depending on the length of the time period, this could potentially cause the whole system to

react too slowly and not work properly. Other issues are that this method may just introduce even

more error into the system than it removes due to imprecise data sent by the flight computer.

Transient Flap Actuation

Another source of error to be addressed is the dynamic movement of the flaps between

fixed positions. This dynamic movement will be taken into account by calculating an average

interim Cd*A value based on the difference between the starting angle and target angle, and

applying it over the time it takes for the flap to actuate.

Global error correction variable

Once the rocket is constructed and the code is ready, a flight test will be performed.

Based on the variation between the actual achieved apogee and the target of one mile, a final

correction variable can be implemented to calibrate the entire system to be more accurate. This is

based on the assumption that we may end up with a precise system that always tends to

overshoot or undershoot. This global error correction should fix that by altering the target apogee

slightly depending on the system’s performance. Implementing this type of correction would

require multiple test flight, being overly costly.

Integration

Testing will be conducted to ensure quick and reliable data transfer from the Telemega

flight computer to the Beaglebone AirBrake controller.


Pseudo – Code

The basic sequence of steps the computer must iteratively perform is listed below.

1. Gather current velocity, tilt angle, and altitude from Telemega.

2. Determine y component of Velocity

3. Determine current air density and create an average value based on current and target

altitude

4. Calculate constant Cd*A needed to achieve apogee

5. Apply correction factor to Value of Cd*A

6. Determine brake angle that matches the given value of Cd*A, current velocity, and

current tilt angle.

7. Actuate brake to calculated angle.

a. If brake is already moving to angle, just update target angle.

b. If brake is moving away stop the previous action and go to new angle.

Test Code

Below is a sample of code written in MATLAB that contains the equation for calculating

the value of Cd*A necessary to reach one mile apogee. The results check is to ensure that the

original function used to solve for ‘k is still accurate. In solving for ‘k’, a new function is

introduced called the lambert W function, or product log function. In MATLAB so far, this

function takes some amount of time to solve which may be problematic in the future onboard

code.

% Altitude Targeting using Projectile Motion with Drag

clc

clear all

%Input Variables

A = 0.19635;

y = 8000;

v = 800;

m = .93243;

g = 32.18504;

p = 0.002134;

%Function Split into parts cuz matlab is annoying.


f1 = (-m*(v^2)*lambertw(-1,(-2*(exp((-2*g*y)/(v^2)))*g*y)/(v^2)));

f2 = 2*g*m*y;

f3 = 2*(v^2)*y;

k = (f1-f2)/f3;

Cd = (k*2)/(p*A);

disp ('Necessary Cd To Achieve 1 mile at current conditions');

disp (Cd);

%Results Check

CdT = 0.1168;

kt = 0.5*p*CdT*A;

yt = (m/(2*kt))*log(((m*g)+(kt*v^2))/(m*g));

disp ('Max Altitude');

disp (yt);

Below is an excel table used to predict apogee with varying methods. Listed so far is

apogee using simple projectile motion with no drag, projectile motion with drag, and also apogee

if the current measured drag were constant. This chart will eventually be used in the

determination of error correction factors when comparisons to Rocksim are made.

Input Variables

Variable SI English

Coefficient of Drag Cd 0.32 Cd 0.4

Rocket diameter d (mm) 98 d (in) 6

Cross Section Area A (m^2) 0.007543 A (ft^2) 0.19635

Weight W (lb) 30

Mass m (kg) 4.535931 m (slug) 0.93243

Velocity V (m/s) 30 V (ft/s) 800

Angle (deg) 0 0

gravity g (m/s^2) 9.81 g (ft/s^2) 32.18504

Cd*A (m^2) 0.002414 ft^2 0.07854

Air Density (kg/m^3) 1.1 (slug/ft^3) 0.002134

k k (kg/m) 0.001328 k (slug/ft) 8.38E-05

Current drag 1.194805 53.64207

Current drag acceleration 0.263409 57.52932


Max Altitude Ymax (m) 45.26652 Ymax (ft) 5702.159

Max Alt w/o Drag Ymax 45.87156 Ymax 9942.508

Max Alt w/ Current Drag Ymax 44.67207 Ymax 3566.876 Table 31: AirBrak e Computer Inputs

IV. Mission Performance Predictions MISSION PERFORMANCE CRITERIA

In order for the mission to be deemed successful, the following criteria must be met:

1. The launch vehicle must reach an apogee height of 5,280 feet above ground level

2. The launch vehicle will be designed to be completely recoverable and reusable.

3. Each independent section of the launch vehicle should land with a maximum kinetic

energy of no more than 75 ft – lbf.

4. Electrical circuitry of the recovery system will be completely independent of any payload

bay circuitry.

5. The recovery system will be equipped with two commercially available altimeters for

redundancy.

6. Any recovery system components must be deployed electronically; motor ejection may

not be used as a form of deployment.

7. An arming switch will arm each altimeter, and be capable to be left on the “on” position

for launch.

8. Each altimeter will have a dedicated power supply.

9. Parachute bays will use removable shear pins.

10. Each independently – landing, untethered segment of the rocket will have an electronic

tracking device.

LAUNCH VEHICLE CHARACTERISTICS SUMMARY

The launch trajectory date of this year’s competition vehicle was simulated using a

Rocksim model displaying the entire configuration plan of the launch vehicle. This model can be

seen in Figure 71 displayed below. The use of this program allowed different overall

characteristics of the rocket to be determined. These values can be found in Table 31 below.

Overall Length (in.) 129.07

Overall Diameter (in.) 6.155

Overall Mass (lb.) 34.67

Stability Margin 2.36


MOTOR SELECTION

For this year’s competition, Cesaroni motors were chosen because they are known for

being both reliable and simple to use. The change to the Cesaroni 4263L1350 – P, from the

initially proposed Cesaroni 3300L3200 – P, was needed due to the addition of thirty – five inches

to the overall length of the launch vehicle. This motor choice brought the projected apogee

height to just over six thousand feet. This overshoot in height allows for design and

manufacturing flexibility, as well as allows sufficient time for the computer of the air – braking

system to converge on the one mile apogee height and deploy flaps appropriately.

The following images demonstrate two different thrust curves, one given by the

manufacturer and one simulated from the flight data provided by Rocksim.

CG location (from

nosecone tip)

88.9

CP location (from

nosecone tip)

103.04

Table 32: Overal l Rock s im Vehicle Values

Figure 81: Rock s im S chematic of Launch Vehicle


Figure 82: S uppl ier-Provided Thrus t Curve


Figure 83: S imulated Thrus t Curve

As can be seen from the above demonstrated plots, the simulated curve matches up very

closely with the thrust curve provided by the manufacturer. The table below outlines the basic

specifications for the motor chosen from the flight data simulated from Rocksim.

Thrust – to – Weight Ratio 10.93

Rail Exit Velocity 80.39 ft/s

Projected Altitude 6068.31 ft.

Maximum Acceleration 36518.2 ft/s2

Motor Burn Time 3.284 s

Maximum Motor Thrust 346.37 lbf

Average Motor Thrust 293.00 lbf

Total Motor Impulse 962.22 lbf

Table 33: Motor S imulated S peci fications


STABILITY ANALYSIS

In order to validate the choice of motor for this competition, Rocksim was used to find

the stability margin of the launch vehicle. In addition to finding the overall stability margin for

the launch vehicle, Rocksim was used to generate graphs analyzing the stability margin as a

function of time throughout the flight of the launch vehicle. This graph, along with a graph of

center of gravity/center of pressure versus time, may be found in the accompanying figure below.

(The graphs below are for the analysis done with no external wind speed; the graphs for other

analyzed wind speeds may be found in Appendix A.)

Figure 84: CG/CP vs . Time, no wind


Figure 85: S tatic Margin vs . Time, no wind

From the above graphs, it can be seen that the static margin of stability always lies in an

acceptable range between launch and apogee height. Therefore, the chosen motor can be deemed

effective for the chosen application.

FLIGHT SIMULATIONS

The flight profile for the launch vehicle, along with any relevant flight data was obtained

by launching the vehicle in Rocksim with the Cesaroni 4263L1350 – P motor chosen for this

application. The following image demonstrates the flight profile of the launch vehicle for five

simulated cases: no wind, 5 – mph wind, 10 – mph wind, 15 – mph wind, and 20 – mph wind.


Figure 86: No Wind Fl ight Profi le

Figure 87: 5 -mph Wind Fl ight Profi le


Figure 88: 10 -mph Wind Fl ight Trajectory


Figure 89: 15 -mph Wind Fl ight Profi le

Figure 90: 20 -mph Fl ight Profi le


From these flight simulations, Rocksim provides data for apogee height, in addition to

maximum drift given certain launch criteria. The table which follows outlines the apogee height

and maximum drift distance for all five launch conditions simulated by the above flight profiles,

as well as other useful data such as maximum velocity and acceleration.

Initial Wind Speed Apogee Height (ft.) Drift Distance (ft). Max. Velocity

(ft/s)

0 mph 6066.5 258.16 762.7

5 mph 6033.4 1078.9 762.6

10 mph 5991.5 1516.9 762.3

15 mph 5813.9 2399.8 760.6

20 mph 5883.0 2164.3 761.3 Table 34: Fl ight S imulation Data

In addition to this data, Rocksim also provides means of analyzing the graphs of different

parameters with respect to time. The following figures demonstrate the plots of the parameters

mentioned above with respect to time for the case with no wind. The graphical data for the

remaining cases can be found in Appendix A.

Figure 91: Alti tude vs . Time, no wind


Figure 92: Range (Drift) vs . Time, no wind

Figure 93: Mach Number vs . Time, no wind

The above graph for Mach number is found by diving the velocity graph by the speed of

sound in air. The relation between Mach number versus time, and velocity versus time, is the

same due the assumption that the speed of sound in the given application is relatively constant.


LANDING ANALYSIS

According to the rules of this competition, a safe landing constitutes a kinetic energy at

landing of no more than 75 ft – lbf. The choice of parachute found in the recovery subsystem

portion of this report was done only after the kinetic energy at landing was found to be less than

what the competition specified. The launch vehicle recovery system was designed to allow the

vehicle to land in one segment, therefore only one kinetic energy calculation per parachute

option had to be taken into account.

The terminal velocity of launch vehicle was calculated using

𝑉 = √2𝐸𝑔𝑐

𝑚

where E is the kinetic energy, gc is the dimensional constant, and m is the total mass of the

launch vehicle to be recovered. A value of 75 ft – lbs was used for the maximum kinetic energy,

since this was the requirement established in the statement of work to determine the minimum

size of the parachute. The steady state velocity under parachute was calculated using

𝑉 = √2𝑚𝑔

𝜌𝐶𝑑𝐴

where g is acceleration due to gravity, ρ is the density of air, Cd is the coefficient of drag of the

parachute, and A is the effective area of the parachute.

The following table outlines the method and values used to determine that the final

parachute selection was within the guidelines of the competition:


Table 35: Des cent Rate Energy Calculation

From the kinetic energy calculations, it was found that the launch vehicle will have to

descend at a rate of 12.69 ft/s on only one parachute in order to have an energy of 75 ft – lb.

With the Iris Ultra, our launch vehicle will have a decent rate of 67.97 ft – lb, which is safely

beneath the maximum value.

V. Interfaces and Integration

INTERNAL VEHICLE INTERFACES

Nylon shear pins will be used to hold the nosecone to the upper airframe during ascent.

At apogee, the nylon pins will be sheared by the ejection event. This allows for the parachute to

be ejected. The following image outlines the above configuration.


Figure 94: Nos econe Interface

The parachute bay airframe will be secured with the AirBrake bay airframe with four

steel pins. The pins will go through the parachute phenolic airframe and be secured onto the rails

on which the electronics bay will slide. The rails will be secured to the walls of the AirBrake

Airframe with high strength epoxy. An L – bracket will be used to mount the rails to the

AirBrake bulkhead. This allows any load to be diverted from the electronics bay to the adjacent

bulkhead. The image immediately following demonstrates this interface configuration.


Figure 95: Parachute Bay to AirBrak e Bay Interface

To secure the AirBrake bay to the propulsion bay, and aluminum ring will be

manufactured by being milled. This aluminum ring will act as a center ring joining the two

airframes. The airframes will be joined on the aluminum ring, and secured using set screws on

both the AirBrake airframe and the propulsion bay airframe. This can be seen in the following

image.

Figure 96: AirBrak e Bay to Propuls ion Bay Interface


EXTERNAL VEHICLE INTERGRATION

The launch vehicle will need to interface with both the payload – retrieving rover and the

ground support equipment in order to ensure a smooth and fully – autonomous process. To

organize the process, the following flowchart was created. This allowed the entire team to be

adequately aware of the whole scope of the project, even when members are mostly focused on

one portion of the project.

As is seen in the flowchart image above, the sequence will begin by the allocated start

switch required for the competition. At this point, a wireless signal will be sent to the

RaspberryPi controlling the rover, initializing the webcam computer vision. The rover’s servos

will begin to actuate the motion and actively search for the payload. Once the payload has been

Sequence Initialized by Start Switch

WiFi Signal Sent to Rover RaspberryPi

Computer Vision on Rover Begins Motion to Search for Payload

Rover Motion Picks Payload with Clamping Arm

Rover Travels to Payload Bay in Vehicle

Payload Deposited in Bay Clamps

Rover Moves Under Its Ramp

Radio Signal Sent to Servo Controller To Close Payload Bay

After Set Time, RaspberryPi Uses Radio Signals to Interface with AGSE Beaglebone

AGSE Sequence Initiated

Vehicle Erected, Ignitor Inserted

Vehicle Launched

Figure 97: AGS E S equence Flowchart


located, the rover will pick it up with its clamped arm, and transport it to the launch vehicle

payload bay. After placing it in the clamps, the rover will transport itself underneath the ramp it

uses to travel to the launch vehicle, in order to shield it from any launch debris.

After the rover is safely put away, the RaspberryPi controlling it will send a radio signal

to the servo controller, allowing for the Payload Bay to close. A set time for this action will be

recorded prior; after this time has passed, the RaspberryPi will communicate with the

Beaglebone controlling the ground support. This will begin the launch vehicle erection and

ignitor insertion. After this entire process is complete, the vehicle will be ready to be launched.

VI. Safety

SAFETY OFFICER

Juan Trujillo is the safety officer for the FIU PantherWorks Space team during the 2015 –

2016 season. He is responsible for taking the steps necessary to ensure the overall safety of the

team members as well as of the public throughout all team activities. He is also responsible for

ensuring compliance with all rules and regulations.

The responsibilities of the Safety Officer include the following:

Monitor team activities with an emphasis on Safety during:

A. Design and construction of vehicle and AGSE

B. Ground testing of vehicle and AGSE

C. Sub-scale launch test

D. Full-scale launch test

E. Competition launch

F. Recovery activities

G. Educational Engagement activities

Lead the team in establishing safety procedures for construction, assembly, launch, and

recovery activities and enforce those procedures

Identify any safety concerns and take appropriate action to mitigate the hazard.

Maintain current revisions of the team’s hazard analyses, failure modes analyses,

procedures, and MSDS/chemical inventory data.

Plan for proper purchase, storing, transporting, and use of all energetic devices.

Ensure compliance with all local, state, and federal laws, as well as NAR and TRA

regulations

Provide a written safety manual to the team that includes hazards, safety plans and

procedures, PPE requirements, MSDS sheets, operator manuals, FAA, NAR and TRA

rules and regulations, and confirm the all team members are complying with the rules set

forth in the safety manual


Brief the team before activities on a safety plan for different environments, materials

used, and tests. Enforce proper use of Protective Equipment during these activities

Establish a risk matrix that determines the risk level of each hazard based on likelihood

of occurrence and severity of the event.

PRELIMINARY CHECKLIST

Procedure Check if safely completed

The NAR/TRA mentor will check that

the flight computers are set to “OFF”

before loading the ejection charges.

The entire parachute loads will be

inspected prior to being loaded into the

rocket, and all flame proof materials will

be checked for wear and holes prior to

loading.

The knots on all shock cords will be

checked, as well as the firmness of their

fit onto the airframe and associated

components of the airframe.

Metal to shock cord interfaces will be

checked for rust and smoothness to

ensure that the cord cannot be torn by

the force of ejection.

The TRA mentor will then build and

load the rocket motor into the fully

assembled rocket, and will ensure that

the igniter cannot have voltage across its

terminals by twisting the ends of the

igniter together.

The TRA mentor and the team will then

fill out a flight card for the rocket, stating

its expected altitude and the certified

impulse of the motor.

The rocket will then be loaded onto the

AGSE, and upon completion of payload

insertion, and once the rocket has been

lifted to a proper launch angle, the

ignitor will be inserted by the AGSE into

the rocket engine.

The TRA mentor will check the igniter

for proper insertion. If the Igniter is

properly inserted, the TRA mentor will

complete wiring the igniter to the launch

control system present at the launch site,

and a team member will arm the flight


computers. Once the continuity beeps

from the flight computers are confirmed,

the team will move back to the minimum

safe distance for the launch. Table 36: Firs t Draft, Procedure Check l is t

SAFETY PROCEDURES

NAR Model Rocket Safety Code

Every team member is required to have read and acknowledged the NAR Model Rocket

Safety Code.

Pre – Launch Briefings

Additional briefings on the high power rocketry code for Level 2 flights will occur prior

to the launch day as well as before any preceding tests flight. A short quiz about safe distances

and the procedure for launching a rocket safely will be given prior to travelling out to the launch

site.

Team members will be given a basic instruction on the nature of launch sites, and will be

actively discouraged from engaging in cell phone texting and conversations while at the launch

site, unless they are far from the minimum safe distance required by NAR/TRA for the site.

All team members will be able to identify the potential hazards of assembling the launch

vehicle for launch and will remain focused and alert so that proper protocol is followed.

Caution statements will be issued with every plan of action for the construction of the

launch vehicle and AGSE. The safety officer will create a general Personal Protection Equipment

guideline for team members working on the rocket to consult before starting work each meeting.

The safety procedure will be inserted into every working document after the section that details

the powered equipment, chemicals, or materials to be used.

Rocket Motor Handling

The Chief Engineer will purchase the motor reloads from an online vendor or locally at a

launch site, and will store the reloads in a separate flame proof canister. The Chief Engineer will

load the rocket motor in accordance with TRA guidelines, and will be responsible for the rocket

motor in its entirety.

Ignitors for the rocket motors will come packaged with the reloads, and the ejection

charges will be filled by the Chief Engineer with small amounts of black power, according to the

engineer’s discretion on the amount of force required to ensure separation. The amount will not


exceed 5 grains of black powder. The black powder will be stored in a separate flameproof

canister from all of the other energetic ingredients, and will have a desiccant loaded into the

canister to aid in the removal and prevention of moisture contamination.

Only team members that have achieved a level 2 certification or higher are allowed to

acquire, store, and manipulate the team’s launch vehicle motors. By having obtained a Level 2

certification, the individual has demonstrated that he or she understands the safety guidelines

regarding motors. The motors shall be stored in accordance with the regulations set forth by

NFPA 1127. The motors for both test and competition launches will be transported by car to the

launch site.

HAZARD ANALYSIS

Risk Assessment Matrix

By researching and analyzing each human interaction with the environment and launch

vehicle system, and by reflecting upon past launching experience, hazards have been identified

and will continue to be brought to the team’s attention. Each hazard has been assigned a risk

level by evaluating the severity of the hazard and the probability that the hazard will occur using

the risk assessment matrix, found in Error! Reference source not found. X below. This type of

hazard analysis will continue to be done and updated as the project moves forward and more

potential hazards are brought to the team’s attention.

A severity value between 1 and 5 has been assigned to each hazard, with a value of 1

being the least severe. In order to determine the severity of each hazard, the outcome of the

mishap was compared to an established set of criteria based on the severity of personal injury,

environmental impact, and damage to the rocket, equipment, or personal property.

A probability value between 1 and 5 has been assigned to each hazard as well, with a

value of 1 being least likely to occur. The probability value was determined for each hazard

based on an estimated percentage chance that the mishap will occur given the following:

All personnel involved have undergone proper training on the equipment being used or

processes being performed.

All personnel have read and acknowledged that they have a clear understanding of all

rules and regulations set forth by the latest version of the safety manual.

Personal protective equipment is used as indicated by the safety lab manual and the

Material Safety Data Sheet (MSDS).

All procedures and safety precautions were correctly followed during construction of the

launch vehicle, testing, pre-launch preparations, and the launch.


All components were thoroughly inspected for damage or fatigue prior to any test or

launch.

In addition to the overall risk assessment matrix provided above, initial risk assessments

have been made for various possible hazards that have been identified at this stage in the project.

Acknowledging the hazards now brings attention to these particular failure mechanisms

that need to be improved upon. The team can take into account these possible failures as the

design continues to evolve and will work to diminish these hazards and to identify other possible

hazards throughout the design and building phases. The risk assessment charts that follow serve

to highlight areas that the team will need focus on moving forward with the project. The

identified hazards can be found below.

Lab and Machine Shop Risk Assessment

There are risks associated with working with machinery, tools, and chemicals that the

team will need to be aware of when manufacturing the launch vehicle. The following table

addresses these hazards. The manufacturing of parts for the launch vehicle and ground support

equipment will be performed mainly on the university campus.

1-4 = Low Risk Severity Value (SV)

4-8= Medium Risk 1-Insignificant 2-Minor 3-Moderate 4-Major 5- Catastrophic

10-25 = High

Risk

Could result in: ● Insignificant injuries, damage to property, or

environmental effects ● Monetary Loss <

$100

Could result in: ● Insignificant injuries or environmental

effects, ● Partial failure of non-critical system

● Monetary Loss >

$100

Could result in: ● Minor injuries ● Moderate environmental

effects ● Failure of non-critical systems

● Monetary Loss >

$500

Could result in: ● Severe injuries ● Reversible environmental

effects or ● Partial mission failure

● Monetary Loss >

$1,000

Could result in: ● Death ● Significant environmental

damage or ● Complete mission failure

● Monetary Loss >

$5,000

Likelihood

Value (LV) 1-Very Unlikely < 1% chance it will happen

1 2 3 4 5

2-Unlikey between 1% and

15% chance 2 4 6 8 10

3-Moderate between 15% and 50% chance

3 6 9 12 15

4-Likely between 50% and

90% chance 4 8 12 16 20

5-Very Likely > 90% chance it will happen

5 10 15 20 25

Table 37: Ris k As s es sment Matrix


Hazard Cause Outcome SV LV Risk Mitigation

Using power

tools such as

drills and

saws

Improper

training or

carelessness

when using

tools

Minor to

severe injuries

such as cuts or

burns

Damage to

component

being worked

on or

equipment

2 2 4

Low

risk

Safety gear appropriate

for each tools must be

worn when using it

Safety glasses must be

worn at all times when

in the machine shop

No untrained team

member may use

power tools unless

supervised by a trained

individual

No team member will

work alone in the

machine shop

Working

with

chemical

components

Chemical

Fumes or

splash

Minor burns or

injuries due to

skin contact or

inhalation of

chemicals

3 1 3

Low

Risk

Chemical containers

will be marked with

safety precautions

specification

MSDS documentation

will be available and

reviewed before use pf

chemical component

Appropriate protective

gear will be worn when

working with

chemicals

Work shall be

completed in a well

ventilated area Table 38: Lab/Machine S hop Ris k As s es sment

AGSE Functionality Risk Assessment

The hazards outlined in the table which follows are the risks linked to all components

comprising the AGSE; including launch pad functionality, rocket erector, and ignitor installation.

Due to the high importance of a stable launch assembly, the system will be rigorously tested

prior to any launches.


Unstable/

Un-level

Launch

Platform

Poor

Construction

Unstable

ground

May cause an

unpredictable

rocket

trajectory

2 2 4

Low

Risk

Extensive testing for

stability will be done

on the AGSE before

use


May impede

rocket launch

Launch Tower

could tip over

during launch

Ensure that everyone

is a safe distance

from Launchpad

established by NAR

rules

Rocket gets

caught on

launch rail

Misalignment

of launch

tower joints

High friction

on launch rail

Rocket may not

achieve a

sufficient

velocity before

exiting launch

rail, may be

damaged on

exit

3 2 6

Medium

Risk

Launch tower will be

inspected during set

up (included in

launch checklist)

Testing and analysis

will be performed

during building

phase to make sure

this will not be an

issue

Sharp

Edges on

components

Manufacturing

stage

Minor cuts

when making,

or transporting

rocket

1 3 3

Low

Risk

Sharp edges will be,

smoothened, and de-

burred

Rocket

Launch

causes a

Brush Fire

Dry

Launching

conditions

Small bush fire 2 1 2

Low

Risk

Have equipment to

extinguish fire in

hand and ready to

use

Wait until safety

range officer

determines is safe to

extinguish fire

Arms

Buckle

under load

of rocket

Material

Failure,

May cause

rocket to jam

upon exit

Launch

Platform may

not reach

required

position

1 3 3

Low

Risk

Analysis and testing

will be performed on

configuration of

components to

ensure this will not

happen

Components

Jam

Material

Failure

AGSE will not

reach desired

launch angle

4 2 8

Medium

Risk

All components will

be checked prior to

launch

Vehicle is

not lifted at

fast enough

rate

Poor motor

choice

Time

requirement for

this portion of

the competition

will not be met

3 3 6

Medium

Risk

Test runs of the

entire AGSE will be

performed to ensure

time limit is met

Motor

Failure

Motor

breaking/Short

Rocket will not

erect and may

fall back down

4 2 8

Medium

Risk

This will be taken

into account when


to horizontal

position

developing the

AGSE further

Redundancy measure

will be implemented

Power Loss Electrical

failure

Faulty

batteries

AGSE will not

function

4 1 4

Low

Risk

Extra batteries will

be brought to launch

Batteries will be

tested prior to launch

Redundancy

measures will be

implemented

Igniter not

properly

installed in

motor

Design or set

up failure in

igniter injector

Possible

damage to

rocket motor

during ignition

Loss of vehicle

5 3 15

High

Risk

Additional

mechanism for

verification will be

installed

Verification of

proper set up of this

part of the AGSE

will be done by

multiple team

members Table 39: AGS E Ris k As s es sment Matrix

Stability and Propulsion Risk Assessment

The hazards associated with the stability and propulsion of the launch vehicle are

outlined the table that follows. The team has multiple members with certifications supporting

that the launch vehicle motors can be safely handled. In addition, key team members also have

previous high – powered rocketry experience, allowing for a safe and stable launch vehicle. This

area is considered a low risk for the team, but it is still important to address any potential

problems that the team may face throughout the project.


Motor

Fails to

Ignite

Faulty motor

Problems

with igniter

injector

system

Delayed

ignition

Faulty or

disconnected

e-match

Rocket does not

launch or

launches at an

unpredicted time

3 2 6

Medium

Risk

Follow NAR safety

code

Wait appropriate

amount of time (60

seconds) before

approaching to check

ignition system

Be prepared to remove

the ignition system

from the rocket motor


Motor

Explodes

on Launch

Pad

Faulty motor Significant

damage to the

rocket

4 1 4

Low

Risk

All personal should be

at a safe distance

allowed by NAR

guidelines

Be prepared to

extinguish possible

fires caused by this

incident

Rocket

does not

reach

sufficient

velocity

before

leaving

launch rail

Rocket

weight to

impulse ratio

is not correct

Unstable launch

and

unpredictable

flight trajectory

3 1 3

Low

Risk

Simulations on

Rocksim are run to

ensure correct motor

selection

Fins shear

during

flight

Poorly

constructed

rocket

Not enough

epoxy used to

secure fins to

the frame

Unstable rocket,

Unpredictable

flight path

4 1 4

Low

Risk

Examine rocket fins

for any issues prior to

launch

Confirm that all

personnel are at a safe

distance during launch

Airframe

buckles

during

flight

Airframe

encounters

stresses over

material’s

specifications

Loss of rocket

Rocket becomes

severely

unstable

4 1 4

Low

Risk

Adequate material

selection that can

sustain stresses much

higher than required

Improperly

aligned fins

Fin are

mounted

incorrectly,

not straight

or unequally

spaced

Rocket severely

unstable, may

spin excessively

during flight

4 1 4

Low

Risk

Fins will be installed

with adequate

tolerances so they do

not negatively affect

flight trajectory

Air brake

system

causes

instability

One of the

flaps may

break

Severely

destabilizes

rocket

4 3 12

High

Risk

Testing and analyses

will be performed to

determine maximum

stresses caused on the

flaps

Fins tolerances will be

designed to withstand

much higher stresses

than necessary

High risk due to fact

that this concept hasn’t

been tested.


Table 40: S tabi l i ty/ Propuls ion Ris k As s es sment

Launch Day Risk Assessment

The following table outlines risks that might potentially be encountered during launch

preparations during either the competition day, or any test launched the team plans to conduct.

Hazard Cause Outcome S

V

L

V

Risk Mitigation

Dropping

Rocket during

transportation

Carelessness

when

handling

rocket

Minimal

damage

(scratched) to

components of

rocket

1 2 2

Low risk

Rocket is designed to

be durable due to the

nature of launching

it, however careful

handling should be

practiced and will be

enforced during the

prelaunch safety

briefing

Black Powder

causes charges

to go off early

Altimeter

failure sends

an incorrect

reading

Could cause

serious injuries

and significant

damage to the

rocket

4 1 4

Low

Risk

All electronics will

be kept OFF during

preparation until last

possible moment.

Parachute

failure during

decent

Parachute does

not deploy

Altimeter

failure

Not enough

pressurization

Wrongly

sized

parachute

Parachute gets

stuck and

can’t deploy

Rocket

reaches

ground with

too great

kinetic energy

causing

damage to

rocket

components

or dangerous

situation to

personnel.

Rocket may

fall too slowly

causing it to

drift great

distances

5 1 5

Medium

Risk

Simulations have

been completed to

confirm that

parachutes have been

properly selected

Ground test will be

perform ion the

parachute ejection

system and on the

parachute itself to

verify that each is

working properly

Table 41: Launch Day Ris k As s es sment


Environmental Hazards Risk Assessment

The hazards addressed in the table which follows are risks from the environment that

could affect the launch vehicle, or any of its components. Several of these hazards resulted in a

moderate risk level, and will remain that way for the remainder of the season. These hazards are

the exception for needing to achieve a low risk level. This is because several of these hazards are

out of the team’s control, such as the weather. In the case that environmental hazards present

themselves on launch day, putting the team at a moderate risk, the launch will be delayed until a

low risk level can be achieved. The hazards that the team can control will be mitigated to attain a

low risk level.


Weather

conditions

such as

rain or

high winds

Nature Damage to

electrical

components

Increased

Drifting

Launch may be

cancelled

3 4 12

High

Risk

Weather forecast will

be checked prior to

launch and plans will

be made accordingly

Design a way to

protect electrical

components from

rain

Avoid launching at

high winds when

possible

Trees/

Ponds/

Swamps

Launch site

proximity to

trees

Rocket may land

wrongly causing

it to get tangled

or damaged

3 1 3

Low

Risk

Not to launch with

high winds that may

cause the rocket to

drift too much.

Simulations have

been completed on

Rocksim with

different wind

velocities use this

information to

determine if rocket is

safe to launch

Extremely

Cold

Temp.

Messes with

the batteries

causing them

to discharge

more quickly

Can cause

fiberglass to

shrink

Discharged

batteries might

cause failure of

AGSE or

electrical

problems

causing failure in

setting up black

powder charges

3 1 3

Low

Risk

Batteries will be

inspected prior to

launch. Extra

batteries will be

purchased and

brought to the launch

to use if necessary


Humidity Climate Black Powder

becomes moist

and fails to ignite

2 4 8

Medium

Risk

Motors will be stored

correctly in a

moisture free

environment Table 42: Environment al Effects Ris k As s es sment

DESIGN FAILURE MODES

Although several failure modes were discussed in the preceding sections, the table shown

below discusses them in more detail. There are areas that the team must pay take into closer

consideration when continuing to develop the design in order to minimize their effect on the

launch vehicle.

Potential Failure

Mode

Cause Consequence Mitigation

Parachute Failure Parachute burns due

to ejection charge

Improper installation

of Kevlar blanket

Ensure blanket

completely wraps

around parachute

Parachute detaches

from shock chord

Vehicle has

uncontrolled descent

leading to

catastrophic failure

Securely tie the

parachutes to the

shock chords;

multiple people will

check know strength

Launch Failure Igniter fails to ignite Motor will not

combust; rocket will

not launch

Ensure continuity;

Properly store

igniters

Motor explodes Rocket will not

launch;

Catastrophic damage

to vehicle and AGSE

Proper storage of

motor

Altimeter Failure Leads break free Signals are not sent to

ejection charges;


of vehicle

Install thicker gauge

wire

Altimeter runs out of

battery power

Ejection charges do

not activate;


of vehicle

Put a new battery in

each altimeter before

each launch; ensure

they are fully charged

External Structural

Failure

Rail button separates

while on launch rail.

Rocket has an

undesirable

trajectory.

Proper installation,

alignment, and

location of rail

buttons.

Fins break during

flight due to drag

force.

Rocket is unstable

during flight

Use proper materials

and construction

techniques for fins.


Upper electronics bay

hatch detaches in

flight.

Damage to

electronics. Rocket

has unstable flight.

Construction of the

electronics bay hatch

will ensure a smooth

contour and will be

firmly attached.

Internal Structural

Failure

Internal components

shift during initial

thrust.

Rocket’s center of

gravity shifts,

resulting in an

unstable flight.

Apply enough epoxy

to secure internal

components.

Couplers fail from

being too short.

Body tube

connections are

weak. Rocket breaks

apart during liftoff.

Ensure couplers are

at least one tube

diameter in length to

hold the rocket

together.

Motor Mount fails Motor flies through

the rocket and

damages components.

Rocket flight is

unstable.

Make the forward

motor mount

bulkhead thick

enough

Ejection Charge

Failure

Ejection charges fail

to ignite

Pressure increase is

not sufficient to eject

airframe components.

Uncontrolled descent

of vehicle.

Ground ejection test

Ejection charge too

large

Potential damage to

internal and external

components of

vehicle

Ground ejection test

Separation Failure Premature separation

of rocket components

Damage to rocket due

to unforeseen forces

acting on the vehicle

Ensure connections

are strong and do not

easily shift around Table 43: Fai lure Mode Analys is

SECTION 4: AGSE CRITERIA GROUND SUPPORT PERFORMANCE CRITERIA

The ground support will be considered a success if the following basic criteria are met:

1) The payload is autonomously retrieved and placed in the appropriate launch vehicle

compartment.

2) The payload compartment is autonomously closed after payload insertion.

3) The launch rail erects the launch vehicle to five degrees off vertical.

4) The ignitor is inserted into the launch vehicle engine.


5) The entire process outlined above occurs completely autonomously in under ten

minutes.

OVERALL AGSE SEQUENCE OF EVENTS

The following table outlines the sequence in which the entire autonomous ground support

equipment will perform its tasks.

Sequence Title Description

0 Visual Inspection Visual inspection of the entire

system. This will ensure that

all components and

connections are safely in

place.

1 Power On - Standby Give power to the AGSE.

System in standby mode.

2 Initialization Initializes system; automation

begins.

3 Awaken MAV Rover wakes.

4 Payload Retrieval Rover leaves housing unit,

seeks and retrieve payload.

5 Payload Delivery Rover travels up the ramp to

the platform and safely

deposits the payload in the

rocket. Rover closes payload

hatch.

6 MAV Housing Rover travels back down

ramp and into housing unit.

Rover powers down.

7 AGSE Automation – Standby

to Launch

Linear actuator begins to

retract, propping rail and

rocket. Ignitor travels into the

rocket.

8 Final Visual Inspection Final visual inspections to

verify all system go.

9 Lift-off Button press – we have lift

off. Table 44: AGS E S equence


LAUNCH RAIL DESIGN

Figure 98: Launch Rai l As s embly in Horizontal S tate

Figure 99: Launch Rai l As s embly in Launch Pos i tion

The main support of the AGSE comes from a member system of machine frames and

locking system. The machine frames will be made from 6061T6 aluminum, along with three

styles of connectors that will link between the frame parts. The base dimensions itself are 9’ x 3’

x 1.5’. The framework will be able to hold all of the working components mostly within its foot-

print, aside from the ramp that the rover will use to reach the launch vehicle. The width of the

platform was set to make the platform stable enough so that the launch vehicle will be impossible

tip over with conditions that would be acceptable to launch in. One of the major advantages of


this design is that the launch vehicle gets placed in the center of the AGSE before launch, giving

the launch vehicle good stability during launch.

The box – in style truss structure adds stiffness to the design without adding too much

weight to the overall design. Currently, this structure weighs roughly 41 lbs. Another benefit of

the design is its ability to come apart and be put together again with screws, which enables

mobility of the system. In addition, many of the parts are being manufactured at an external

company with higher precision tools than our facilities have. The final product would have high

tolerances for the final assembly.

The following table outlines overall dimensions for the current launch rail assembly.

Dimension At resting position At launch position

Length 108.00” 108.00”

Width 68.81” 68.81”

Height 31.86” 104.46”

Weight 61.47 lbs 61.47 lbs Table 45: Overal l Launch Rai l Dimens ion

Launch Rail Motion

The original, proposed launch rail design had a stationary hinge at the base of the ground

support equipment, and used a motor and gearbox to raise the rail from the initial horizontal

position to the final position of 5 degrees off vertical. The newer design uses a double connecting

rod system. The linear actuator will pull the rail towards the center of the AGSE and a secondary

support beam that is pinned further up the rail will cause the rail to move to the desired angle.

Initial Geometry Selection

The geometry of the system was selected by optimizing the positions of certain

components, and working around those. The team considered that the over which the support

beam would attach to the rail would be at the center of gravity of the rocket, shown by “Ca” in

the Figures 86 ad 87 below. Since the linear actuator has a specific stroke length (“La” in the

Figure 87), this was fixed the overall displacement of the linear actuator’s attachment point to the

rail. In order to avoid the support beam being horizontal at its starting position, the attachment

was placed a distance denoted by “H” in Figures 86 and 87 on the ground support equipment.

Using the change in the overall length in the geometry, given by “L” in the following images,

and the initial and final angles of the rail; the following formula was created to determine the

length of the launch vehicle supporting rail, denoted by “S”.

𝐶𝑎 + √𝑆2 − 𝐻2 = 𝐿𝑎 + cos(85°) + √𝑆2 − (𝐶𝑎 ∗ sin(85°) + 𝐻)2


Figure 100: Ini tial Launch Rai l configuration

Figure 101: Final Launch Rai l configuration

Using a computational solver, the team was able to solve for several different build

configurations in order to find the ideal configuration given the stroke length of commercially

available linear actuators. The team concluded in placing the 12 inch linear actuator along the

rail. Because the predicted center of gravity was calculated about 40 inches away from the base

of the launch vehicle, the “Ca” ended ups equal to 24 inches. This gave “H” a final value of 18


inches, which would for a large enough angle to reduce the holding stress on the beam as the

vehicle sits on it awaiting launch.

Figure 102: Launch configuration

Furthermore, this configuration allows part of the launch rail to sit in between the overall

framework, lowering the overall center of gravity. This is demonstrated by the image above.

Motion and Stability Analysis

To allow for the movement of the rail, a door hanger mechanism for sliding doors will be

purchased and used as shown in blue and red in the following figure. Each door rail is rated for

200 pounds.


Figure 103: S l iding motion mechanis m

The linear actuator that will be used for this application is rated to move at 0.98 in/s with

no load, and 0.78 in/s at maximum load capacity. The system will be able to fully erect the

vehicle into its final launch position in a minimum of 40 seconds. The importance of the speed at

which this movement occurs is critical for the time constraint that the AGSE has given the MAV

competition regulations. A graph of the angular velocity of the rail during its motion is shown in

the figure below.


Figure 104: Launch Rai l Angular Veloci ty

To determine the stability of the platform, static equilibrium is used to determine when

the sum of the forces and moments are no longer equal to zero in every direction. For simplicity,

the center of gravity of the launch vehicle is used as the reference point for the following

calculations.

Friction force: Ff = µs*N

ΣMcg = 0

where Ff is the friction force, µs is the coefficient of static friction, N is the normal force, and Mcg

is the moment about the center of gravity. With the center of gravity being in the middle, the

smallest moment arm would be laterally.

Thus, the moment at the center of gravity of the ground support with the launch vehicle

included would be: Mcg = 100lb*(1.5ft) = 150lb*ft (approximate weight)

To be able to tip the assembly over, it is assumed that the frictional force would have to

be greater that the moment force:

0.0000E+00

1.0000E+00

2.0000E+00

3.0000E+00

4.0000E+00

5.0000E+00

6.0000E+00

0.000 10.000 20.000 30.000 40.000

sec

Angular velocity of the rail (in degrees/s)

Rail_assemb-1 AngularVelocity2 (deg/sec) Ref.Coordinate System:


150lb*ft = F*(4ft), F = 37.5 lb, approx. center of gravity of the rocket

Once the vehicle is in launch position, the only thing that would foreseeably provide an

external force to the system would be the atmosphere. If wind would generate 37.5 lbs of force

on the vehicle, the launch would not be able to safely proceed.

Ignitor Insertion and Locking

The original ground support design used a secondary beam attached to the rail that would

move along the AGSE and lock into a niche in the AGSE to keep the rail steady. The newer

design, however, uses stops and the static load of the linear actuator to keep the rail steady. The

linear actuator is rated for 400 lbs static force, which is sufficient for the application in which is

it being used. There will also be a stopper in the rail that will prevent the wheel from over-

extending past the desired 85 degrees, regardless of the linear actuator. This will create a

mechanical lock between the stoppers, in addition to the linear actuator that will hold the rail in

place. The supporting member of the rail will also have two side bars that will help reduce any

lateral movement that may otherwise disrupt the vehicle launch.

The ignitor insertion system consists of a repurposed motorized telescoping radio antenna

from a car. The ignitor would be attached to the end, then inserted the 24 inches up the chosen

motor in approximately 10 seconds, thereby igniting the motor.

Components

Framework


Figure 105: Framework Bas e Body

The entire framework assembly may be will able to be disassembled and reassembled for

transport, and may be placed back together with only a wrench. The ground support frame will

be manufactured using piping frame purchased from Misumi. We are using No. 28 FFAU High

Rigidity Type Piping.

Figure 106: Framing Piping Cros s S ection

The High Rigidity Type piping has a cross – sectional area of 11.93 inches – squared.

This will be sufficient enough to prevent any sort of buckling caused by load applied. The piping

is rated for 1/2 pound-per-foot. Since 80 feet worth of piping is required, this calculates to be

roughly 40 pounds worth of piping needed to build the frame.


Figure 107: Bas ic Piping S peci fications

Ramp

Figure 108: Rover Ramp

The ramp that would allow the rover to deliver the payload to the launch vehicle before

launch is made of sheet aluminum, with a supporting structure underneath for strengthening. It is

set a 30 degree angle, which is the maximum allowable angle the integration of the rover would

allow. This angle will be experimented with to increase the speed of the rover, while retaining as

small of a footprint as possible.

Joints

The connecting piece used between the various beams are machine frame joints; they

were designed to connect and attach using the naturally high coefficient of friction of aluminum.

On the much thinner beams, the stress of the load is distributed over a large surface area rather

than a small bolt hole. In total, there are three different connecting joints being used for this

application, as can be seen in the following three images. They consist of three parts that clamp

onto the aluminum extrusions, creating the final joint.


Figure 109: Two piece joint

This joint is capable of connecting two beams together, and will be used to create the

outer truss framework of the launch rail.

Figure 110: Three piece joint


The joint pictured above will be used to connect three extruded beams together. It will be

used to connect the corners of the truss framework.

Figure 111: Articulated joint

This is the last type of joint that will be used for this launch rail. It will be smoothed and

greased in order to reduce the coefficient of friction and the tension on the bolt. This joint will be

used to allow the rotation movement needed to erect the launch vehicle into its launch

configuration.


Telescoping Antenna

Figure 112: Repurpos ed car antenna

The process by which this antenna would insert the antenna is explained in the launch rail

motion section above.

Linear Actuator

Figure 113: PA-04 Linear Actuator

The planned linear actuator to be used for this application will be the one pictured above

from Progressive Automations. The PA – 04 linear actuator has an IP – 66 rating, and is

enclosed, allowing for both water and dust resistance from the actuator’s inner components.

Having dust tight enclosures on components is always beneficial to the reliability of any system.


Figure 114: PA-04 Linear Actuator dimens ions

Input Voltage 12V DC

Stroke 40 inches

Force 400 lbs

Speed 0.98"/sec (400 lbs)

Protection Class IP66

Operational Temperature 26ºC~+65ºC

Noise db<45(A)

Duty Cycle 20%

Limit Switch Built In, Non-Adjustable

Current (full load) (full load) 12 A

Mounting Holes 0.40"

Screw Type ACME

Housing Type Aluminum allow

Wire Length 60"

Fully Retracted 7.87" + 40”

Fully Extended 7.87" + 40” + 40” Table 46: PA-04 S peci fications

The member that connects the 80-20 rail to the PA-04 Linear Actuator is at a 14 degrees

incline, and it is estimate that our rocket and rail system will weigh at 50lbs at its center of

gravity. By using trigonometry, the team found that a 145 pound-force is needed to raise the

launch vehicle into its final configuration. The PA-04 Linear Actuator is rated for a 400 pound-

force load, which is roughly 2.7 times more force than is required. This is gives us a safe margin,

ensuring the launch vehicle will be safely erected. The static load is rated the same as the push

load, therefore, the actuator will be able to firmly hold the launch vehicle in place.


Figure 115: 12V DC, S peed vs . Load

At sequence initiation, the linear actuator is expected to be fully extended, and fully

retracted at launch. Because the PA-04 model is rated for 0.98 in/s, with a 40-inch stroke, the

team expects roughly 41 seconds of driving time.

Figure 116: 12V DC, Current vs . Load

Since the heaviest force the linear actuator will experience is 145 pounds, we expect a

maximum current draw of 8 amperes.


Component CFD analysis

Several CAD simulations were run the various members in order to make sure all loads

would be safely supported. These analysis simulation may be seen in the following images.

Figure 117: Load on Main S upporting Beam

This is a simulation of the supporting member of the rail. The load would be distributed

across the beam because the door rail being used would bear the load it was rated for, and pass it

on to the supporting member underneath. For optimum safety purposes, an overestimation of 50

lbs of force from the launch vehicle was simulated on the rail. Each member would only see half

of that force, so 25 lbs was used for the static simulation. The factor of safety was about 5.5,

which is well within reason for safety.


Figure 118: S upporting beam buck l ing s imulation

The vertical members of the launch rail assembly would only undergo compression,

therefore a buckling simulation was conducted on these members. This simulation was run under

an overestimated load of 75 lbs for safety reasons. The results from the simulation indicated a

safe choice in member, giving a factor of safety of 3.2. Further design development could result

in the use of a high rigidity variant of this beam, where the thickness would be three times the

thickness of the current beam being used. This decision will be evaluated at a later point of the

design development.


Overall Assembly Summary

Figure 119: Top View, Ground s upport

Figure 120: S ide View, Ground s upport

Figure 121: Ground S upport As s embly Components

Part No. Name Description Weight (lbs)


1 Sheet Metal Ramp A sheet metal ramp

with bracing

underneath. MAV

will travel and deliver

payload to rocket by

climbing the ramp

5

2 Linear Actuator 48-inch linear

actuator that will

retract and erect the

railing.

42

3 10ft 80/20 Aluminum

Launch Rail

Rocket will sit on this

railing. Railing will

provide stability for

lift off.

2.5

4 Ignition Plate Steel Plate that

protects electronics

from exhaust.

5

5 Guide Rails and

wheels

Guide rails for wheels

to travel and allows

Launch Rail’s

movement.

5

6 Frame Aluminum frame

where components

are mounted to and

provides ground

clearance.

25

7 Stability Rails Aluminum tubes that

provide extra

stability.

15

8 (not pictured) Battery Provide power to

AGSE System. 12V

20Ah

10.5

9 (not pictured) Integrated System Beaglebone

Microcontroller used

to automate AGSE.

This includes various

electronics, circuits,

relays, and wiring.

10

TOTAL WEIGHT 120 Table 47: Ground S upport Component Des cription

As is seen in the above table, the total weight of the system does not surpass the

maximum allowed weight of 150 lbs.


ROVER DESIGN

The payload retrieving rover will be four – wheeled, servo controlled, and computer -

controlled by a RaspberryPi. Servos will be used to control the autonomous motion of the rover,

in addition to controlling the extendable claw used to retrieve the payload and place it in its

respective compartment. Python will be used to write the code for rover retrieval process. The

payload will be searched for by a webcam and computer vision through openCV – python. After

the payload has been located by the rover, it will navigate to it and retrieve it with the claw.

The payload compartment in the launch vehicle will be marked with a specific color

combination recognizable by the rover’s computer vision, thus guiding it to the correct position.

The structure of the rover will be 3D printed in order to reduce weight and add flexibility the

structure’s design process. The rover will have the ability to be started and paused through a

wireless connection with the Raspberry Pi.

Selected Components

RaspberryPi

Figure 122: Ras pberryPi computer

The Raspberry Pi was chosen for this project for several reasons. The RaspberryPi is a

small and versatile microcontroller. It is able to run a full version of a linux operating system due

to its embedded linux platform. In addition, this microcontroller is more geared towards video


processing, which is necessary for this given application. The advantage of working in a linux

environment is the ability to use commercially available equipment such as USB webcams,

Bluetooth, and wireless dongles. OpenCV python was chosen because it can be used natively and

without modification on the RaspberryPi.

Webcam

Figure 123: S tandard Webcam

Due to the flexible nature of the RaspberryPi, the linux operating system, and openCV-

python, the team will be able to buy an ordinary USB webcam.

Wheels

To accommodate different terrains, the team has decided to use four wheels in the rover

design. Wheels with some tread will be used to providetraction while moving and retrieving the

payload.


Continuous Servos

Figure 124: Continuous S ervo

To actuate the wheels, continuous servos will be used instead of motors. Continuous

servos are easier to use with microcontrollers than motors. In addition, servos require

significantly less power to operate than motors. The lower power consumption allows for smaller

batteries, which alleviates overall weight. The servos we are using have a fair amount of torque

which is useful when moving the rover.


Standard Servo

Figure 125: S tandard S ervo

Regular servos will be used to actuate the claw. The webcam will be mounted on the

claw that will retrieve the payload, which will move both laterally and axially to center the

payload in the frame of the camera. Servos will also be used to actuate the gripping mechanism

that will carry the payload to the rocket. The servos are useful for this function for many of the

same reasons that justify their use to move the rover. The have low power consumption and they

are easy to use with microcontrollers.

Overall Process

The rover will be comprise of four wheels, two of which will be powered by continuous

servos. On the rover there will be a claw with a webcam will be mounted atop it. This will serve

to track and retrieve the payload. The claw will laterally and axially around the area to track the

payload. The team has written openCV code that is able to track an object and give its location

within the frame. The position of the payload will be used to move the servos on the claw and

center the payload in the middle of the frame.

While the payload is being tracked, the rover will be moving towards the payload. Once

the rover is a certain distance from the payload, it will stop and actuate the claw to pick up the

payload. Once in possession of the payload, the rover will make its way to the launch vehicle.

The launch vehicle will be marked with a color cue that will be recognized by the openCV

program. Once at the launch vehicle payload compartment, the claw will deposit the payload in

the compartment.


OpenCV Tracking Code

1. import numpy as np

2. import cv2

3.

4.

5. def run_main():

6. cap = cv2.VideoCapture('rover.mp4')

7.

8.

9. # Read the first frame of the video

10. ret, frame = cap.read()

11.

12.

13. # Set the ROI (Region of Interest).

14.

15. c,r,w,h = 900,650,70,70

16. track_window = (c,r,w,h)

17.

18.

19. # Create mask and normalized histogram

20. roi = frame[r:r+h, c:c+w]

21. hsv_roi = cv2.cvtColor(roi, cv2.COLOR_BGR2HSV)

22. mask = cv2.inRange(hsv_roi, np.array((0., 30.,32.)), np.array((180.,255.,255.)))

23. roi_hist = cv2.calcHist([hsv_roi], [0], mask, [180], [0, 180])

24. cv2.normalize(roi_hist, roi_hist, 0, 255, cv2.NORM_MINMAX)


25. term_crit = (cv2.TERM_CRITERIA_EPS | cv2.TERM_CRITERIA_COUNT, 80, 1)

26.

27.

28. while True:

29. ret, frame = cap.read()

30.

31.

32. hsv = cv2.cvtColor(frame, cv2.COLOR_BGR2HSV)

33. dst = cv2.calcBackProject([hsv], [0], roi_hist, [0,180], 1)

34.

35.

36. ret, track_window = cv2.meanShift(dst, track_window, term_crit)

37.

38.

39. x,y,w,h = track_window

40. cv2.rectangle(frame, (x,y), (x+w,y+h), 255, 2)

41. cv2.putText(frame, 'Tracked', (x-25,y-10), cv2.FONT_HERSHEY_SIMPLEX,

42. 1, (255,255,255), 2, cv2.CV_AA)

43.

44.

45. cv2.imshow('Tracking', frame)

46.

47.

48. if cv2.waitKey(1) & 0xFF == ord('q'):

49. break


50.

51.

52. cap.release()

53. cv2.destroyAllWindows()

54.

55.

56. if __name__ == "__main__":

57. run_main()

AGSE VERIFICATION

The following table outlines the rules given by the Centennial Challenge portion of the

competition, and the methods the team will use to verify each competition requirement is met.

Requirement

Number


3.3.2.1.1 Teams will position their

launch vehicle

horizontally on the AGSE.

The launch rail of the vehicle

has been designed to start in a

horizontal position, and lift

the launch vehicle into its

final position.

The continual development of

the launch rail design will

always assume a horizontal

starting position.

3.3.2.1.2 A master switch will be

activated to power on all

autonomous procedures

and subroutines.

The official launch controller

will house a master switch

capable of providing power to

all AGSE processes.

The team will be aware of the

schematic needed to make sure

the master switch will be able

to perform its function

accordingly.

3.3.2.1.3 All AGSEs will be

equipped with a pause

switch in the event that a

judge needs the

AGSE to be temporarily

halted for any reason. The

pause switch halts all

AGSE procedures and

subroutines. Once the

pause switch is

deactivated the AGSE

resumes operation.

The final launch controller

will house a pause switch

capable of starting and

stopping the entire

autonomous process at will.

The team will test this switch

before competition day to

ensure it is capable of halting

all autonomous procedures.

3.3.3.2 All AGSE systems shall

be fully autonomous. The

The design of the AGSE has

been carried out such that all

The team will continue the

development of the AGSE as a


only human interaction

will be if the judge pauses

the AGSE.

integrated systems are fully

autonomous.

fully autonomous system. The

system will be tested before the

competition day to make sure

all systems have been

integrated properly.

3.3.3.3 The AGSE shall be

limited to a weight of 150

pounds or less and volume

of 12 feet in height x

12 feet in length x 10 feet

in width.

The current design has been

developed taking into account

the limit in volume and

weight provided by this

metric.

The team will keep this

limitation in mind if any

changes are applied to the

current AGSE design.

3.3.4 3.4.1.1 As one of the goals

of this competition is to

develop equipment,

processes, and

technologies that could be

implemented in a Martian

environment, the AGSE

and any related

technology cannot employ

processes that would not

work in such

environments.

Therefore, prohibited

technologies include:

3.3.4.1.2 Sensors that rely

on Earth’s magnetic field

3.3.4.1.3 Ultrasonic or

other sound-based sensors

3.3.4.1.4 Earth-based or

Earth orbit-based radio

aids (e.g. GPS, VOR, cell

phone).

3.3.4.1.5 Open circuit

pneumatics

3.3.4.1.6 Air breathing

systems

The current AGSE design has

none of the prohibited

components outlines by the

competition rules.

Any alteration in the current

AGSE design will be done

without any of the technology

prohibited by this requirement.

3.3.5.1 Each launch vehicle must

have the space to contain

a cylindrical payload

approximately 3/4 inch

inner diameter and 4.75

inches in length. The

payload will be made of

¾x 3 inch

The payload bay designed for

the launch vehicle has been

dimensioned such that the

payload to be provided by the

competition will fit

comfortably in the

compartment designed for it.

The team will make sure any

changes in this compartment

are done in accordance with

the payload dimensions

provided by this requirement.


Schedule 40 PVC tubing

filled primarily with sand

and may include BBs,

weighing approximately 4

ounces and capped with

domed PVC end caps.

Each launch vehicle must

be able to seal the payload

containment area

autonomously prior to

launch.

3.3.5.2 A diagram of the payload

and a sample payload will

be provided to each team

at time of acceptance into

the competition. In

addition, Teams may

construct practice

payloads according to the

above specifications;

however, each team will

be required to use a

regulation payload

provided to them on

launch day.

The team will construct

practice payloads in

accordance with the

dimensions given for the

payload by the competition

rules.

The practice payloads will be

used to test the process of

payload insertion into the

launch vehicle. This process

will be tested multiple times

for optimum efficiency.

3.3.5.3 The payload will not

contain any hooks or other

means to grab it.

The practice payload will not

have any means to ease its

ability to be grabbed.

The team has no plans to use

any hooks to practice grabbing

the payload.

3.3.5.4 The payload shall be

placed a minimum of 12

inches away from the

AGSE and outer mold line

of the launch vehicle in

the launch area for

insertion, when placed in

the horizontal position on

the AGSE and will beat

the discretion of the team

as long as it meets the

minimum placement

requirements.

The method of obtaining and

inserting the payload will be

designed such that the

autonomous process may be

retrieved at a minimum

distance of 12 inches away

from the AGSE.

The team will run practice runs

of the autonomous system to

ensure a minimum distance of

12 inches may be successfully

achieved.

3.3.5.5 Gravity-assist shall not be

used to place the payload

within the rocket.

The current payload insertion

method will not use any form

of gravity – assist.

The team has no plans to

include gravity – assist as a

method of insertion for the

payload.


3.3.5.6 Each team will be given

10 minutes to

autonomously capture,

place, and seal the

payload within their

rocket, and erect the

rocket to a vertical launch

position five degrees off

vertical. Insertion of

igniter and activation for

launch are also included in

this time.

Design measures are being

taken so that the payload

retrieval and launch rail

systems take place under the

competition delegated time.

The entire integrated system

will be tested and timed so the

10 minute requirement is not

exceeded.

3.3.6.1.1-2 These requirements were

covered in sections

3.3.2.1-2

3.3.6.1.3 A safety light that

indicates that the AGSE

power is turned on. The

light must be

amber/orange in color. It

will flash at a frequency

of 1 Hz when the AGSE is

powered on, and will be

solid in color when the

AGSE is paused while

power is still supplied.

The safety light will be

visibly incorporated on the

side of the launch rail to show

that power is being delivered

to the system.

This light function will be

tested with the entire AGSE

assembly to test functionality.

3.3.6.1.4 An all systems go light to

verify all systems have

passed safety verifications

and the rocket system is

ready to launch.

A green light will be visibly

incorporated onto the side of

the launch rail. This light will

turn on when the master

arming switch is activated.

The functionality of the light

will be tested to ensure proper

functionality.

Table 48: AGS E Verification


SECTION 5: PROJECT PLAN BUDGET PLAN

Launch Vehicle

Figure 126: Launch Vehicle Cos t Chart

Nosecone section

4%

Parachute Bay

16%

Electronics Bay

28%

AirBarke System

4%

Propulsion Bay

33%

Payload

4%

Other

11%

Cost Break Down

Nosecone section Parachute Bay Electronics Bay AirBarke System Propulsion Bay Payload Other


AGSE

Table 49: Comprehens ive Launch Vehicle Budget


Figure 127: AGS E Budget Chart

Table 50: Comprehens ive AGS E Budget

FUNDING PLAN

For the 2016-2017 NASA USLI Competition, the team plans on raising funds from both

local and national contributors.

The team will call, or personally visit, different types of engineering companies and firms

and present them our funding proposal. The team hopes to foster a sponsorship between FIU and

these companies by showcasing our ideas. This, in turn, will further FIU’s relationship with

AGSE, 85%

AGSE Budget

Rover AGSE


various STEM employers located in South Florida, as well as increasing exposure of these

companies.

On campus, the team will display last year’s competition launch vehicle in a zone with

heavy foot traffic. This will inspire students and awaken their curiosity in rocketry and of the

building process involved in making a functioning launch vehicle. The team will have

conversations with these students and answer all the questions they have; and hopes that by

educating the student body at FIU, they will be more susceptible to contributing towards this

project. The team also plans on 3D printing charms and keychains for fundraising for this

competition.

ASME has grants and opportunities for teams that take on projects such as NASA USLI.

The team will write a proposal and submit an application for these grants. FIU also has grants

that the team will apply for as the well.

In the past, the local community has been supportive of ASME and the PantherWorks

Space team. By outreaching to the local public schools, (see Educational Plan and Engagement)

the team hopes to generate enough public interest for contributions.

Web presence is a pivotal part of fundraising. The team has begun creating a social media

(Facebook, Reddit, Instagram, etc.) and a GoFundMe and Kickstarter. By presenting a proposal,

the team looks forward to generating support from people around the nation. The team will also

post up images and tutorials showing the different techniques used in making this year’s

competition launch vehicle.

TIMELINE

The development of a timeline for the project is imperative to the overall success of the

launch vehicle. Without a clear plan of action, the team will have no definite direction on what

needs to be accomplished next. A timeline provides a form of accountability for the team,

allowing all members the opportunity to be aware of what tasks need to be accomplished by what

time. Two separate timelines are included in this report, a general timeline developed for the

entire scope of the project, and a detailed timeline developed for the purpose of the preliminary

and critical design reviews. As the overall design continues to mature, more details may be

added to the timeline. In addition, further deadlines will be added as for the flight readiness

review as critical design review nears. These references may be found in Appendix B.


EDUCATIONAL ENGAGEMENT

The PantherWorks Space team strives for educational outreach. The team strongly

believes that by educating the generations that come after us, the team is also encouraging a

better and brighter future.

The goal of the PantherWorks Space team is to teach middle school – aged students the

basics of rocketry; and potentially interest them in pursuing a career in science, technology,

engineering, or mathematics. The area around Florida International University has many middle

schools with minority students that would not normally pursue higher education. With that in

mind, the team has partnered with FIU’s new community outreach called “Engineers on

Wheels”. This program was created with the intent of providing local K-12 students with fun and

interactive presentations on varying fields of engineering.

This year’s Education Engagement Plan focuses on the team visiting surrounding public

schools and give a simple lecture on the basics of rocketry. The team will focus on the

Newtonian motion of rockets and explain the different forces, that a launch vehicle would

experience in flight. The team will also encourage the students to come up and brainstorm

different ideas for propellants, aerospace structures, and methods of launch by asking different

questions in a group discussion. This will engage the students and have them share their

creativity.

Once the lecture is done, the team will introduce a “Build-A-Rocket” workshop, where

the students will design their own launch vehicle out of a 2-Liter bottle and construction paper.

Once done, they will be launched by compressing water in the 2-Liter bottle with a bike pump.

For older students, the team will present last year’s competition launch vehicle, show them the

various techniques the team used to manufacture it, and have a basic presentation on rocketry.

Afterwards the team will have a workshop where the students make paper airplanes and have a

competition to see which model traveled the furthest (this will help students visualize how

different geometric shapes have different aero properties).

Each visit at a school will last for three hours. One visit has already been successfully

conducted, and two more are planned for this year. Another three visits will be conducted before

the competition as well.


SECTION 6: APPENDICES

Appendix A – Stability Analysis Graphs

5 – MPH WIND

Figure 128: CG/CP vs . Time, 5 -mph wind

Figure 129: S tatic Margin vs . Time, 5 -mph wind


Figure 130: Alti tude vs . Time, 5 -mph wind

Figure 131: Range (Drift) vs . Time, 5 -mph s peed


Figure 132: Mach Number vs . Time, 5 -mph wind

10 – MPH WIND



Figure 134: S tatic Margi n vs . Time, 10 -mph wind



Figure 136: Range (Drift) vs . Time, 10 -mph wind

Figure 137: Mach Number vs . Time, 10 -mph


15 – MPH WIND





Figure 141: Range (Drift) vs . Time, 15 -mph



20 – MPH WIND



Figure 146: Range (Drift) vs . Time, 20 -mph



Appendix B – Timelines

GENERAL TIMELINE


DETAILED PDR/CDR TIMELINE

fiu_nsl_pdr

Documents

vehicle summary

launch vehicle design

design of launch vehicle

launch vehicle verification

vehicle criteria changes

launch vehicle assembly

team summary

agse criteria changes