2012 – 2013 georgia tech ramblin’ rocketeers flight ... gatech usli frr.pdf · g 2012 – 2013...

2012 – 2013 Georgia Tech Ramblin’ Rocketeers Flight Readiness Review

G 2012 – 2013 GEORGIA TECH RAMBLIN’ ROCKETEERS

FLIGHT READINESS REVIEW

Georgia Institute of Technology 2 of 187 Ramblin’ Rocketeers

PAGE INTENTIONALLY LEFT BLANK




Table of Contents

Table of Contents ............................................................................................................................ 3 Table of Figures .............................................................................................................................. 8 Table of Tables ............................................................................................................................. 10 1. Introduction ........................................................................................................................... 12

School Information and NAR Section Contacts ............................................................ 12 1.1. Work Breakdown Structure ............................................................................................ 13 1.2. Launch Vehicle Summary .............................................................................................. 14 1.3.

1.3.1. Overview ................................................................................................................. 14 1.3.2. Changes since CDR ................................................................................................ 14

Payload Summary .......................................................................................................... 14 1.4.1.4.1. Overview ................................................................................................................. 14 1.4.2. Changes Since CDR ................................................................................................ 15 1.4.2.1. Payload Changes Since CDR .............................................................................. 15 1.4.2.2. Avionics Changes Since CDR............................................................................. 15

2. Project L.S.I.M. Overview ..................................................................................................... 16 Mission Statement .......................................................................................................... 16 2.1. Requirements Flow Down .............................................................................................. 16 2.2. Mission Objectives and Mission Success Criteria ......................................................... 17 2.3. System Requirements Verification Matrix (RVM) ........................................................ 17 2.1. Mission Profile ............................................................................................................... 30 2.1.

3. Launch Vehicle ...................................................................................................................... 32 Overview ........................................................................................................................ 32 3.1.

3.1.1. Mission Criteria ...................................................................................................... 32 System Design Overview ............................................................................................... 33 3.2. Recovery System ............................................................................................................ 42 3.3.

3.3.1. Altimeters ................................................................................................................ 45 3.3.2. Arming Switches ..................................................................................................... 46 3.3.3. Parachute Dimensions ............................................................................................. 47 3.3.4. Drift Profile Analysis .............................................................................................. 48 3.3.5. Kinetic Energy of Launch Vehicle ......................................................................... 50 3.3.6. Ejection Charges ..................................................................................................... 51 3.3.7. Testing..................................................................................................................... 52

Structure ......................................................................................................................... 53 3.4.3.4.1. Construction ............................................................................................................ 54 3.4.2. Payload Integration ................................................................................................. 55




3.4.3. Avionics Integration ................................................................................................ 55 3.4.4. Section Integration .................................................................................................. 56

Launch Vehicle Performance Analysis .......................................................................... 56 3.5.3.5.1. Altitude Predictions and Motor Selection ............................................................... 56 3.5.2. Stability ................................................................................................................... 59 3.5.3. Testing..................................................................................................................... 60 3.5.4. ...................................................................................................................................... 61 3.5.5. ...................................................................................................................................... 61 3.5.6. ...................................................................................................................................... 61

Intimidator 5 Kit Mass Breakdown ............................................................................... 61 3.6. Interfaces and Integration ............................................................................................... 62 3.7.

3.7.1. Interface with the Ground ....................................................................................... 63 3.7.2. Interface with the Ground Launch System ............................................................. 63

Launch Vehicle Operations ............................................................................................ 63 3.8.3.8.1. Launch Checklist .................................................................................................... 63

4. Flight Experiment .................................................................................................................. 65 Introduction to the Experiment and Payload Concept Features & Definition ............... 65 4.1. Accomplishments Since CDR ........................................................................................ 66 4.2.

4.2.1. Important Changes .................................................................................................. 66 4.2.2. Test Launch Lessons Learned ................................................................................. 66 4.2.2.1. Summary of Science Team Payload .................................................................... 66 4.2.2.2. Report of Failures and Occurrences .................................................................... 67 4.2.2.3. Integration ........................................................................................................... 67 4.2.2.4. Sensor detachment ............................................................................................... 67 4.2.2.5. Openlog File Writing .......................................................................................... 67 4.2.2.6. Results and Future Mitigation ............................................................................. 67 4.2.2.7. Integration Results ............................................................................................... 67 4.2.2.8. Sensor data .......................................................................................................... 68 4.2.2.9. OpenLog Risk Mitigation .................................................................................... 68

Science Background ....................................................................................................... 69 4.3.4.3.1. Important Highlights ............................................................................................... 69

Experiment Requirements and Objectives ..................................................................... 69 4.4.4.4.1. Success Criteria ....................................................................................................... 69 4.4.2. Requirements .......................................................................................................... 70 4.4.3. Hypothesis and Premise .......................................................................................... 74 4.4.4. Experimental Method and Relevance of Data ........................................................ 75

Testing plan .................................................................................................................... 76 4.5.




4.5.1. Overview ................................................................................................................. 76 4.5.2. MR Fluid Creation and Validation of Theory ......................................................... 79 4.5.3. MR Fluid Shear Stress Characterization: Two Plate Test ..................................... 80 4.5.4. Working Ground Model .......................................................................................... 82 4.5.5. Sensors .................................................................................................................... 82

Design review ................................................................................................................. 82 4.6.4.6.1. Viscosity Test Rig ................................................................................................... 82 4.6.2. Ground Test – MR fluid production and manipulation .......................................... 84 4.6.3. Hardware and build progress .................................................................................. 85 4.6.3.1. Sensing ................................................................................................................ 85 4.6.3.2. Solenoids ............................................................................................................. 89 4.6.3.3. Microcontroller .................................................................................................... 91

Payload Relevance and Science Merit ........................................................................... 91 4.7. RGEFP ........................................................................................................................... 93 4.8.

4.8.1. RGEFP-Specific Design Work ............................................................................... 93 4.8.1.1. Containment Box................................................................................................. 93 4.8.1.2. Computer ............................................................................................................. 95 4.8.1.3. Weights ................................................................................................................ 96 4.8.1.4. Equipment Layout for Take-off, in Flight, and Landing ..................................... 98

Flight Experiment Integration ...................................................................................... 100 4.9.5. Flight Avionics .................................................................................................................... 104

Avionics Overview ....................................................................................................... 104 5.1. Avionics Success Criteria ............................................................................................. 106 5.2. SIDES Design Approach .............................................................................................. 107 5.3.

5.3.1. SIDESboard .......................................................................................................... 108 5.3.2. SIDES Electrical Harness ..................................................................................... 109 5.3.3. Master IMU ........................................................................................................... 110 5.3.4. Science Experiment Computer ............................................................................. 110 5.3.5. Telemetry .............................................................................................................. 110

De-scope Options ......................................................................................................... 111 5.4. Power Budget ............................................................................................................... 111 5.5. EM Interference ............................................................................................................ 112 5.6. Transmission Frequencies and Protocols ..................................................................... 112 5.7. Software Maturity ........................................................................................................ 113 5.8. De-scope Option: Flight Computer Definition ............................................................ 113 5.9. Avionics Testing and Reliability Assurance ............................................................ 115 5.10. Ground Station .......................................................................................................... 116 5.11.

5.11.1. Purpose .................................................................................................................. 116 5.11.2. Function ................................................................................................................ 117




5.11.3. Design Considerations .......................................................................................... 119 5.11.3.1. Choice of Antenna ............................................................................................. 119 5.11.3.2. Choice of Camera .............................................................................................. 120 5.11.3.3. Motor Sizing ...................................................................................................... 120 5.11.3.4. Software Maturity ............................................................................................. 122 5.11.3.5. Effects of Excess RF Radiation on the Recovery Avionics .............................. 127

Avionics Mechanical Integration.............................................................................. 128 5.12.6. General Safety ..................................................................................................................... 129

Vehicle Safety and Environment .................................................................................. 129 6.1.6.1.1. Overview ............................................................................................................... 129 6.1.2. Mission Assurance ................................................................................................ 130

Payload Safety .............................................................................................................. 132 6.2. Personnel and Environmental Hazards ........................................................................ 136 6.3.

7. Project Budget ..................................................................................................................... 141 Funding Overview ........................................................................................................ 141 7.1. Current Sponsors .......................................................................................................... 142 7.2. Actual Project Cost ....................................................................................................... 142 7.3.

7.3.1. FRR Budget Summary .......................................................................................... 142 7.3.2. System-Level Budget Summary ........................................................................... 143 7.3.3. Flight Hardware Expenditures .............................................................................. 144 7.3.3.1. Flight Hardware Expenditure Overview ........................................................... 144 7.3.3.2. Flight Hardware Cost Breakdown ..................................................................... 145

8. Project Schedule .................................................................................................................. 147 Schedule Overview ...................................................................................................... 147 8.1. Critical Path Chart: CDR to PLAR .............................................................................. 148 8.2. Schedule Risk ............................................................................................................... 150 8.3.

8.3.1. High Risk Items .................................................................................................... 150 8.3.2. Low-to-Moderate Risk Tasks ............................................................................... 151

9. Educational Engagement Plan and Status ........................................................................... 153 Overview ...................................................................................................................... 153 9.1. Atlanta Makers’ Faire ................................................................................................... 153 9.2. FIRST Lego League and Tech Challenge .................................................................... 153 9.3.

References ................................................................................................................................... 156 Appendix I: Gantt Chart.............................................................................................................. 157 Appendix II: Launch Checklist ................................................................................................... 159 Appendix III: Science Overview ................................................................................................ 164 Appendix IV: Ground Test Plan ................................................................................................. 177




Appendix V: Science MFOs and Drawings ................................................................................ 179 Appendix VI: Altimeter Wiring Harness Schematic .................................................................. 187




Table of Figures

Figure 1. 2012 – 2013 project work breakdown structure. .......................................................... 13 Figure 2. Flow down of requirements. .......................................................................................... 16 Figure 3. Project L.S.I.M. mission profile. ................................................................................... 31 Figure 4: Internal Layout of the Launch Vehicle .......................................................................... 43 Figure 5: Drogue Parachute Assembly ......................................................................................... 44 Figure 6: Main Parachute Assembly ............................................................................................. 45 Figure 7: Electronic Altimeter Schematic ..................................................................................... 46 Figure 8: Featherweight Screw Switches ...................................................................................... 47 Figure 9: Payload Integration Structure ........................................................................................ 55 Figure 10: Avionics Integration Structure .................................................................................... 56 Figure 11: L1390 Altitude and Thrust vs. Time ........................................................................... 58 Figure 12: Launch Vehicle Stability vs. Time .............................................................................. 59 Figure 13: 45% Scale Test Rocket and Flight .............................................................................. 60 Figure 14: Intimidator 5 Kit Landing Mass Breakdown ............................................................... 62 Figure 15: Correcting for Piezo drift ........................................................................................... 68 Figure 16: FFT of corrected data showing peak around 26 Hz ................................................... 68 Figure 17: LSIM testing logic, illustrating a simple relationship of information between the test sequences and emphasizing that they flow down from the pursuit of the LSIM hypothesis. ....... 78 Figure 18: Preliminary static testing of MR fluid mixtures in magnetic fields ........................... 80 Figure 19: Shear stress of a fluid using the two-plate test (Source: Wikipedia) ......................... 81 Figure 20: Piezo-electric sensor used for detecting anchor force oscillations ............................. 86 Figure 21: Piezo-electric sensor circuit. The sensor is modeled as a variable-voltage source at 300 Hz. While 300 Hz is a theoretical maximum for the reading speed of the microcontroller, data was logged at a rate between 88-96 Hz. ................................................................................ 87 Figure 22: Top and bottom view of sensor prototype circuit. Leads soldered to the piezo-electric sensors are attached to the blue terminals, while pins go to the microcontroller for data logging and analog reading. This prototype supports two sensors and is approximately 3 inches by 5 inches. Final boards may be much smaller. ................................................................................. 87 Figure 23: data showing sensor drift and a method of correction by distributing the data around the overall mean. ........................................................................................................................... 89 Figure 24: frequency spectrum for the entire dataset. .................................................................. 89 Figure 25: 4x4 solenoid driver. Two drivers can be linked together per microcontroller to control 32 solenoids directly. Approximately 3 inches by 2 inches. ............................................ 90 Figure 26: Arduino Mega microcontroller with major dimensions. ............................................ 91 Figure 27: Ideas for the containment box, illustrating some support elements and a possible electrical conduit. .......................................................................................................................... 94 Figure 28: bottom mounting bracket for the USLI sounding rocket. A larger version is intended to be used in the containment box. This piece attaches to the box - a second part attaches to the canister and snaps into the bracket. ............................................................................................... 95 Figure 29: Current weight budget with totals, and broken out by known subassemblies. .......... 97




Figure 30: Summary of the weight budget to report subassembly totals. .................................... 98 Figure 31: Equipment layout for containment box, 6 canisters, laptop and crew for all stages of flight. ............................................................................................................................................. 99 Figure 32: Payload Assembly .................................................................................................... 100 Figure 33: Payload Base with 150N of loading .......................................................................... 101 Figure 34: Factor of Safety vs. Total Load from SolidWorks SimulationXpress and generated trend line equation....................................................................................................................... 103 Figure 35: SIDES system layout ................................................................................................ 108 Figure 36: SIDESboard bottom side view ................................................................................. 109 Figure 37: SIDESboard top side view ....................................................................................... 109 Figure 38: Xbee transceiver unit ................................................................................................ 110 Figure 39: Antenna performance as a function of range ............................................................ 113 Figure 40: Generalization of flight computer software .............................................................. 114 Figure 41: Diagram of a helical antenna .................................................................................... 119 Figure 42: Typical radiation pattern for a helical antenna ......................................................... 119 Figure 43: Canon Powershot SX260 ......................................................................................... 120 Figure 44: High-Level Software Process .................................................................................... 123 Figure 45: Updating Rocket State ............................................................................................... 124 Figure 46: Updating Servo Position ............................................................................................ 125 Figure 47: Updating Camera Zoom ............................................................................................ 126 Figure 48:Transmit Rocket Location .......................................................................................... 127 Figure 49. System expenditure summary at CDR. ..................................................................... 143 Figure 50. Sub-system Testing/Development Breakdown. ........................................................ 144 Figure 51. Sub-System Flight Hardware Breakdown. ................................................................ 145 Figure 52. Flight Systems flight hardware breakout. .................................................................. 146 Figure 6. Critical Path Chart from CDR to PLAR ...................................................................... 149 Figure 20. Participation at the Atlanta Makers' Faire. ................................................................ 153 Figure 21: Previous FIRST Lego League outreach event. .......................................................... 153 Figure 56: FLL Regional Event at Wheel High School .............................................................. 154 Figure 57: FLL Regional Straw Rocket Activity ........................................................................ 155 Figure 58: Plot of B field magnitude in MR fluid versus magnitude of vector 𝝁𝟎𝑯, for iron volume concentrations of 10, 20, and 30 percent ....................................................................... 165 Figure 59: Shear stress of ideal Bingham plastic (and MR fluid model) versus shear rate 𝒅𝒗𝒅𝒏, compared to ideal Newtonian liquid ........................................................................................... 167 Figure 60: Microgravity time as a function of launch angle from horizon ................................ 170 Figure 61: Slosh regimes and similarity parameters .................................................................. 172 Figure 62. Schematic and free-body diagram of slosh dynamic model ...................................... 173 Figure 63. Base Plate .................................................................................................................. 179 Figure 64. Second and Top Plate ................................................................................................ 180 Figure 65. Side view of main structure ....................................................................................... 182 Figure 66. Trimetric view of main structure ............................................................................... 182 Figure 67. Top view of structure with 90 Degree L Brackets .................................................... 183




Figure 68. Side view of 90 Degree Brackets .............................................................................. 184 Figure 69. Test Structure with Base Plates ................................................................................. 185

Table of Tables

Table 1. Mission Objectives and Mission Success Criteria for the L.S.I.M. mission .................. 17 Table 2. Launch Vehicle RVM ..................................................................................................... 18 Table 3. Flight Systems RVM ...................................................................................................... 26 Table 4. Flight Avionics RVM ..................................................................................................... 29 Table 5: Mission Success Criteria................................................................................................. 33 Table 6: Launch Vehicle System Requirements ........................................................................... 34 Table 7: Launch Vehicles Properties ............................................................................................ 48 Table 8: Recovery System Properties ........................................................................................... 48 Table 9: Drift Estimates ................................................................................................................ 49 Table 10: Recovery Characteristics .............................................................................................. 49 Table 11: Kinetic Energy at Drogue Parachute Deployment ........................................................ 50 Table 12 : Kinetic Energy at Main Parachute Deployment .......................................................... 51 Table 13: Black Powder Properties............................................................................................... 52 Table 14: Black Powder Masses ................................................................................................... 52 Table 15: Success Criteria ............................................................................................................ 53 Table 16: Failure Modes ............................................................................................................... 53 Table 17: Altitude as a Function of Motor Selection (Constant Dry Mass) ................................. 57 Table 18: Overall Weight Breakdown .......................................................................................... 61 Table 19: Intimidator 5 Kit Landing Masses ................................................................................ 62 Table 20: Methods currently available for damping slosh. .......................................................... 65 Table 21: Elements of the theoretical modeling for the LSIM payload ...................................... 69 Table 22: LSIM success criteria from the Requirements Verification Matrix ............................ 70 Table 23: LSIM Requirements ..................................................................................................... 70 Table 24: Scientific method fulfillment for LSIM ....................................................................... 76 Table 25: Test sequences and descriptions, included options de-scoped since PDR .................. 77 Table 26: List of MR fluid ingredients ........................................................................................ 79 Table 27: Payload Assembly Dimensions ................................................................................. 100 Table 28: Data from SolidWorks SimulationXpress, highlighting the data from assumptions. 102 Table 29: Avionics requirements ............................................................................................... 105 Table 30: Avionics Success Criteria .......................................................................................... 107 Table 31. SIDES Power Budget.................................................................................................. 111 Table 32: Major Flight Computer Components ......................................................................... 115 Table 33: Ground station requirements ...................................................................................... 117 Table 34: Risk Identification and Mitigation Steps ................................................................... 130 Table 35: Risk Assessment Matrix with Risk Class ................................................................... 131 Table 36. Launch vehicle failure modes. ................................................................................... 131




Table 37. Payload hazards and mitigation ................................................................................. 132 Table 38. Payload safety failure modes ..................................................................................... 135 Table 39: Environmental Hazards, Risks, and Mitigation ......................................................... 137 Table 40. Summary of sponsors for the Ramblin. Rocketeers ................................................... 141 Table 41. List of current sponsors of the Ramblin' Rocketeers. ................................................. 142 Table 42. FRR Project Budget Summary. .................................................................................. 143 Table 43. Design milestones set by the USLI Program Office. .................................................. 147 Table 44. Identification and Mitigations for High-Risk Tasks. .................................................. 150 Table 45. Low to Moderate Risk items and mitigiations. ........................................................... 152 Table 46: Microgravity times for fall heights ............................................................................ 171 Table 47: Similarity parameters for simplified flight profile of the launch vehicle .................. 172




1. Introduction

School Information and NAR Section Contacts 1.1.

Team Summary Sc

hool

Info

& P

roje

ct

Title

School Name Georgia Institute of Technology

Team Name Ramblin’ Rocketeers

Project Title Liquid Stabilization in Microgravity

(LSIM)

Launch Vehicle Name Vespula Mk II

Payload Option 1,2 0F1

Team

Info

rmat

ion

Project Lead / Team

Official

Richard

Safety Officer Tony, Joseph

Team Advisors Dr. Eric Feron

Dr. Marilyn Wolf

NA

R

Info

rmat

ion

NAR Section Primary: Southern Area Rocketry

(SoAR) #571

Secondary: GA Tech Ramblin’

Launch vehicle Club #701

NAR Contacts Primary: Matthew Vildzius

Secondary: Jorge Blanco

1 The Ramblin’ Rocketeers’ LSIM payload is applicable to both the Option 1 and Option 2 payload options listed in the 2012-2013 USLI Handbook. On its own, the LSIM payload is intended to be an engineering payload demonstrating a novel technology; additionally, the LSIM payload can be scaled up and will be shown to meet the requirements to compete for the Option 2 payload option.




Work Breakdown Structure 1.2.

In order to effectively coordinate design efforts, the project is broken down along technical

discipline lines that emulate typical programs in the Aerospace industry. Each sub-team has a

general manager supported by several technical leads and subordinate members. Team

memberships were selected based on the individuals’ areas of expertise as well as personal

interest. Figure 1 shows the work breakdown structure.

Figure 1. 2012 – 2013 project work breakdown structure.




Launch Vehicle Summary 1.3.

1.3.1. Overview

The Ramblin’ Rocketeers’ launch vehicle has a gross-lift off weight of approximately 45 pounds

and features a 75 mm L1350 solid motor. The launch vehicle is an Intimidator 5 kit with a

custom payload integration structure. The recovery system utilizes a 30” drogue parachute

slowing the launch vehicle down to 64.78 ft/s and a 120” main parachute to slow the launch

vehicle down to 15.12 ft/s.

1.3.2. Changes since CDR

The following changes have been made since the Preliminary Design Review:

• Due to several launch failures the custom Vespula Mk II vehicle design has been de-

scoped to the Intimidator 5 kit.

Payload Summary 1.4.

1.4.1. Overview

The Ramblin’ Rocketeers will design, build, test, and fly a system for damping liquid slosh

through the use of magnetorheological fluid. This fluid will be actuated with solenoids and

driven to a pre-defined state in the Liquid Stabilization in Microgravity (LSIM) experiment.

Further, Flight Systems will implement a network of SIDESboards for distributed sensor

networks, empowering LSIM, and collecting valuable engineering data. A substantial ground

station for observation and telemetry is planned to support the flight of the launch vehicle.

Additionally, the Ramblin’ Rocketeers will pursue the NASA payload options 1 and 2 in the

design, construction, testing, and flight of a primary science experiment and Reduced Gravity

Education Flight Program. This payload will test the feasibility and practicality of systems to

manipulate magnetorheological (MR) fluids in microgravity for the purpose of demonstrating

possible methods for reducing propellant slosh in low-gravity environments.




1.4.2. Changes Since CDR

1.4.2.1. Payload Changes Since CDR

• All use of cameras for the science payload has been de-scoped.

1.4.2.2. Avionics Changes Since CDR

• Temperature and Strain Gauge nodes no longer necessary with switch from Vespulla MkII to Intimidator kit

• MasterIMU now uses a SIDESboard instead of a Maple, extra computing power not necessary because only two nodes use the SIDES network

• Master Clock node has been descoped, since there are only two nodes to synchronize • RS485 Hardware will use simplified control software reflecting the simplifications to the

SIDES network

Infrared will no longer be used as a primary means of measuring slosh in the experiment for the

launch vehicle. A camera independent of the avionics apparatus may be used – however the

primary sensor is seen to be a vibration sensor placed into the base bolt and integrated into a

SIDES node. The precise details of ground testing have been reviewed in depth and many

changes as to the specifics have been made as testing platforms have been developed. These

should result in high quality ground testing data. This data will be used to complete the final link

in an expanded theory describing MR fluid. While operating under several assumptions and

simplifications, this expanded theory should aid greatly in the development of control software

for the flight experiment.




2. Project L.S.I.M. Overview

Mission Statement 2.1.

The mission of the Mile High Yellow Jackets is:

To maintain a sustainable team dedicated to the gaining of knowledge through the designing,

building, and launching of reusable launch vehicles with innovative payloads in accordance with

the NASA University Student Launch Initiative Guidelines.

Requirements Flow Down 2.2.

The requirements flow down is illustrated in Figure 2. As illustrated by the requirements flow

down, the Mission Success Criteria flow down from the Mission Objectives of Project A.P.E.S.

All system and sub-system level requirements flow down from the either of the Mission

Objectives, Mission Success Criteria, or the USLI Handbook.




Mission Objectives and Mission Success Criteria 2.3.

Table 1. Mission Objectives and Mission Success Criteria for the L.S.I.M. mission

MO Mission Objective

MO-1 An altitude of 5,280 ft above the ground is achieved.

MO-2 Create an environment in which to test microgravity payloads.

MO-3 Reduction in the sloshing motion of a propellant simulatn in microgravity with a magnetic

fluid.

MO-4 Successful recovery of the launch vehicle resulting in no damage to the launch vehicle.

MSC Mission Success Criteria Source Verification

Method

MSC-1 Minimum Mission Succes: Achieve an altitude of

5,280 ft., with a tolerance of +320 ft./-640 ft. MO-1

Testing

MSC-2 Minimum Mission Succes: Achieve a microgravitiy

environment of ± 0.1 G MO-2

Testing

MSC-3 Minimum Mission Success:Sucessfully record video

of flight experiment during microgravity and start/stop

the experiment without mechanical and electrical

failures.

MO-3

Testing

MSC-4 Full Mission Succes: Successful matching of the

damping ratio for ringed baffles in the wave amplitudes

experienced during flight to within ±30%.

MO-3

Testing

MSC-5 Minimum Mission Success: The Launch Vehicle is

recovered with no damage to the structure of the launch

vehicle.

MO-4,USLI

Handbook 1.4

Testing

MSC-5.1 Full Mission Succes:The Launch Vehicle is recovered

with no damage to the skin of the launch vehicle.

MSC-7, MO-4 Testing

System Requirements Verification Matrix (RVM) 2.1.

Table 2, Table 3, and Table 4 list the requirements verification matrix for each subsystem.




Table 2. Launch Vehicle RVM

Requirement

No. Requirement Source

Verification

Method

Design

Feature Status

LV-1

The Launch Vehicle

shall carry a

scientific or

engineering

payload.

USLI Handbook 1.1,

MO-2 Inspection

iMPS

standardized

payload

interface

In Progress

LV-1.1

The maximum

payload weight

including any

supporting avionics

shall not exceed 15

lbs.

LV-1 Inspection

Maximum

Parachute

Sizing

In Progress

LV-1.2

The Launch Vehicle

shall have a

maximum of four

(4) independent or

tethered sections

USLI Handbook 1.5 Inspection

Three (3)

sections:

nosecone,

payload, and

booster

In Progress

LV-2

The Launch Vehicle

shall carry the

payload to an

altitude of 5,280 ft.

above the ground.

USLI Handbook 1.1,

MO-1 Testing

Modified tube

fins for

straight flight,

motor sizing

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-2.1

The Launch Vehicle

shall use a

commercially

available solid

motor using

ammonium

perchlorate

composite

propellant (APCP).

USLI Handbook

1.11 Inspection

Use of a

commercially

available

solid motor

In Progress

LV-2.2

The total impulse

provided by the

Launch Vehicle

shall not exceed

5,120 N-s.

USLI Handbook

1.12 Inspection

A motor with

a maximum

motor class of

"L" shall be

used

In Progress

LV-2.3

The Launch Vehicle

shall remain

subsonic throughout

the entire flight.

USLI Handbook 1.3 Analysis Motor Sizing In Progress

LV-2.4

The Launch Vehicle

shall carry one

commercially

available barometric

altimeter for

recording of the

official altitude


Commercially

available

altimeter

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-2.5

The amount of

ballast, in the

vehicle's final

configuration that

will be flown in

Huntsville, shall be

no more than 10%

of the unballasted

vehicle mass.

USLI Handbook

1.14 Inspection

Proper motor

selection for

gross lift-off

weight of the

launch

vehicle.

In Progress

LV-2.5

The Launch Vehicle

shall have

aerodynamic

stability margin of

1.5 to 3 cailbers

prior to leaving the

launch rail.

LV-2 Analysis

Modified

tube-fins for

aerodynamic

stabilization.

In Progress

LV-3

The Launch Vehicle

shall be safely

recovered and be

reusable.

MSC-7.1 Testing

Parachute

Sizing and

real time

Ground

Station

tracking

In Progress

LV-3.1

The Launch Vehicle

shall contain

redundant

altimeters.


Ground

testing of

altimeter

ejection.

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-3.2

The recovery

system shall be

designed to be

armed on the pad.

LV-3 Inspection Arming

Switches In Progress

LV-3.3

The recovery

system electronics

shall be completely

independent of the

payload electronics.


The recovery

system

electronics

shall be

entirely

independent

of from all

other systems.

In Progress

LV-3.4

Each altimeter shall

be armed by a

dedicated arming

switch which is

accessible from the

exterior of the

vehicle airframe

when the vehicle is

in the launch

configuration on the

launch pad.


Recovery

system design

shall

incorporate

one (1)

independent

arming switch

for each

altimeter

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-3.5


have a dedicated

power supply.


Recovery

system design

shall

incorporate

independent

power

supplies for

each

altimeter.

In Progress

LV-3.6

Each arming switch

shall be capable of

being locked in the

"ON" position for

launch.

USLI Handbook 2.8 Testing

The arming

switches will

be designed to

use a key to

change the

state of the

switch.

In Progress

LV-3.7

Each arming switch

shall be a maximum

of six (6) feet above

the base of the

Launch Vehicle.


Arming

switches shall

be located

near the

booster

section of the

launch

vehicle

In Progress

LV-3.8

The Launch Vehicle

shall utilize a dual

deployment

recovery system.


Utilization of

a drogue and

main

parachute

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-3.9

Removable shear

pins shall be used

for both the main

and drogue

parachute

compartments

USLI Handbook

2.10 Inspection

Plastic shear

pins will be

installed in

the recovery

compartments

.

In Progress

LV-3.10

All sections shall be

designed to recover

within 2,500 ft. of

the launch pad

assuming 15 MPH

winds.

USLI Handbook 2.3 Analysis

Parachute

sizing will

incorporate

descending

velocities and

drift

restrictions.

In Progress

LV-3.11

Each section of the

Launch Vehicle

shall have a

maximum landing

kinetic energy of 75

ft-lbf.


Properly sized

main

parachute to

ensure

landing

kinetic

energies

below 75 ft.-

lbf

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-3.12

The recovery

system electronics

shall be shielded

from all onboard

transmitting

devices.

LV-3 Testing

Proper

shielding

shall be

incorporated

into the

design to

protect the

electronics

from payload

interference.

In Progress

LV-4

The Launch Vehicle

shall be launched

utilizing

standardized launch

equipment

LV-3 Inspection

Use of

standard 1515

rail buttons

and 8 foot

launch pad

rail.

In Progress

LV-4.1

The Launch Vehicle

shall be capable of

being launched by a

standard 12 volt

direct current (DC)

firing system and

shall require no

external circuitry or

special ground

support equipment

to initial launch.


Use of

standard

igniters.

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-4.2

The Launch Vehicle

shall not require any


special ground

support equipment

to initiate the launch

other than what is

provided by the

range.

USLI Handbook

1.10 Testing

Use of

standard

igniters, 1515

rail buttons,

and 8 foot

launch rail.

In Progress

LV-4.4

The Launch Vehicle

shall have a pad stay

time on one (1)

hour.


Follow

manufacturers

recommendati

ons for power

In Progress

LV-4.5

The Launch Vehicle

shall be capable of

being prepared for

flight at the launch

site within two (2)

hours from the time

the waiver opens.


Easy

assembly of

the rocket

structure and

easy

integration of

the payload

and avionics.

In Progress

LV-4.6

The Launch Vehicle

shall be compatible

with either an 8 foot

long, 1 in. rail

(1010), or an 8 feet

foot long, 1.5 in. rail

(1515), provided by

the range.


Utilization of

1515 rail and

rail interfaces

for launch

In Progress




Table 3. Flight Systems RVM

Requirement

Number

Requirement Definition Source Verification

Method

Design

Feature

Status Verification

Source

Document

FS-1 The flight systems team

shall design and build

the LSIM Payload MO-3 Inspection LSIM

payload In

Progress MO-3

FS-2 The LSIM payload shall

be designed to fly on a

SLP rocket

USLI Handbook

3.1.1 Inspection LSIM

payload In

Progress

USLI Handbook

3.1.1

FS-4 The Flight Systems

Team shall produce a

working system for

manipulating MR fluid

in LSIM.

MSC-3 Testing Solenoids

and Control Algorithms

In Progress MSC-3


Team shall ensure that

all avionics are properly

shielded from the LSIM

payload.

MSC-3 Testing

Faraday cages and webbing tied to

ground on the harness

Not Started MSC-3


Team shall design all

LSIM components and

avionics such that they

may be easily integrated

with the Modular

Payload System of the

payload bay in the

rocket.

MSC-3 Inspection Mounting system Complete MSC-3




Requirement

Number


Method

Design

Feature

Status Verification

Source

Document


Team shall conform to

all weight, power, and

dimensional

requirements as per the

rocket design.

MSC-3 Analysis TBD In Progress MSC-3

FS-7.1 The Experiment and

Avionics, with

mechanical supports,

shall weight no more

than 15 lbf.

LV-1.1 Inspection TBD In Progress LV-1.1

FS-8 The flight computer

shall execute all tasks

necessary to the

operation of the LSIM

payload and avionics.

MSC-3 Inspection Maple SIDES node

In Progress MSC-3

FS-9 The LSIM payload shall

have a dedicated power

supply. MSC-3 Inspection SIDES node In

Progress MSC-3


Team shall ensure

redundancy and

reliability of all internal

electrical hardware.

MSC-3 Inspection SIDES network

In Progress MSC-3




Requirement

Number


Method

Design

Feature

Status Verification

Source

Document


Team shall provide for

payload operation with

up to 1 hour of wait on

the launch pad and 2

hours of wait during

preparation of the

Rocket.

USLI Handbook

1.6 Inspection TBD In

Progress

USLI Handbook

1.6


Team shall provide for

electrical operations to

begin at the beginning

of the flight trajectory.

MSC-3 Inspection TBD In Progress MSC-3


Team shall ensure that

the LSIM payload is

shut down safely during

the deployment phase of

the flight trajectory.


FS-14 Data from the LSIM

payload shall be

collected, analyzed, and

reported by the team

using the scientific

method.

USLI Handbook

3.2 Inspection

Data logging in

SIDES network

In Progress

USLI Handbook

3.2




Requirement

Number


Method

Design

Feature

Status Verification

Source

Document

FS-15 The LSIM payload will

be designed to be

recoverable and be able

to launch again on the

same day without any

repairs or modifications.

USLI Handbook

3.5 Inspection

Appropriate mounting

to the payload

interface.

In Progress

USLI Handbook

3.5

Table 4. Flight Avionics RVM

Requirement


Verification

Method

Design

Feature Status

FA-1 All Flight Avionics shall

have sufficient power

sources to survive 1-hour

pad stay in additon to

normal operation

requirements

USLI Handbook 1.7 Testing Power Supply In Progress

FA-2 The Flight Computer shall

collect video of the flight

experiment during

microgravity

MSC-3 Testing Camera In Progress

FA-3 The Flight Computer shall

collect Launch Vehicle

position data and

environment conditions (e.g.

acceleration).

MO-4 Testing IMU, GPS In Progress




Requirement


Verification

Method

Design

Feature Status

FA-4 The Flight Avionics shall

downlink telemetry

necessary to a Ground

Station for the recovery of

the Launch Vehicle

USLI Handbook

2.11 Teting

GPS, Ground

Station, Xbee In Progress

FA-5 The GPS coordinates of all

independent Launch Vehicle

sections shall be transmitted

to the Ground Station

USLI Handbook

2.11.1 Teting

GPS, Ground

Station, Xbee In Progress

FA-6 The Flight Avionics shall

operate on an independent

power supply from the

recovery system.

USLI Handbook

2.12 Inspection Power Supply In Progress

Mission Profile 2.1.

Figure 3 illustrates the mission profile for Project L.S.I.M. In order to achieve the desired

microgravity environment, the launch vehicle will continue through for one (1) second until

deployment of the drogue parachute. This post-apogee delay will yield approximate 4.5 seconds

of microgravity to perform the L.S.I.M.

G 2012 – 2013 GEORGIA TECH RAMBLIN’ ROCKETEERS FLIGHT READINESS REVIEW


Figure 3. Project L.S.I.M. mission profile.




3. Launch Vehicle

Overview 3.1.

The purpose of the launch vehicle is to carry a scientific payload to one mile in altitude and

safely return the vehicle to the surface of the Earth. Embracing innovative and out-of-the-box

thinking, the Ramblin’ Rocketeers launch vehicle will have the ability to carry a wide range of

payloads, from scientific experiments to engineering flight demonstrations. A rib and stringer

payload mounting rig enables easy integration for payload. The launch vehicle has a five inch

outer diameter and is 9 feet, 10 inches in length. The launch vehicle is composed of three

sections; the nose cone, the payload section, and the booster section. The science payload will be

housed in the payload section of the rocket and the avionics will be housed in the booster section

above the motor.

The launch vehicle will utilize a dual-deployment recovery system that will minimize the drift of

the launch vehicle by mitigating the effects of unpredictable wind conditions with a drogue chute

descent. However, the overall purpose of the recovery system, to minimize damage to the launch

vehicle from impact with the ground, will be maintained by a main chute deployed closer to the

ground. The drogue parachute will be housed in the section connecting the booster and payload

sections, while the main parachute will be located between the payload section and nose cone.

Both parachutes are made of rip-stop nylon. To ensure successful chute deployment, redundant

systems will be used. Each chute will feature two independent black powder ejection charges

with corresponding redundant igniters and StratoLogger altimeters. The powder charges will be

ignited using low-current electronic matches with independent power supplies at the command of

the altimeters.

3.1.1. Mission Criteria

The criteria for mission success are shown in Table 3.




Table 5: Mission Success Criteria

Requirement Design feature to satisfy that

requirement

Requirement

Verification Success Criteria

Provide a suitable

environment for the

payload.

The payload requires a steady, but

randomly vibrating platform to

test the L.S.I.M. system.

Unsteadiness in the motor's thrust

and launch vehicle aerodynamics

cause vibrations. In addition,

deployment of the drogue

parachute will be delayed one

second to maximize time in

microgravity.

By measuring the

acceleration with the

payload's

accelerometers.

The L.S.I.M.

system reduces a

recordable amount

of sloshing.

To fly as close to a

mile in altitude as

possible without

exceeding 5,600 ft.

A motor will be chosen to propel

the vehicle to a mile in altitude.

Through the use of

barometric

altimeters.

The altimeters

record an altitude

less than 5,600 ft.

The vehicle must be

reusable.

The structure will be robust

enough to handle any loading

encountered during the flight.

Through finite

element analyses

and structural

ground testing of

components.

The vehicle

survives the flight

with no damage.

System Design Overview 3.2.

lists the derived system-level requirements in order to meet the success criteria. The requirement

numbers reference the requirements in the 2012-2013 USLI Handbook.




Table 6: Launch Vehicle System Requirements

Requirement


Verification

Method

Design

Feature Status

LV-1

The Launch Vehicle

shall carry a

scientific or

engineering

payload.

USLI Handbook 1.1,

MO-2 Inspection

iMPS

standardized

payload

interface

In Progress

LV-1.1

The maximum

payload weight

including any

supporting avionics

shall not exceed 15

lbs.

LV-1 Inspection

Maximum

Parachute

Sizing

In Progress

LV-1.2

The Launch Vehicle

shall have a

maximum of four

(4) independent or

tethered sections


Three (3)

sections:

nosecone,

payload, and

booster

In Progress

LV-2

The Launch Vehicle

shall carry the

payload to an

altitude of 5,280 ft.

above the ground.

USLI Handbook 1.1,

MO-1 Testing

Modified tube

fins for

straight flight,

motor sizing

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-2.1

The Launch Vehicle

shall use a

commercially

available solid

motor using

ammonium

perchlorate

composite

propellant (APCP).

USLI Handbook

1.11 Inspection

Use of a

commercially

available

solid motor

In Progress

LV-2.2

The total impulse

provided by the

Launch Vehicle

shall not exceed

5,120 N-s.

USLI Handbook

1.12 Inspection

A motor with

a maximum

motor class of

"L" shall be

used

In Progress

LV-2.3

The Launch Vehicle

shall remain

subsonic throughout

the entire flight.

USLI Handbook 1.3 Analysis Motor Sizing In Progress

LV-2.4

The Launch Vehicle

shall carry one

commercially

available barometric

altimeter for

recording of the

official altitude


Commercially

available

altimeter

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-2.5

The amount of

ballast, in the

vehicle's final

configuration that

will be flown in

Huntsville, shall be

no more than 10%

of the unballasted

vehicle mass.

USLI Handbook

1.14 Inspection

Proper motor

selection for

gross lift-off

weight of the

launch

vehicle.

In Progress

LV-2.5

The Launch Vehicle

shall have

aerodynamic

stability margin of

1.5 to 3 cailbers

prior to leaving the

launch rail.

LV-2 Analysis

Modified

tube-fins for

aerodynamic

stabilization.

In Progress

LV-3

The Launch Vehicle

shall be safely

recovered and be

reusable.

MSC-7.1 Testing

Parachute

Sizing and

real time

Ground

Station

tracking

In Progress

LV-3.1

The Launch Vehicle

shall contain

redundant

altimeters.


Ground

testing of

altimeter

ejection.

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-3.2

The recovery

system shall be

designed to be

armed on the pad.

LV-3 Inspection Arming

Switches In Progress

LV-3.3

The recovery

system electronics

shall be completely

independent of the

payload electronics.


The recovery

system

electronics

shall be

entirely

independent

of from all

other systems.

In Progress

LV-3.4


be armed by a

dedicated arming

switch which is

accessible from the

exterior of the

vehicle airframe

when the vehicle is

in the launch

configuration on the

launch pad.


Recovery

system design

shall

incorporate

one (1)

independent

arming switch

for each

altimeter

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-3.5


have a dedicated

power supply.


Recovery

system design

shall

incorporate

independent

power

supplies for

each

altimeter.

In Progress

LV-3.6

Each arming switch

shall be capable of

being locked in the

"ON" position for

launch.


The arming

switches will

be designed to

use a key to

change the

state of the

switch.

In Progress

LV-3.7

Each arming switch

shall be a maximum

of six (6) feet above

the base of the

Launch Vehicle.


Arming

switches shall

be located

near the

booster

section of the

launch

vehicle

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-3.8

The Launch Vehicle

shall utilize a dual

deployment

recovery system.


Utilization of

a drogue and

main

parachute

In Progress

LV-3.9

Removable shear

pins shall be used

for both the main

and drogue

parachute

compartments

USLI Handbook

2.10 Inspection

Plastic shear

pins will be

installed in

the recovery

compartments

.

In Progress

LV-3.10

All sections shall be

designed to recover

within 2,500 ft. of

the launch pad

assuming 15 MPH

winds.


Parachute

sizing will

incorporate

descending

velocities and

drift

restrictions.

In Progress

LV-3.11

Each section of the

Launch Vehicle

shall have a

maximum landing

kinetic energy of 75

ft-lbf.


Properly sized

main

parachute to

ensure

landing

kinetic

energies

below 75 ft.-

lbf

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-3.12

The recovery

system electronics

shall be shielded

from all onboard

transmitting

devices.

LV-3 Testing

Proper

shielding

shall be

incorporated

into the

design to

protect the

electronics

from payload

interference.

In Progress

LV-4

The Launch Vehicle

shall be launched

utilizing

standardized launch

equipment

LV-3 Inspection

Use of

standard 1515

rail buttons

and 8 foot

launch pad

rail.

In Progress

LV-4.1

The Launch Vehicle

shall be capable of

being launched by a

standard 12 volt

direct current (DC)

firing system and

shall require no


special ground

support equipment

to initial launch.


Use of

standard

igniters.

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-4.2

The Launch Vehicle

shall not require any


special ground

support equipment

to initiate the launch

other than what is

provided by the

range.

USLI Handbook

1.10 Testing

Use of

standard

igniters, 1515

rail buttons,

and 8 foot

launch rail.

In Progress

LV-4.4

The Launch Vehicle

shall have a pad stay

time on one (1)

hour.


Follow

manufacturers

recommendati

ons for power

In Progress

LV-4.5

The Launch Vehicle

shall be capable of

being prepared for

flight at the launch

site within two (2)

hours from the time

the waiver opens.


Easy

assembly of

the rocket

structure and

easy

integration of

the payload

and avionics.

In Progress




Requirement


Verification

Method

Design

Feature Status

LV-4.6

The Launch Vehicle

shall be compatible

with either an 8 foot

long, 1 in. rail

(1010), or an 8 feet

foot long, 1.5 in. rail

(1515), provided by

the range.


Utilization of

1515 rail and

rail interfaces

for launch

In Progress

Recovery System 3.3.

The purpose of the recovery system is to minimize damage to the launch vehicle from impact

with the ground. The launch vehicle will use a dual-deployment recovery system to mitigate the

effects of unpredictable wind conditions on drift with a drogue chute descent. The drogue

parachute will be housed in the compartment connecting the booster and payload sections, and

the main parachute will be located between the payload section and nose cone, as illustrated

below in Figure 4. The launch vehicle will be armed on the launch pad using two arming

switches, one for each independent altimeter and ejection charge. For the purpose of simulation,

the launch vehicle has been modeled using the Open Rocket Software, with both parachutes

made of rip-stop nylon.




Figure 4: Internal Layout of the Launch Vehicle

During descent, 30 feet of Kevlar webbing will connect the parachutes to the launch vehicle. The

drogue parachute will be housed in a cylindrical compartment in the rear section between the

payload and booster sections as illustrated in Figure 4. This compartment has an outer diameter

of 5.25 inches and a length of 10 inches. A bulkhead in the rear payload section will house the

ejection wells and also serve to take the impulse of the gun powder blast. The drogue parachute’s

retention mechanics includes a U-Bolt placed between the two ejection wells on the underside of

the payload section, as well as a U-Bolt in the booster section thrust plate. In addition, a shock

cord connecting the booster section and main rocket body together. At deployment, the ejection

charges will separate the booster section from the main rocket, releasing the drogue parachute as

well.




Figure 5: Drogue Parachute Assembly

The main parachute will be placed in a section above the payload bay. The section has an outer

diameter of 5.25 inches and a length of 12 inches. The main parachute’s ejection wells will be

placed such that the impulse is imparted on the payload section and the nose cone is separated

from the main rocket – pulling the main parachute out. Shock cords will connect the main

parachute to the nose cone and the payload section of the launch vehicle, ensuring that the all

sections remain together during descent. The main parachute assembly is illustrated in Figure 6.




Figure 6: Main Parachute Assembly

The parachute casings are made of G10 fiberglass, and the bulkhead under the main chute is

made of plywood. Two-inch stainless steel U-Bolts will be drilled into the bulkheads, and will be

used to attach the shock cords. 1/16” nylon rod will be used as the four shear pins to keep both

the main and drogue chute compartments together during flight until the parachutes are

deployed. PVC end-caps will be used to direct the ejection charges in order to protect the casing

from thermal shock, and a NOMEX shield will protect the parachutes. The charges will be

ignited using an e-match.

3.3.1. Altimeters

To ensure successful chute deployment, redundant systems will be used. Each chute will feature

two independent black powder ejection charges with corresponding redundant igniters and

StratoLogger altimeters. The altimeters will ignite the ejection charges through the use of low-

current electronic matches using independent power supplies. The components which compose




each altimeter system are independent of all payload electronics. The altimeters and all recovery

electronics have a pad stay time of at least an hour. The system setup for each altimeter is shown

below in Error! Reference source not found.. The electrical drawing for the wiring harness is

located in Appendix VI.

Figure 7: Electronic Altimeter Schematic

In addition, the recovery electronics wiring will be protected from transmitting devices in the

rocket through faraday cages and shielding integrated into the wiring harnesses, these devices are

discussed further in the Avionics. Ground testing will determine whether transmission

interference will affect the altimeter devices directly.

3.3.2. Arming Switches

The altimeters and the recovery systems will be activated on the launch pad with two arming

switches. Each arming switch activates one of the two independent altimeter systems. The

arming switches will be located at the base of the payload section which is approximately four

feet above the bottom of the launch vehicle. The arming switches will be Featherweight Screw

Switches and is illustrated below in Figure 8. The Screw Switches are locked in the ON position

when the middle screw is screwed in and completes the circuit. In addition to having a simple

activation and de-activation method, the Screw Switches are very lightweight and small.




Figure 8: Featherweight Screw Switches

3.3.3. Parachute Dimensions

The sizing of the main parachute is determined by the weight of the launch vehicle and the

kinetic energy constraint of the launch vehicle when it touches down. Based on the LV-3.10 and

LV-3.11 requirements, the launch vehicle should not experience more than 75.0 ft-lbf of kinetic

energy upon landing, this places an upper limit on the landing velocity to be approximately 15.12

ft/s. The main parachute is 12 feet and the drogue parachute is two feet.

Table 7 and Table 8 outline the dimensions and properties of the constraining launch vehicle

properties and the properties of the parachutes.




Table 7: Launch Vehicles Properties

Launch Vehicle Properties

Weight of launch vehicle 45 lb

CD of Launch vehicle 0.75

Max Kinetic Energy 75.0 ft-lbf

Table 8: Recovery System Properties

Properties Main Parachute Drogue Parachute

Diameter (ft) 12 2

Surface Area (ft^2) 113.10 3.14

Estimated CD 1.40 1.20

Target Descent Rate (ft/s) 13.75 89.11

3.3.4. Drift Profile Analysis

Drift profile analysis is the method used to estimate and constrain the landing site for the launch

vehicle. Based on how long the launch vehicle will be in flight and the wind speed at launch, the

range can be estimated. Using the equations below, the drift of the launch vehicle under the main

and drogue parachutes can be determined. The results are shown below in Table 9

𝐷𝑟𝑖𝑓𝑡 = 𝑇𝑖𝑚𝑒𝑖𝑛 𝑓𝑙𝑖𝑔ℎ𝑡 ∗ 𝑉𝑤𝑖𝑛𝑑 (1)

𝑇𝑖𝑚𝑒𝑖𝑛 𝑓𝑙𝑖𝑔ℎ𝑡 =

𝐴𝑙𝑡𝑚𝑎𝑥

𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑠𝑝𝑒𝑒𝑑 (2)

𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦 = �

2𝑚𝑔𝜌𝐴𝐶𝑑

(3)




Table 9: Drift Estimates

Launch Vehicle Drift Estimates

Wind Speed (mph)

Drift (ft)

Drogue Parachute Main Parachute Total Drift 5 393.35 266.65 660.01

10 786.70 533.31 1320.01

15 1180.06 799.96 1980.02

20 2018.18 1216.11 3234.29

The descent velocity of the launch vehicle will be estimated using the terminal velocity. The

terminal velocity is the constant speed of a free-falling object when the drag due to air resistance

prevents further acceleration. The values are listed below in Table 10.

Table 10: Recovery Characteristics

Recovery Systems Properties Drogue Parachute Main Parachute

Diameter (ft) 2.00 Dimensions (ft) 12.00

Flight Time (s) 53.64 Flight Time (s) 36.36

Terminal

Velocity (ft/s) 89.11

Terminal

Velocity (ft/s) 13.75

Horizontal

Drift (ft) 1180.06

Horizontal

Drift (ft) 799.96

Total Drift (ft) 1980.02




3.3.5. Kinetic Energy of Launch Vehicle

Kinetic energy calculations were performed using the equation below.

𝐾𝐸 =12𝑚𝑣2 (4)

Using the masses of the separate sections, the kinetic energies can be calculated using the

velocity of the system at different points in the mission. The Kinetic Energies of separate

sections after the drogue chute is deployed are given below in Table 11. After the drogue chute is

deployed, the launch vehicle has separated only between the booster section and the payload

section, so the payload and nosecone sections are treated as one part. The velocities listed are the

terminal velocities under the drogue parachute once the

Table 11: Kinetic Energy at Drogue Parachute Deployment

Launch Vehicle Section Weight (lb.) Velocity (ft/s) Kinetic Energy (ft-lbf)

Nose Cone 1.59 89.11 196.07

Payload 16.45 89.11 2028.51

Booster 19.61 89.11 2418.18

The Kinetic Energies of the separate sections after the deployment of the main chute and landing

are given below in Table 12. After the main chute is deployed all three sections have separated

and their separate masses were used in the calculations. All sections will have the same velocity

due to the shock cord tethers.




Table 12 : Kinetic Energy at Main Parachute Deployment

Launch Vehicle Section Weight (lb.) Velocity (ft/s) Kinetic Energy (ft-lbf)

Nose Cone 1.59 13.75 4.67

Payload 16.45 13.75 48.30

Booster 19.61 13.75 57.58

3.3.6. Ejection Charges

To eject the parachutes, redundant black powder charges will be used. The containers housing

the chutes will also be pressurized in order to ensure chute deployment. Due to the different

requirements for the drogue and main chutes, two sets of calculations will be needed.

The amount of black powder used in the ejections charges can be calculated through Equation

(5) below. Once the amount of black powder is determined the values can then be tested before

flight. The equation relates weight of black powder to the ejection pressure, volume of the

container, black powder combustion gas constant, and the black powder combustion temperature.

The constants used are listed below in Table 13.

𝑙𝑏 𝑜𝑓 𝐵𝑙𝑎𝑐𝑘 𝑃𝑜𝑤𝑑𝑒𝑟 =

𝑃𝑟𝑒𝑠𝑠𝑢𝑟𝑒 ∗ 𝑉𝑜𝑙𝑢𝑚𝑒𝑅𝑇

(5)

Using the pressurization of 10 psig and 9 psig as a structural maximum for the main and drogue

chute compartments, the resulting black powder masses are calculated to be 5 grams and 2 grams

for the main and drogue chutes, respectively, as illustrated below in Table 14. The masses used

will depend on the final container dimensions, which were estimated at 5.25 inches in radius and

12 and 10 inches in length for the main and drogue, respectively. The force required for

separation with the given number of Nylon shear pins would be 446 lbf for the main chute and

393 lbf for the drogue chute.




Table 13: Black Powder Properties

Constant Value

Combustion Gas

Constant 22.16 ft lbf/ lbm °R

Combustion

Temperature 3307 °R

Table 14: Black Powder Masses

Main Drogue

Total

Pressurization 10 psig 9 psig

Ejection force 446lbf 393lbf

Black Powder 5 grams 2 gram

3.3.7. Testing

In order to ensure the safety and viability of the calculations made in determining the black

powder masses, ground testing was completing before flying the launch vehicle recovery system.

The black powder testing was successfully conducted on the Vespula Mk II rocket. Since the

recovery sections of the Intimidator 5 kit and the Vespula Mk II are identical, the recovery

system of the Intimidator 5 kit is validated.

Due to the explosive nature of black powder charge testing, the tests for this launch vehicle were

coordinated with the campus security and the Georgia Tech Fire Marshal. For the black powder

test, the rocket was placed horizontally on the ground on a relatively smooth surface to minimize

unwanted static friction irrelevant to a flight environment. Table 15 lists the conditions for test

success and failure.




Table 15: Success Criteria

Success Criteria

Ejection charge ignites

Shear pins break

Launch vehicle moves half the distance of shock cord

Table 16: Failure Modes

Failure Criteria

The fiberglass of the tube coupler shatters due to the charge.

The shear pins don’t shear, and the launch vehicle stays intact.

The NOMEX/cloth shield fails and the parachute is burned.

The E-matches fail to ignite the black powder.

Structure 3.4.

The purpose of the launch vehicle is to carry a microgravity research payload to a mile in altitude

and safely return to the surface of the Earth. Additionally, the launch vehicle will also be

designed to carry a wide range of possible experiments, so that the rocket can be reused in the

future. The overall design is to be as flexible as possible, encouraging reuse for future research

and multiple launches. The rocket has been constructed and is ready for a test launch scheduled

for Saturday, March 23rd. The objective of the test launch is to verify the recovery system with

delayed apogee ejection and to collect preliminary acceleration data from the RGEFP payload.

The launch vehicle is a Performance Rocketry 5 inch Intimidator kit. The construction was

carried out by the Rocket Team, under the Supervision of Richard Zappulla, an experienced and

certified Level 2 High-Powered Rocket flyer. The payload is designed around the constraints of

the vehicle payload section, which are five inches in diameter and 30 inches in length.




3.4.1. Construction

The launch vehicle has a 98mm motor mount and adapters for both a 54mm and 75 mm motor

have been constructed to allow for a wide range of useable motors. The kit is made of all

fiberglass for strength and durability. The tubing is 1/16 in. thick, the centering rings and

bulkheads are 1/8 in. thick, the fins are 3/16 in. thick and factory beveled, and the nosecone is

constructed of fiberglass. The fins are attached with through-wall construction and securely

bonded to the motor mount and body tube with fiberglass-reinforced epoxy. The fins transmit

most of the force from the motor to the booster section, so the centering rings are not critical,

however they are bonded using the same techniques used for the fins. The fiberglass, at all

fiberglass-epoxy joints, was scuffed with 80 grit sandpaper to create a good bonding surface. US

Composites epoxy is reinforced with chopped (1/4 in.) or milled (1/16 in.) fiberglass filler and

fumed silica filler was used for all structural bonds on the rocket.

The recovery system attaches to the booster section with a length of 1/4 in. steel cable (wire

rope) rated for 6000 lb. breaking strength. Steel cable was used instead of a U-bolt because of

the small clearance between the 98mm motor mount and the 5 in. body tube. A loop of cable

extends through the centering ring, and a 1 foot section on either side is epoxied to the motor

mount. This method has been tested on other rockets and is at least as strong as the U-bolt bolted

to the centering ring. The long length of cable allows the force of the recovery system to be

distributed over a very large area of the motor mount tube compared to a relatively small area of

a centering ring. For a redundant recovery system attachment and a redundant motor retention, a

tapped forward closure will be used in addition to the cable attachment point. The rocket was not

built with a "ziperless" design because of the limitations that it would impose including motor

length and payload volume. A zipper is unlikely because the tubing is all fiberglass.

The attachment points to the payload bay use a more traditional U-bolt attachment. The U-bolts

are bolted through two doubled up 1/8 in. bulkheads for a total of 1/4 in. of fiberglass, and steel

fender washers are used on the back to further spread out the load from the recovery system.




The bulkhead assembly will be bolted to the end structural plates of the MPS assembly. All U-

bolts are 1/4 in. steel rated for over 2000 lb. breaking strength. The shock cord is one inch nylon

webbing rated for 4000 lb. and will be attached to the other components of the recovery system

with 1/4 in. steel quick links rated for at least 2000 lb. A length of shock cord is epoxied directly

to the nose cone as opposed to sealing it with a bulkhead to allow the addition of weight for final

adjustment, and to allow the installation of electronics in future missions.

3.4.2. Payload Integration

The payload will be integrated using a rib and stringer design. The stringers will be three #8

threaded rods and the ribs will be made of plywood. The ribs will be secured using #8-32 nuts

and #8 washers on both the top and bottom of each rib. This flexible design allows for the ribs to

be moved based on the needs of the science payload. At the top and bottom of the payload

integration structure are two half inch plywood bulkheads. The bulkheads are attached to the

body of the rocket using four L-brackets on both the top and bottom of each bulkhead. There is a

total of 16 L-brackets on the payload section bulkheads. The payload integration structure is

pictured below in Figure 9 with L.S.I.M. specific integration already components mounted.

Figure 9: Payload Integration Structure

3.4.3. Avionics Integration

Due to the relatively short payload section of the Intimidator 5 kit the avionics will be mounted

in the booster section above the motor. Similarly to the payload integration structure, the

avionics integration structure is composed of a bulkhead and a rib separated by three #8 threaded




rods. The bulkhead is secured by 8 L-brackets, four on the top and four on the bottom, and

features a steel U-bolt to secure the drogue parachute. The rib is secured using four L-brackets

on the bottom side. The avionics integration structure is pictured below in Figure 10.

Figure 10: Avionics Integration Structure

3.4.4. Section Integration

The three sections of the rocket, namely the nose cone, payload, and booster sections, will be

separated by two parachute bays made of G-10 fiberglass. These bays, one for the drogue

parachute and the other for the main parachute, will serve as structural elements as well as sealed

compartments for recovery purposes. At the end of each section is a sealing bulkhead with a U-

bolt to which adjacent sections of the launch vehicle are tethered, in addition to recovery devices.

Launch Vehicle Performance Analysis 3.5.

3.5.1. Altitude Predictions and Motor Selection

Mission performance predictions are based on a projected rocket mass of approximately 15.7 kg.

This estimate does not include the mass of the motor case and propellant. However, the mass of

the rocket motor case and propellant is factored into the flight performance simulations that have

been conducted. The motor selected for flight, which will result in the vehicle attaining the




target altitude, is an Aerotech L1390. The Aerotech L1390 has a total impulse of 3949.0 Ns and

will keep the launch vehicle sub-sonic throughout flight.

Altitude simulations were performed in Open Rocket on a model representing the launch

vehicle's dimensions, mass, and Center of Gravity (CG) location. Table 17 below lists the

altitude output from various simulations. Additionally, Figure 11 is a plot of altitude and thrust

versus time for the selected flight motor.

Table 17: Altitude as a Function of Motor Selection (Constant Dry Mass)

Rocket Mass (without motor) Motor Altitude (ft)

15.7 kg L1390 (Aerotech) 5,272

15.7 kg L1720 (Cesaroni) 5,010

15.7 kg L1482 (Loki) 5,210

15.7 kg L1520 (Aerotech) 5,023

15.7 kg L1355 (Cesaroni) 5,190




Figure 11: L1390 Altitude and Thrust vs. Time




3.5.2. Stability

At liftoff, the rocket has a stability margin of 2.3 calibers. The motor selected for flight provides

a launch rail exit velocity of 65 ft/s, which is sufficient for stability. The CG versus Center of

Pressure (CP) (or stability margin) for the flight is plotted below in Figure 12. The stability

margin increases to approximately three calibers during the coast phase of flight. Regarding

sensitivities, simulations were produced with wind speeds up to 20 MPH. At the maximum

tested wind speed, the lowest simulated maximum altitude was 5,200 feet, and the lowest launch

rod exit stability margin was 0.75, with a coast phase margin of 2.5.

Figure 12: Launch Vehicle Stability vs. Time




3.5.3. Testing

The sub-scale launch tested the Vespula Mk II modified tube fin design. The successful sub-

scale test occurred on Saturday, October 13 and is illustrated below in Figure 13. The modified

tube fin design was de-scoped when the switch to the Intimidator Kit was made.

Figure 13: 45% Scale Test Rocket and Flight

At this time, the competition launch vehicle has not been flown due to previous in-flight failures

of the Vespula Mk II launch vehicle. The test flight of this vehicle will be completed before the

FRR telecom, and additional vehicle data will be presented during the presentation. The

presentation will include a drag assessment and the comparison and validity between flight

results and predicted results obtained via analysis tools.




Intimidator 5 Kit Mass Breakdown 3.6.

The mass breakdown for the Intimidator 5 Kit is illustrated below in Table 18 and the landing

masses for the Intimidator 5 Kit are illustrated below in Table 19. The graphical breakdown of

the landing masses are illustrated below in Figure 14. The values obtained for the booster,

payload, and nosecone sections were obtained through weighing each section individually while

everything was assembled with the exception of the motor. In addition, the values for the drogue

chute, main chute, shock cords, and the motor case are also actual weights obtained from a scale.

Table 18: Overall Weight Breakdown

Component Weight (lb.) Quantity Total Weight (lb.)

Booster 11.2 1 11.2

Payload(w/ nosecone) 8.071 1 8.071

Main Parachute 1.7 1 1.7

Drogue Parachute 0.109 1 0.109

Shock Cord 0.755 2 1.51

Motor 7.87 1 7.87

Motor w/out propellant 3.41 1 3.41

Science 10 1 10

Avionics 5 1 5

Total Take-Off

Weight

45.46




Table 19: Intimidator 5 Kit Landing Masses

Section Landing Mass (lb.)

Nosecone 1.619

Payload 16.452

Booster 19.61

Total 37.68

Figure 14: Intimidator 5 Kit Landing Mass Breakdown

Interfaces and Integration 3.7.

The interfaces between the launch vehicle and the ground, and ground launch system, shall be

described such that the operation of interfacing the launch vehicle with these systems can be

correctly carried out to ensure optimal launch vehicle performance, with maximum safety to the

USLI team, and so that a sustainable architecture can be developed to show new members the

necessary action items of launch vehicle/ground/ground launch system integration.

1.619 lb.

16.452 lb.

19.61 lb.

Nosecone

Payload

Booster




3.7.1. Interface with the Ground

The launch vehicle will have a GPS tracking system that will deliver real-time telemetry, as well

as the launch vehicle’s landing location, to the ground tracking station via an XBEE radio

transmitter. When the power system is locked to the ON position on the launch pad, the XBEE

will begin transmitting telemetry data.

3.7.2. Interface with the Ground Launch System

The launch vehicle will have attached large launch lugs, so that it can fit within a launch rail with

an aluminum 1515 T-slotted extrusion, of a minimum length of 8 feet. The launch vehicle will be

placed on a launch stand designated by the LCO after being inspected and certified flight-worthy

by the RSO. After proper assembly and insertion of the motor, inspection and certification, and

attachment to the launch stand, the electronics necessary for the payload and recovery system,

will be activated and locked into position. The altimeter will announce the readiness of the

electronics and payload system via a series of beeps. The launch vehicle will be launch using

standardized launch equipment including a standard 12 volt direct current firing system.

Launch Vehicle Operations 3.8.

It is the responsibility of Launch Operations to create comprehensive guides and checklists to

ensure proper operation of the launch vehicle and the safety of the USLI team. Proper operation

of the launch vehicle requires that certain protocols and procedures are observed by the Ramblin’

Rocketeers team during assembly and launch.

3.8.1. Launch Checklist

The Launch Checklist ensures that all tasks necessary for a successful launch are completed and

completed in the most efficient order. The Launch Checklist has both a performer and an

inspector to ensure all tasks are completed correctly. In addition, there is a Troubleshooting

Chart to address common problems when preparing and launching rockets. The Launch




Checklist remains largely unchanged from the previous year in which the launch vehicle was

prepared for launch in one hour. Because of this the Ramblin’ Rocketeers are confident that the

time needed to prepare the launch vehicle for launch will remain well below the two hour

requirement, LV-4.5. The Launch Checklist can be found in Appendix II.




4. Flight Experiment

Introduction to the Experiment and Payload Concept Features & Definition 4.1.

With the rise of entrepreneurial space flight, many new exotic spacecraft are being designed for

the purpose of finding a profit in space. Many of these spacecraft will be equipped with liquid

fuel propulsion and attitude control systems, or will seek to store large quantities of liquid

propellant. These liquids present difficulties in the design and operation of a spacecraft because

in low gravity, the fluids will be dominated by a combination of capillary/inertial/gravity

gradient forces and will respond to perturbations. The response of stored liquids to such

perturbations is termed slosh, and slosh is known to 1) alter the inertia matrix of a spacecraft and

2) to hamper the use of vents and propellant feed lines. Some methods of controlling slosh are

listed in Table 20.

Table 20: Methods currently available for damping slosh.

Damping Method Description

Tank geometry The choice of tank geometry (cylindrical, spherical, toroidal, etc) is known to

have an impact on slosh damping through viscous effects.

Ring baffles Annular disks along the circumference of a tank that impede slosh and may be

given various camber geometries.

Lids and mats Lids and mats float on a free surface of the liquid and impede slosh.

Floating cans Cans impede slosh by absorbing and dispersing the kinetic energy of the liquid.

Expulsion bag or

diaphragm

Bags and diaphragms reduce slosh by containing the propellant and forcing it

into propulsion feed lines.

Non-ring baffles Non-ring baffles are baffles that do not necessarily follow a tank circumference,

e.g. cruciform baffles.

Flexible baffles Flexible baffles are baffles made of flexible materials that deform under the

inertia of sloshing liquids.

While present methods of reducing slosh may be very effective in some flight regimes, there are

design issues inherent to some of these systems. For baffles – perhaps the most effective




dampers for the additional inert mass – instabilities can occur during launch if propellant levels

are below the lowest baffle as in the case of the Saturn I. Similarly, such problems could occur

in low gravity situations where the baffles are rendered ineffectual from lack of contact with the

liquid. However, with the expense of mechanical complexity and inert mass, expulsion bags and

diaphragms can be used to avoid such instabilities. The Ramblin’ Rocketeers intend to provide

another alternative solution by demonstrating the use of magnetorheological (MR) fluid as a

moveable, deformable baffle and potentially a diaphragm equivalent.

Accomplishments Since CDR 4.2.

Since CDR, the team has pushed forward with hardware development. Despite delays due to

shipping and receiving of parts, as well as longer-than-predicted manufacturing times, ground

testing is about to begin and should be completed before April. A full science payload for the

rocket is on track to be completed on April 2, just before the RGEFP Technical Experiment Data

Package deadline of April 3.

Additionally, a science package was flown on the first test flight, giving important insight to

improve the science payload data acquisition. A second, upgraded package more similar to the

electronics that will be flown on the flight at Huntsville launched on March 9.

4.2.1. Important Changes

All use of cameras for the science payload has been de-scoped. Bench-testing of the solenoid

control system is prioritized over the two-plate testing, so that important progress is made

towards finalizing both the USLI and RGEFP payloads.

4.2.2. Test Launch Lessons Learned

4.2.2.1. Summary of Science Team Payload

The science team flew an Arduino Mega, OpenLog, and two piezo-electric vibration sensors to

obtain preliminary data on the launch vehicle environment. The power supply was a 6V AA




battery array with DC connector. Payload tubes for the purpose of mass simulation were also

flown but were not within the scope of the science team interests of the full-scale test flight.

4.2.2.2. Report of Failures and Occurrences

4.2.2.3. Integration

The science electronics suffered a lack of integration consideration in design. A vertical

integration scheme was necessary for the Arduino Mega, however the protoboard shield and

power supply had cumbersome dimensions and securing these items increased the risk of failure.

A protoboard pieces was cut unnecessarily, endangering solder joints. Finally the Mega was

mounted vertically with duct tape and redundant zip-ties, securing it to the rocket structure.

4.2.2.4. Sensor detachment

A sensor was observed to have been detached from the terminals at some point during the

tumbling phase of the rocket trajectory. However, this sensor appears to have had its supporting

circuitry shorted to Arduino ground because of the manner of integration and a lack of electrical

tape on the Mega USB B header. The result was saturated readings of the secondary (radial)

piezo-electric sensor. This was a non-critical failure.

4.2.2.5. Openlog File Writing

A secondary file was begun at some point in the writing of data. There is no time-step matchup

between these files. Thus it is difficult to understand where in the timeline of the launch vehicle

trajectory that the open log reset. This is a potential failure risk for the SIDES nodes that could

disrupt the coherence of time data. The data are still available for analysis but are made difficult

to interpret.

4.2.2.6. Results and Future Mitigation

4.2.2.7. Integration Results




One of the two zip-ties where broken during the spin of the launch vehicle. Fortunately a second

zip-tie provided redundancy. Integration systems resistant both to typical axial loads and

potential rotational loading should be considered for future launches.

4.2.2.8. Sensor data

Sensor data was recovered and is being analyzed. The data recovered are presented in Figure 15

and Figure 16.

Figure 15: Correcting for Piezo drift Figure 16: FFT of corrected data showing peak around 26 Hz

Sensor data might be improved by better mating of the sensors to the mounting brackets or

rocket structure. The data show strong peaks and piezo drift and should be useful in comparison

of future datasets.

4.2.2.9. OpenLog Risk Mitigation

To mitigate the risk of openlog errors – i.e. starting to write a new file with dissimilar time stamp

– or other issues that might arise unexpectedly in the SIDES network, as well as to better analyze

the data, an accelerometer will be flown on the science board so that the launch vehicle

accelerations can be precisely correlated with vibration data.




Science Background 4.3.

A complete science background is included in Appendix III.

4.3.1. Important Highlights

In the science background, several important relationships are developed. These relationships

are given in Table 21.

Table 21: Elements of the theoretical modeling for the LSIM payload

LSIM element Relationship

Response of MR fluid

to a magnetic field

(motion as it

rigidifies)

Longitudinal Slosh

Model ∆�̈� = 𝑔 −

𝑏2𝑚�Δ�̇�� −

𝑘𝑚∆𝐿

Lateral Slosh Model �̈� = �−

𝑔𝑘(𝑘𝐿 + 𝑚𝑔)� 𝜃 − �

𝑏1𝑚� �̇�

Damping of MR

Fluid if it is a rigid

baffle

By experiment and in correspondence with reference material tabulated data and

plots

Damping of

Container

From reference material,

𝛿 = 4.98𝜈1/2𝑅−3/4𝑔−1/4

Experiment Requirements and Objectives 4.4.

4.4.1. Success Criteria

Minimum and maximum success criteria have been defined for the LSIM payload. Table 22 lists

these criteria.




Table 22: LSIM success criteria from the Requirements Verification Matrix

LSIM Success Criteria

Minimum Successfully record video of flight experiment during microgravity and start/stop the

experiment without mechanical and electrical failures.

Maximum Successful matching of the damping ratio for ringed baffles in the wave amplitudes

experienced during flight to within ±30%.

4.4.2. Requirements

The requirements set for the LSIM experiment to satisfy both the goals of Ramblin’ Rocketeers

and the USLI requirements are listed in Table 23. Flight systems (experiment and avionics) are

now budgeted to be 15 lbf.

Table 23: LSIM Requirements

Requirement

Number

Requirement

Definition

Source Verification

Method

Design

Feature

Status Verification

Source

Document

FS-1 The flight systems

team shall design and

build the LSIM

Payload

MO-3 Inspection LSIM payload

In Progress MO-3

FS-2 The LSIM payload

shall be designed to

fly on a SLP rocket

USLI Handbook

3.1.1 Inspection LSIM

payload In

Progress

USLI Handbook

3.1.1


Team shall produce a

working system for

manipulating MR

fluid in LSIM.

MSC-3 Testing Solenoids

and Control Algorithms

In Progress MSC-3




Requirement

Number

Requirement

Definition

Source Verification

Method

Design

Feature

Status Verification

Source

Document


Team shall ensure

that all avionics are

properly shielded

from the LSIM

payload.

MSC-3 Testing

Faraday cages and webbing tied to

ground on the harness

Not Started MSC-3


Team shall design all

LSIM components

and avionics such

that they may be

easily integrated with

the Modular Payload

System of the

payload bay in the

rocket.

MSC-3 Inspection Mounting system Complete MSC-3


Team shall conform

to all weight, power,

and dimensional

requirements as per

the rocket design.

MSC-3 Analysis TBD In Progress MSC-3




Requirement

Number

Requirement

Definition

Source Verification

Method

Design

Feature

Status Verification

Source

Document

FS-7.1 The Experiment and

Avionics, with

mechanical supports,

shall weight no more

than 15 lbf.

LV-1.1 Inspection TBD In Progress LV-1.1

FS-8 The flight computer

shall execute all tasks

necessary to the

operation of the

LSIM payload and

avionics.

MSC-3 Inspection Maple SIDES node

In Progress MSC-3


shall have a

dedicated power

supply.

MSC-3 Inspection SIDES node In Progress MSC-3


Team shall ensure

redundancy and

reliability of all

internal electrical

hardware.

MSC-3 Inspection SIDES network

In Progress MSC-3




Requirement

Number

Requirement

Definition

Source Verification

Method

Design

Feature

Status Verification

Source

Document


Team shall provide

for payload operation

with up to 1 hour of

wait on the launch

pad and 2 hours of

wait during

preparation of the

Rocket.

USLI Handbook

1.6 Inspection TBD In

Progress

USLI Handbook

1.6


Team shall provide

for electrical

operations to begin at

the beginning of the

flight trajectory.



Team shall ensure

that the LSIM

payload is shut down

safely during the

deployment phase of

the flight trajectory.





Requirement

Number

Requirement

Definition

Source Verification

Method

Design

Feature

Status Verification

Source

Document

FS-14 Data from the LSIM

payload shall be

collected, analyzed,

and reported by the

team using the

scientific method.

USLI Handbook

3.2 Inspection

Data logging in

SIDES network

In Progress

USLI Handbook

3.2


will be designed to

be recoverable and be

able to launch again

on the same day

without any repairs

or modifications.

USLI Handbook

3.5 Inspection

Appropriate mounting

to the payload

interface.

In Progress

USLI Handbook

3.5

4.4.3. Hypothesis and Premise

The hypothesis posed in the LSIM experiment is that –

If a baffle can be manipulated during the flight of a spacecraft, then unstable slosh

can be actively damped.

The experiment will apply radial magnetic fields to the propellant tank to manipulate and rigidify

the MR fluid during the microgravity phase of the launch vehicle trajectory to perform Liquid

Stabilization in Microgravity – LSIM. The launch vehicle ascent will provide a high vibrational

intensity environment in which to test the anti-slosh system. Furthermore, the use of diaphragms

and propellant bags are eliminated with the assumption that:

Trading mechanical complexity for electrical complexity is preferable from a reliability

standpoint.




Therefore, the Ramblin’ Rocketeers will implement a design to apply these concepts to both the

launch vehicle and RGEFP.

4.4.4. Experimental Method and Relevance of Data

The experimental method for LSIM requires a multi-step approach for ground testing, flight

testing, and RGEFP. The purpose of ground testing will be to characterize the shear stress

behavior of MR fluid of different composition and magnetic field configuration, the

manipulation of MR fluid, and preliminary data on slosh damping ability. Flight testing will

provide actual data on the capability of the MR fluid system to dampen slosh, especially in the

microgravity environment. RGEFP would seek to explore a “big-picture” system that actively

attempts to remove any stray MR fluid as propellant simulant is pumped out of the tank. In any

of the test cases, an optimal mixture of MR fluid will enable an application of active control to

maneuver MR fluid into position in flight. The testing cases are organized by the team testing

matrix for LSIM, which is designed to enable comparative analysis of the results and to verify

completion of the data set. Following the scientific method, the test matrix outlines control

experiments and baseline comparisons to develop a qualified understanding of MR fluid in the

context applicable to LSIM. A summary of scientific method fulfillment is given in Table 24.




Table 24: Scientific method fulfillment for LSIM

Method step Fulfillment

Question What are options for electrically damping slosh?

Research Study of MR fluid and a review of “The Dynamic Behavior of Liquids

in Moving Containers”

Hypothesis If a baffle can be manipulated during the flight of a spacecraft, then

unstable slosh can be actively damped.

Test Ground testing plan and test matrix, flight test, RGEFP

Analysis Data examination, post-processing, and analysis

Communicate SLP documentation and VTC

Furthermore, in an improvement over previous experimental design, the team intends to fly a

control experiment as part of the flight test, permitting greater validating capability for the

effectiveness of the damping system.

Testing plan 4.5.

4.5.1. Overview

To accomplish the objectives of LSIM, several distinct testing sequences are necessary. Key to

the success of LSIM is ground testing, where MR fluid mixtures will be characterized and

manipulated with solenoids. Following on these tests are the USLI flight test and separately the

RGEFP project. However, at this juncture of the project some testing has been de-scoped,

namely shake-table testing of the “bench test” platform to demonstrate slosh reduction in 1-gee.

This test was de-scoped due to the complexity and time constraints of procuring an appropriate

shake-table.




Table 25: Test sequences and descriptions, included options de-scoped since PDR

Test Sequence Explanation Purpose

MR fluid characterization A “two-plate” test

To determine experimentally the

viscosity performance of MR

fluid mixtures.

Bench testing Solenoid operations

To develop a method of control

for rigidifying and raising MR

fluid within a canister.

Descoped sequences Shake table testing Formerly, to test slosh reduction

in 1-gee.

Launch Vehicle test USLI flight test

Control and experiment test

inside the launch vehicle to

determine comparative reduction

in slosh.

RGEFP Microgravity University

cooperative project

Up-scaled testing and feasibility

study of MR fluid slosh

reduction.

A brief description of each testing sequence is given in Table 25 above and the relationship

between these testing sequences is illustrated in Figure 17 below.




Figure 17: LSIM testing logic, illustrating a simple relationship of information between the test sequences and emphasizing that they flow down from the pursuit of the LSIM hypothesis.

Ground testing will serve four general purposes: (1) the creation of MR fluid, (2) the verification

and validation of theory and control systems, (3) the characterization of MR fluid, and (4) the

development of a working model for flight testing. For the successful completion of ground

testing, the team will create an optimal mix of MR fluid. An optimal mix will depend on the

fluid's balance between rigidity and fluidity for manipulation under a magnetic field, such that

the MR fluid is easily moved to an appropriate location in the tank. Verifying the Ramblin'

Rocketeers' solution and theory of using MR fluid as a baffle to dampen unstable slosh will go

through two phases. During phase one, only MR fluid will be subjected to a magnetic field.

Phase two will include water along with MR fluid being subjected to a magnetic field. The

results from these phases will indicate whether the solution is feasible by observing the

controllability of MR fluid by a magnetic field as well as observing differences between MR

fluid and the propellant simulant. By characterizing the MR fluid, the team will understand the




various properties of the MR fluid such as its exerted shear force and how it changes under a

magnetic field. The characterization process will include testing the force and viscosity of the

MR fluid and observing preliminary slosh damping. Finally, a working ground model will be

developed using the results from (1), (2), and (3) with constraints for flight experimentation.

4.5.2. MR Fluid Creation and Validation of Theory

MR fluid can be created from three ingredients: carrier oil, magnetic particles, and surfactant.

Table 26 provides example MR fluid ingredients in the design space.

Table 26: List of MR fluid ingredients

Carrier Oil Magnetic Particles Surfactant

Mineral Oil IRON100 Powder Citric Acid

Nanometer particulate ferrofluid IRON325 Powder Oleic Acid

FE100.29 Powder Soy Lecithin

Fe304 M1 Powder

For a preliminary ground test in search of better understanding the behavior of MR fluid –

thereby making more informed decisions on the design space – the team opted to use mineral oil,

IRON325 powder, and oleic acid. By trial and error testing, the team created a stable MR fluid

mixture using the aforementioned ingredients. The team created two mixtures of differing

viscosities. While some sources had presented the iron concentration as 60% by mass, the

preliminary tests found it necessary to increase this percentage. The first mixture resulted to be

too fluid with 17 grams of mineral oil, 1 gram of oleic acid, and 56 grams of IRON325 powder

(76% by mass). The second mixture resulted to be too viscous with 16 grams of mineral oil, 1

gram of oleic acid, and 56 grams of IRON325 powder (77% by mass). The ingredients were

measured using a scale accurate to a gram. Future measurements will use a more accurate scale.

From trial and error testing, the team created an MR fluid testing matrix that will test every




possible combination between ingredients as well as small deviations from the trial and error

test. For example, the team will gradually decrease the iron percentage by mass while gradually

increasing the mineral oil percentage until the optimal mixture – a mixture that appears to be

rigid enough to act as a baffle and manipulative enough to move readily – has been attained.

Each mixture will be static tested by neodymium magnets and good mixtures may be tested with

solenoids as ground testing improves. Validation of theory and control of MR fluid will occur if

there is a change in the MR fluid's viscosity under a magnetic field.

Figure 18: Preliminary static testing of MR fluid mixtures in magnetic fields

From the results of preliminary testing, the composition of MR fluid is likely to be changed to

using carrier oil made of ferrofluid. Ferrofluid is a mixture nanometer-scale ferromagnetic

particles in oil with a surfactant. However, unlike MR fluid, ferrofluid does not have as high a

percentage of pure iron and does not rigidify in the same manner as MR fluid. As carrier oil, the

team hypothesizes that ferrofluid will increase the mobility and useability of the MR fluid

mixture; even with 60% and greater mass ratios of iron powder. Furthermore, smaller iron

particulates may also increase the mobility of the MR fluid. Greater mobility than the initial

mixtures is preferred such that the MR fluid may be moved to the final baffle location using

solenoids, and eventually for the mobility desired for RGEFP.

4.5.3. MR Fluid Shear Stress Characterization: Two Plate Test




The team will determine the shear stress MR fluid exerts inside and outside magnetic fields to

better understand how to manipulate the MR fluid as desired. To determine the shear stress, the

team will perform a two-plate test with and without magnetic field acting upon the MR fluid.

This test was chosen because of its simplicity; other tests such as a barometer test were

considered for measuring the MR fluid's viscosity and force – they turned out too complicated to

realize.

The two-plate test consists of two plates: a bottom plate, which is fixed to the ground and a top

plate, which is free to move. A load sensor will be placed on the top plate to measure the reaction

force that is generated. The current plate choice is acrylic.

Figure 19: Shear stress of a fluid using the two-plate test (Source: Wikipedia)

A control test will be performed by only having two plates together with a load sensor on the top,

moving plate to calculate the frictional force by the plates themselves. For accurate and

consistent results, a mechanical pulling device will be used to pull the top plate. Once a control

has been measured, a quantity of MR fluid will be placed between the two plates and the same

procedure will repeat with and without the MR fluid under a magnetic field. These tests will

characterize the force that MR fluid will generate when it is under a magnetic field and when it is

free of a magnetic field. A complete ground testing plan description is included in the Ground

Test Plan appendix.




4.5.4. Working Ground Model

The team will develop three methods of MR fluid control: an array of solenoids, a movable

solenoid, and a fixed solenoid. For the launch vehicle and RGEFP, solenoid arrays appear to be

the best current option.

4.5.5. Sensors

Piezo-electric vibration sensors will meaure the variation in the force applied to the mounting

brackets.

Design review 4.6.

4.6.1. Viscosity Test Rig

A two-plate test rig was designed to characterize the viscosity of MR fluid. The design of this

test rig underwent many revisions in order to meet measurement and budget requirements. The

objective of this rig is to measure the force of fluid acting on the plate and ultimately, the

viscosity value of the MR fluid. The reaction force and the viscosity constant are related by the

following equation:

𝐹𝑓𝑙𝑢𝚤𝑑��⃗𝐴

= −𝜂𝑉0𝐷

(1)

In Eq.(1), A is the surface area of the plates, V0 is the pulling velocity of the top plate, D is the

distance between the plates, and η is the viscosity coefficient. The surface area of the plate is

31.08 in2, and the distance between the plates is estimated to be about .25 in. A motor is used to

control the velocity, so V0 is also known. Ffluid is measured with a nichrome wire device. This

device measures force through changes in resistivity induced by small changes in cross-sectional

area resulting from tension in the wire. The measurement limits for this device are set by the

accuracy of electrical interface hardware and calibration testing.




One of the design requirements is to simulate a non-frictional surface. The aim is to ensure that

the only force acting on the plate is the fluid’s shear force. This is accomplished by adding

railings between the plates and the base of the structure. Although the railings do not completely

remove friction, they minimize it enough so that friction is negligible. Part drawings for the test

rig are available in the Appendix. The drawings shows locations of the railings’ attachment

points and the tap needed as well. The railings require a #8 type screw, and a ¼ in tap is needed

for hard woods and acrylic sheets with this type of screws. The dimensions and locations of the

features on these parts are mainly driven by two parameters: solenoid strength and test time.

Because the magnetic strength of the solenoid decreases drastically as the distance increases, the

plates need to be close to the solenoid base. This requires short railings. The group also wants to

maximize the contact time between the plates and the MR fluid, so the plates are designed to be

long and skinny. Solenoids are aligned along the length of the plate to produce a uniform

magnetic field during testing.

Construction of the test rig depends on a number of assumptions and is subjected to revision for

alternative methods if necessary. First, super wood glue will be used to connect wood pieces. If

this is not sturdy enough, elbow brackets will be used to connect the corners of the wood pieces.

The motivation for using glue instead of brackets is saving money. Second, wood pieces and

acrylic sheets will also be glued together with epoxy. This method is the norm for connecting

wood to acrylic sheet, and it saves space and money. A #8 type screw will be used to tighten the

railings to the woods and acrylic sheets. The railings attachment holes are designed to provide a

bit of a leeway. Attachment holes on the structure pieces and test plates can be off by about .05

in. A motor will be used to pull the top plate. The motor will stand on a piece of wood that has

been designed with a height that will perfectly align the motor with the top plate. The tolerance

for the height offset is ±.04 in. String and hook will be used to round out the pulling mechanism.

This testing rig was delayed due to the design process and the winter break. Ordering and

construction are planned to begin coincident with CDR and testing could begin around the time

of CDR VTC in late January.




4.6.2. Ground Test – MR fluid production and manipulation

In order to meet the system functional requirements, the MR fluid must have certain properties

and also adhere to certain standards. The fluid must be sufficiently rigid when magnetized to not

shear significantly or break due to fluid slosh; however, the fluid must also not be excessively

resistant to motion when moving and shearing against a wall, so that it may be moved into

position by the magnets. It is known that the size of iron particles makes the largest difference in

the rheometry of the fluid. For instance, one batch of low quality fluid that was created earlier

during preliminary testing, with larger iron particles than is typical of MR fluids, was found to be

extremely resistant to motion.

Therefore, to find a high-quality fluid with intermediate properties, it is wished to test iron

powders with particles of mean diameter between 0.1 µm and 10 µm. In addition, to ensure

purity, we shall attempt to purchase all powders from well-known sources. For example, one

option being explored is a purchase of carbonyl iron powder from BASF. For the carrier fluid,

mineral oil or hydraulic oil are both known to be fairly typical choices; the properties of the fluid

should not be significantly affected by which is chosen. The final choice of components and their

proportions will be made based on the results of the two-plate testing, as well as qualitative

experience from attempts to move the fluid by manually moving magnets.

The MR fluid will then be manipulated by solenoids in a ground testing platform that permits the

placement and use of solenoids to control the MR fluid.

The foundation of the bench test is built from Maker Beam parts. This allows for great

configurability and flexibility. Four 30 cm Maker Bars are the corner stands of the test rig; they

are placed vertically about 20 cm apart from each other. Four 20 cm Maker Bars are placed

horizontally in between the vertical 30 cm Maker Bars and are attached to give support. There

are three thin acrylic plates: one acts as the main base and the other two have holes cut for the

beaker to fit inside the plates. The base plate sits on 4 flat “L” Maker Beam brackets about 2 cm




above from the base. The 4 “L” Maker Beam brackets are attached to the four 20 cm horizontal

Maker Bars. The beaker is placed in the middle of the base acrylic plate and the solenoids are

placed radially around the beaker on the base plate. Four 90 Degree brackets are used to hold the

second plate 2 cm above the first plate. The second plate has a hole cut in the middle to allow for

the beaker to pass through the middle. More solenoids can be placed around the second acrylic

plate. The same is done with the third acrylic plate as was done with the second acrylic plate, but

4 cm above the first, base acrylic plate.

This design using Maker Beam parts allows for future design modification and addition of parts.

Once more data has been collected, the team can attach a vibration motor underneath the first

base plate to simulate slosh. Solenoids can be added, moved, or removed from each of the three

acrylic plates during testing as test results shed more light on what is needed for more accurate

testing. The top two acrylic plates with holes in them can be moved up or down the 30 cm maker

beam bars to adjust the height at which the solenoids interact with the beaker.

4.6.3. Hardware and build progress

4.6.3.1. Sensing

The damping coefficient of the slosh-reduction system will be estimated by measuring the

anchor force decay1F

2 of the experiment. At this moment, the sensors will be measuring the

frequency and magnitude of these forces (i.e. vibrations). These sensors are considered as

Experimental hardware.

2 On the recommendation of p. 108, The Dynamic Behavior of Liquids in Moving Containers: with Applications to Space Vehicle Technology. Ed. H. Norman Abramson. NASA, 1966. (NASA SP-106)




The sensors are piezo-electric elements which

produce voltage as a function of deformation,

especially deflections of the sensor tip. The

base of the sensor is affixed to both the

experiment canister and lower mounting

hardware. The mounting hardware is designed

to permit tip deflection in at least one direction.

This configuration was tested on the USLI

sounding rocket without any liquid within the

rocket experiment canister and the data is still

under review. At the time of writing, the

Figure 20: Piezo-electric sensor used for detecting anchor force oscillations 2F

3

method of sensing the reduction in slosh is open for revision pending further information. Figure

20 gives the sensor and the major dimensions – the thickness is on the order of 0.125 mm. 3F

4 To

operate the sensors, a simple op-amp circuit has been selected, modeled, and built to limit the

voltage and boost the current of the sensor so that a microcontroller, such as an arduino, may

read the sensor data on analog input pins. The circuit is given in Figure 21. A prototype built for

the USLI sounding rocket is given in Figure 22.

3 Image from Sparkfun, https://www.sparkfun.com/products/9196 4 Sensor datasheet: http://dlnmh9ip6v2uc.cloudfront.net/datasheets/Sensors/ForceFlex/LDT_Series.pdf

https://www.sparkfun.com/products/9196

http://dlnmh9ip6v2uc.cloudfront.net/datasheets/Sensors/ForceFlex/LDT_Series.pdf




Figure 21: Piezo-electric sensor circuit. The sensor is modeled as a variable-voltage source at 300 Hz. While 300 Hz is a theoretical maximum for the reading speed of the microcontroller, data was logged at a rate between

88-96 Hz. With six (6) canisters, six (6) of these circuits would be needed to measure the anchor force

oscillations for each canister. The power supply for the op-amp is sourced from the

microcontroller and does not need to interface directly from the aircraft.

Preliminary data illustrates voltage drift and

may help with analysis of tip deflection should

force magnitude be deemed important;

however, variation of the raw data is currently

under analysis for frequency spectra and to

understand the change in spectra over time. It

is current thought by the team that the change

in spectra over the course of the experiment

may illustrate the decay of the anchor force, as

the momentum of the water is changing as a

damped oscillation. The relative frequency of

the resting structure surrounding the

Figure 22: Top and bottom view of sensor prototype circuit. Leads soldered to the piezo-electric sensors are

attached to the blue terminals, while pins go to the microcontroller for data logging and analog reading.

This prototype supports two sensors and is approximately 3 inches by 5 inches. Final boards may

be much smaller.




experiment in the sounding rocket is used as a baseline for that portion of the experiment; no

such baseline of data currently exists for the aircraft, but may be available as a test case. The

data collected from the test launch of the USLI sounding rocket on the prototype piezo-electric

sensor circuit is given in Figure 23 and Figure 24. The analysis of this data is not finalized and

the noise and peaks are not yet understood. An accelerometer will be added to the rocket

payload for increased clarity – it is not yet certain whether an accelerometer will be flown with

the aircraft payload or whether the team will wish to log the provided aircraft acceleration data.




Figure 23: data showing sensor drift and a method of correction by distributing the data around the overall

mean.

Figure 24: frequency spectrum for the entire dataset.

4.6.3.2. Solenoids




The solenoids for the experiment currently consist of 3 50-turn layers of 32 AWG copper magnet

wire for a total of 150 turns per solenoid. This is approximately 3 times more turns than

necessary to produce the impedance needed to pull 10W at 5VDC, implying that the power draw

will be lower due to greater impedance. This decision was made to increase the field strength

generated by the solenoid. Solenoids may need to be resized if not enough power is available at

the containment box panel – also, not every canister within the containment box will have

solenoids, especially if only six (6) canisters can be flown totally as the experimental controls

must be included. An iron core is used inside of the solenoid diameter. The solenoids current

constructed are manufactured by hand, and are classified as Experimental hardware for the

aircraft.

The solenoids would be controlled by solenoid drivers mounted to one or several

microcontrollers. The currently selected driver is given in Figure 25.

Figure 25: 4x4 solenoid driver. Two drivers can be linked together per microcontroller to control 32 solenoids directly. Approximately 3 inches by 2 inches. 4F

5

5 Image from Sparkfun, https://www.sparkfun.com/products/11352

https://www.sparkfun.com/products/11352




4.6.3.3. Microcontroller

Currently proposed as the microcontroller for the experiment is the Arduino Mega. It is not

certain whether multiple microcontrollers will be necessary. If so a USB multiplexer may be

required but is not yet chosen. The microcontroller is illustrated in Figure 26.

Figure 26: Arduino Mega microcontroller with major dimensions.

Payload Relevance and Science Merit 4.7.

The top priority of the Flight Systems team during project development was to create a payload

concept leveraging team expertise while pursuing achievable and NASA-relevant experiments.

Previously, the 2009-2010 project investigated moving oxygen gas with an electromagnet –




essentially a steady-state siphon for paramagnetic materials. The 2011-2012 USLI team

investigated active platform electromagnetic stabilization, developing control algorithms for

magnetic levitation during flight. After review by Flight Systems and the Georgia Tech

Ramblin’ Rocketeers, the team decided that the most relevant primary payload would be to

demonstrate the use of MR fluids in anti-slosh applications using technology development from

the 2009-2010 and 2011-2012 Georgia Tech USLI experiments. Combining technologies from

the previous projects, the new LSIM payload will demonstrate a possible method to combat

propellant sloshing. The benefits of such an anti-slosh system would be most applicable in deep-

space long-duration missions. In such missions, large quantities of fuel must be stored and/or

transported with cargo/personnel. A major issue in low-gravity environments for propellants is

sloshing, where fluid begins to float freely in space relative to the propellant tanks. Sloshing

may cause loss of pressurization in propellant feed systems, potentially creating dangerous

propulsion failures. The current solution is to create a moveable and deformable baffle from MR

fluid. Using electromagnets, the controlled fluid may then be used to dampen the propellant

oscillations. Systems might be needed to insure that the fluid is removed from the propellant,

and a magnetic siphon could be used if the mixing between fluid and propellant is minimal. This

is the basis for the RGEFP experiment discussed later in this document.

Generally however, the LSIM experiment is a science and engineering payload that involves

phenomena from several fields, primarily magnetism, rheology and viscous flow, as well as near-

inviscid fluid dynamics. Among the goals of LSIM is to develop a scientific model

encompassing all of the above fields in order to understand the interactions between the various

components of the system. This will be achieved by combining theory with experimentation and

testing. Data will be collected for variables such as MR fluid position, MR fluid shear stress, and

simulant position and acceleration as a function of time, rocket acceleration, and electromagnet

currents and positions. Collecting this experimental data will enable changes in the applied

control scheme to be made according to the observed data, as well as allowing for refinement of

the dynamic and scientific model of the MR fluid-propellant simulant system. A full explanation




of the science of slosh and MR fluid relevant to LSIM is included in the Science Background

appendix.

For MR fluids, the primary focus of research in current years has been on the properties of the

MR fluids themselves, and on their interactions with solid objects or containers, rather than on

their interactions with other fluids. Therefore, the LSIM experiment should give insight into this

less-studied subject. In addition to the above modeling, there are other scientific benefits of this

experiment. The behavior of MR fluids in microgravity has been of significant interest, with the

InSPACE experiment on the International Space Station being a large-scale investigation on this

topic. However, engineering applications of the fluid specifically in microgravity do not seem to

have been investigated to the same extent. Microgravity is one of the places where MR fluid is

likely to be most effective, as settling of iron particles and thus degradation of integrity does not

occur in the near-absence of gravity. Therefore, the LSIM experiment allows for investigation of

actual applications of MR fluids in microgravity, as well as scientific modeling of the MR fluid-

simulant dynamics.

RGEFP 4.8.

The Ramblin’ Rocketeers were selected for RGEFP. Reporting on the progress of this portion of

LSIM will be to the JSC Microgravity University and TEDP documentation.

4.8.1. RGEFP-Specific Design Work

4.8.1.1. Containment Box

The containment box serves the role of supporting loads from aircraft accelerations, supporting

the experiment canisters, providing interfaces for electrical hardware, containing any leaks from

the canisters, and providing access to the canisters in the case that the initial canisters may be

exchanged for differently configured canisters during the flight.




The redesigned containment box is 4’ long in the forward direction of the aircraft, 3’ along the

span, and 2’ tall. An illustration of the current containment box idea is given in Figure 27.

Figure 27: Ideas for the containment box, illustrating some support elements and a possible electrical conduit.

According to the aircraft ICD, the Georgia Tech containment box could be placed near any of the

panels, as all panels provide 115V AC and 28VDC.

To simplify the design of the box, the lower components to mate the canisters to the box will be

rescaled versions of hardware already designed for the USLI sounding rocket. These mounting

brackets will be resized for a 4”-maximum diameter tube and will provide a lower support for the

canisters. So that the canisters are not torqued by their own inertia, an upper support will be




attached to the containment box lid that will drop over the top of each canister. Further supports

and trusses will be added as needed to support the manage inertia inside the box. The box,

internal supports, and brackets are classified as Experimental hardware.

The design of the mounting bracket for the USLI sounding rocket, and therefore a smaller

version of the bracket intended for use in the aircraft, is given in Figure 28.

Figure 28: bottom mounting bracket for the USLI sounding rocket. A larger version is intended to be used in the containment box. This piece attaches to the box - a second part attaches to the canister and snaps into the

bracket. 4.8.1.2. Computer

A laptop will be used for logging data and sending commands to the microcontroller(s) within

the containment box.




4.8.1.3. Weights

Figure 29 and Figure 30 give a preliminary weight budget for known assemblies and

subassemblies. The average density of polycarbonate used in the weight computation is 1200

kg/m3. Currently, CBE for weight is 205.30 lb; this is felt to be conservative since the CBE for

weight without contingencies is 164.85 lb – a margin of 17.6%. Lighter weight solutions for the

containment box and mounting are being developed.




Figure 29: Current weight budget with totals, and broken out by known subassemblies.




Figure 30: Summary of the weight budget to report subassembly totals.

4.8.1.4. Equipment Layout for Take-off, in Flight, and Landing

Currently there is no apparatus designed to hold extra canisters to be swapped with the six (6)

canisters that begin the flight during take-off in the containment box. However, the general

configuration during take-off, flight, and landing for the containment box and internal canisters

is the same. This configuration is given in Figure 31.




Figure 31: Equipment layout for containment box, 6 canisters, laptop and crew for all stages of flight. The precise placement of operators during parabolas is viewed as non-essential so long as the

cabling from containment box to laptop remains intact. Crew may need to change canisters

during hypergravity – precise configuration for this activity is not currently known.




Flight Experiment Integration 4.9.

The payload includes all experimental components. A possible configuration for the payload is

shown in Figure 32. The assembly is made of four parts: the base bolt, the base, the payload

plug, and the payload. General dimensions for the payload are listed in Table 27. .

Figure 32: Payload Assembly

Table 27: Payload Assembly Dimensions Parameter Value

Base Diameter 4.97”

Total Height 10.58”

Payload Height 8.95”

Base Thickness 0.1”

The experiment is housed in a PVC plastic pipe that is connected to a base. The payload base is

designed to be the only load bearing component of the payload assembly.

Base Bolt

Payload Plug

Payload

Base




Figure 33: Payload Base with 150N of loading

The base rests in the rib of the structure and holds all of the weight of the payload and any

sensors used. It is made of Delrin plastic and manufactured using injection molding. The payload

base can support roughly 60.85lbs of load before failure. It is designed to support an assumed

maximum load of 30.425lbs with a factor of safety of 2. This load comes from the assumption

that the payload weighs no more than 3lbs accelerated at 10 times the acceleration due to gravity.

Figure 33 shows the stress distribution through the base, using SolidWorks SimulationXpress

Wizard.




Table 28: Data from SolidWorks SimulationXpress, highlighting the data from assumptions

Trial Total Load (lbf) Max Stress (psi) Factor of Safety

1 2.248 337.503 27.07

2 4.496 675.131 13.53

3 6.744 1012.653 9.02

4 8.992 1350.257 6.77

5 11.24 1687.804 5.41

6 13.488 2025.392 4.51

7 15.736 2362.954 3.87

8 17.984 2700.457 3.38

9 20.232 3038.089 3.01

10 22.48 3375.607 2.71

11 24.728 3713.220 2.46

12 26.976 4050.785 2.26

13 29.224 4388.348 2.08

14 31.472 4725.908 1.93

15 33.72 5063.411 1.80




Table 28 shows data taken from SolidWorks SimulationXpress for the payload base. This data

was interpolated to find the maximum load of the payload base. Figure 34 shows the factor of

safety plotted versus the total load on the payload base. The graph and equation allow the

approximate maximum load to be determined mathematically before constructing the first

prototypes.

Figure 34: Factor of Safety vs. Total Load from SolidWorks SimulationXpress and generated trend line equation

y = 60.849x-1

0.00

5.00

10.00

15.00

20.00

25.00

30.00

0 5 10 15 20 25 30 35 40

Fact

or o

f Saf

ety

Total Load (lbs)

Factor of Safety vs. Total Load Payload Base (Delrin 2700)

Factor of Safety




5. Flight Avionics

Feedback is essential to any meaningful design work. In recent years, the Ramblin’ Rocketeers

have implemented a number of unique launch vehicle designs, each with the intention of finding

solutions to particular problems. However, with only limited visual feedback available, it is

difficult, if not impossible to gauge the success of a design or to detect any unanticipated failure

modes. A system that could accurately describe the state of the rocket throughout its flight

would then be enormously valuable. To be effective, such a system would have to be capable of

not only recording data from multiple sources but also able to temporally connect the data. This

would provide the user insight into the interactions between different factors in addition to the

individual measurements. Due to the potential complexity of such a design, the system also

needs to be tolerant to the potential failure of any singular functional unit. This would ensure that

even if some information is lost, the system will still yield meaningful feedback from tests.

Finally, it would be helpful for such a system to be extensible. It is impossible now to envision

all of the potential use cases for such a system. Designing it to be easily adapted to the needs of

future projects would help ensure its success and longevity.

Avionics Overview 5.1.

The avionics are designed to accommodate the primary science payload LSIM, in addition to

supporting structural and aerodynamic analysis of both the advanced fin design and the ‘rib and

stringer’ fuselage design of the Vespulla MkII. To accomplish this goal, SIDES (Simultaneous

Independent Data Logging & Experiment System) was developed to maximize the data extracted

from each flight while reducing the risk of failure of a larger avionics system. SIDES

architecture allows for a flexible, complex, and fault tolerant distributed data collection system

for the Ramblin’ Rocketeers launch vehicle.




Table 29: Avionics requirements

Requirement

Number

Requirement

Definition

Source Verification

Method

Design

Feature

Status Verification

Source

Document

1. The flight avionics

shall collect data

required for a

successful payload

experiment.

USLI Handbook

1.7 Testing Data logger Complete

2. Key elements of the

flight systems shall

operate on independent

power supplies.

MSC-3 Testing SIDES nodes Complete

3. Power supplies should

allow for successful

payload operation

during launch vehicle

flight with up to 1 hour

of pad stay and 2 hours

of standby time during

launch vehicle

preparation.

MO-4 Testing

Battery systems and

power management

Complete

4. The flight avionics

shall be capable of

being attached to the

launch vehicle

structure.

USLI Handbook

2.11 Testing Mechanical

Interfaces Complete




Requirement

Number

Requirement

Definition

Source Verification

Method

Design

Feature

Status Verification

Source

Document

5. GPS coordinates of the

launch vehicle shall be

transmitted to a ground

station.

USLI Handbook

2.11.1 Testing GPS, Ground

Station, Xbee Complete

6. Each avionics node

shall be capable of data

logging with or

without a clock pulse.

USLI Handbook

2.12 Inspection

Flight Software and Data

Logger

Complete

7. Each avionics node

shall operate at some

equal or reduced

functionality during

RS485 communication

failure

USLI Handbook

2.12.1 Inspection

Flight Software

and Redundant

Node Hardware

Complete

Avionics Success Criteria 5.2.

The success of the Ramblin’ Rocketeers avionics team will be defined in two ways: minimum

success criteria that will be accomplished if the requirements are accomplished, and maximum

success criteria that will be met if everything goes according to plan. Maximum success will

include collecting diagnostic data for the launch vehicle, such that design feedback is available

for iterating the most effective launch vehicle design, while minimum success is limited to

successfully collecting and storing the LSIM payload data for recovery and analysis of the data.




Table 30: Avionics Success Criteria

Requirement

Numbers


Method

Design

Feature

Status Verification

Source

Document

1. The avionics system is

functional throughout

the flight and if failures

do occur the entire

system does not go

down.

Analysis,

Testing

Complete

2. The ground station

should be capable of

receiving

supplementary data

transmitted from the

launch vehicle.

Analysis,

Testing

Complete

3. The ground station

should detect the

location of the launch

vehicle throughout the

flight, and track the

location of the landing

for recovery purposes.

Analysis,

Testing

Complete

SIDES Design Approach 5.3.

SIDES utilizes a distributed network of microcontrollers to accomplish diverse tasks. Each node

in the distributed network is capable of operating independently of other nodes. To support this,




each node has a self-contained power supply and data logging capability. This approach reduces

risk by preventing the failure of any node from propagating through the SIDES network.

Distributed data logging presents a synchronization challenge when compiling distributed data.

The integration of the data when clock skew is present becomes much more difficult and often

involves resampling and interpolating the data to obtain useful results. By providing a

synchronization clock signal, the local data logging rates can be easily adjusted to prevent clock

skew.

In ideal operating conditions, the individual nodes of the SIDES network will be able to

communicate over a bus. For noise immunity, the bus will be a differential pair. To optimize

the trade between failure tolerance and weight, electrical harness weight will be reduced by using

a one-to-many, multi-drop bus rather than a point-to-point solution. Software control of the

multi-drop bus nodes will reduce the risk associated with centralized communication while

maintaining the weight advantages of a multi-drop bus.

5.3.1. SIDESboard

The SIDESboard standardizes the nodes, and

helps ease implementation of the electronics.

The SIDESboard contains all the features

necessary at each avionics node to successfully

complete the mission. The SIDESboard has a

standard harness connector, data logging SD

(secure digital) card, battery monitoring circuit,

isolated clock input and a standard mechanical

footprint. The SIDESboard firmware

incorporates a standard set of libraries. These

libraries allow programmers to focus on the

function of the specific node rather than having

Figure 35: SIDES system layout




to code the same functionality each time. The communication bus for the SIDESboard is

handled by an RS485 transceiver. The RS485 format is differential for noise rejection, bi-

directional to save weight in harness wiring, and multi-drop to reduce wiring complexity while

also saving weight. Risk of communication failure is considered to be acceptable for the

purposes of saving weight, because the consequences are low-impact by design. Figure 36 and

Figure 37 depict the SIDESboard PCB (Printed Circuit Board) design, supporting the features

listed above.

Figure 36: SIDESboard bottom side view

Figure 37: SIDESboard top side view

5.3.2. SIDES Electrical Harness




The SIDES electrical harness within the Intimidator kit is extremely simple. Only the Telemetry

and MasterIMU can be wired together, and as such only require a handful of connections routed

between the two standard harness connectors.

5.3.3. Master IMU

The master IMU will house a triple axis accelerometer, gyroscope and magnetometer IMU, and

RS485 hardware. In particular, the Master IMU will facilitate sending data of interest to the

Telemetry node to be forwarded to the ground station.

5.3.4. Science Experiment Computer

The LSIM payload requires multiple vibration sensors, and will also actuate solenoids used to

control the MR fluid during the flight. Due to the switch from Vespulla MkII to the Intimidator

kit, integration changes have made connecting the Science Experiment Computer to the SIDES

network unreasonably difficult. Therefore, this unit will be equipped with an accelerometer to

ensure it can independently determine microgravity timing.

5.3.5. Telemetry

The Telemetry node fulfills the requirement 5 of transmitting the GPS data from the launch

vehicle to the ground station. The Telemetry node will make use of an Xbee GPS transceiver

and a SIDEDboard to log the GPS data while the Xbee is transmitting the data. An example of

the Xbee is depicted in Figure 38.

Figure 38: Xbee transceiver unit




De-scope Options 5.4.

As part of the Ramblin’ Rocketeers’ Flight Systems package for the previous competition cycle,

a computer handling telemetry and GPS was built and flown. This computer has the capability

to run a solenoid driver and read the vibration sensors. Should the SIDES network need to be

de-scoped, this substitute hardware already exists and can be inserted into the system design with

minor modification. More info on this computer may be found in “De-scope Option: Flight

Computer Definition”.

Power Budget 5.5.

Table 31 details the power budget for SIDES.

Table 31. SIDES Power Budget.




EM Interference 5.6.

Faraday cages and shielding webbing may be used to mitigate the risk of EM Interference from

both the telemetry devices and experiment solenoids. More analysis will be needed to determine

the amount of shielding require – for example, placement of electrical harness off-axis of the

experiment solenoids can negate much of the EMI risk inherent in a magnetics experiment. The

precise placement of harness has not been determined as of this point in the launch vehicle

development. Redundancy and robustness is key to the SIDES network and any node failures

should be survivable – further the ground station will provide an added redundancy through more

accurate communications should signal strength from the launch vehicle experience dramatic

fluctuations.

Transmission Frequencies and Protocols 5.7.

The telemetry system is designed to utilize two Xbee PRO 900-XSC modules for one-way

communication from the launch vehicle to the ground station. Using a simple, loss-tolerant

protocol with reliable delivery ensures the data is received if at all possible and that the

information is correct. The SIDES node controlling the Xbee module on-board the launch

vehicle will utilize a 900MHz monopole-monopole vertically polarized rubber duck antenna with

2 dBi gain and 100mW of power. This antenna’s performance is depicted graphically in Figure




39. Receipt of GPS data via radio to the ground station will satisfy the recovery requirement and

bolster kinematics data of the launch vehicle trajectory.

Figure 39: Antenna performance as a function of range

Software Maturity 5.8.

Flight software for the SIDES nodes has been prototyped and robust implementations are still in

development. Ground station software is in development and the progress achieved thus far is

discussed below.

De-scope Option: Flight Computer Definition 5.9.

The following text is pulled from the Ramblin’ Rocketeers 2012 FRR documentation regarding

the flight computer, planned as a de-scope option for the SIDES network.

Flight Computer

The flight computer will be an Arduino Mega which utilizes the ATMEGA 2560AU processor.

The chip has sufficient I2C, serial, and analog inputs to read data from all sensors and log to an




SD card based on Sparkfun’s OpenLog break-out board. Additionally, the chip will run the

Fastrax UP501 GPS module and send the data to an Xbee PRO for transmission to the ground

station. An OpenLog board will provide logging capabilities. The board will be programmed in

the Arduino language, a subset of C++ with some additional libraries. Figure 40 provides a

generalization of proposed flight computer software. Table 32 lists the major components

utilized in this design.

Figure 40: Generalization of flight computer software




Table 32: Major Flight Computer Components

Part Number Component Picture Description

1

The flight computer microprocessor, the

ATmega 2560

2

The GPS receiver, the Fastrax UP501

GPS module

3

The Xbee PRO 900-XSC module for

communication between launch vehicle

and ground station

4

The OpenLog board will provide logging

capability

Avionics Testing and Reliability Assurance 5.10.

Testing was performed on flight systems hardware in order to ensure in-flight success

while recording data and transmitting telemetry information. Test cases were written for each

major sensor to ensure proper hardware functionality. These test cases will be of significant




value for debugging purposes. This will in turn lead to gains in the longevity of the system as

well as in efficiency of programming and future design. This process also led to the creation of a

document detailing the current utilization of input and output resources on each of the

SIDESBoards. Having this document will ensure that no pins will have overlapping utilization

and will facilitate future extensibility by making explicit remaining resources.

Ground Station 5.11.

Amateur rocketry is a test bed for novel aerospace designs; however, normal launches provide

little feedback beyond basic feasibility. This open loop makes it difficult to refine ideas and

identify meaningful or effective designs. While in many cases, acquiring such feedback could be

prohibitively expensive, many performance criteria for vehicles can be acquired through

relatively cheap means with some effort. Detailed visual observation of a launch vehicle can

provide meaningful insight into launch vehicle stability and other important design

considerations. Today even cheap digital cameras can provide levels of detail necessary to give

meaningful vehicle feedback.

Past missions flown by the Ramblin’ Rocketeers have encountered interesting performance

anomalies and have fallen victim to speculation due to limited data collection and some

provocative still camera images. By visually tracking the launch vehicle, unusual flight and

structural characteristics can be positively documented and close the design loop by providing

feedback for the next design iteration.

5.11.1. Purpose

The ground station is designed to ensure communication with and visual observation of the

launch vehicle. Communication quality will be ensured through the use of a high-gain directional

antenna. A digital video camera will be used to observe the launch vehicle throughout its flight.

The ground station will also feature a detachable GPS unit used to make recovery of the launch

vehicle easier.




5.11.2. Function

Both the antenna and camera will be mounted on an alt-azimuthal mount. The mount will have

motors enabling automated rotation of the platform in both of its degrees of freedom. The motion

of the mount will be controlled by a microcontroller that will also be part of the ground station.

In addition to controlling the motors, the controller will also perform the wireless communication

that will receive signals from the launch vehicle via the antenna.

To effectively accomplish its objectives, the ground station must actively track the launch

vehicle throughout its flight. This will be accomplished in one of two ways. The first would use

telemetric data received from the launch vehicle to create a model of the vehicle’s motion. The

second would use a stereo camera system to create disparity maps of the launch vehicle’s motion

and translate these into a series of distance measurements. This could then be used to create a

similar model of motion. The camera zoom will also be adjusted throughout the flight to account

for the changing distance between the base station and the launch vehicle and attempt to

maintain a near constant level of detail.

Table 33: Ground station requirements

Requirement Design Feature

Satisfying

Requirement

Requirement Verification Success Criteria

Accurately receive

telemetric from launch

vehicle

High-gain

direction antenna

Analysis of received signals Sufficient information

for modeling motion

and retrieving launch

vehicle is received




Requirement Design Feature

Satisfying

Requirement

Requirement Verification Success Criteria

Maintain constant visual

tracking of launch

vehicle

High optical

camera,

motorized mount

and control

algorithm

Review of captured video Launch vehicle remains

in FOV through apogee

Provide relative position

information of launch

vehicle for recovery

Detachable GPS

module

Successfully locate launch

vehicle

Successfully locate

launch vehicle




5.11.3. Design Considerations

5.11.3.1. Choice of Antenna

Figure 41: Diagram of a helical antenna

Deciding on the proper type of antenna requires two opposing design characteristics: the

directionality and gain of the antenna. Choosing a higher gain antenna will allow for a greater

range of operation but would give a smaller beam width. This would increase dependence on the

tracking algorithm for ensuring signal quality. A helical antenna offers a good compromise

between these two considerations, with typical examples offering a half power beam width of

20-60° and boresight gains of 8-22 dB. This beam width would give some cushion for latency in

the tracking algorithm. The gain would also be sufficient to ensure good signal quality even

under non-line-of-sight propagation at considerable distance, such as might be the case after

landing.

Figure 42: Typical radiation pattern for a helical antenna




5.11.3.2. Choice of Camera

The choice of video camera posed a similar design decision. Much like an antenna, a camera

provides a certain angular window of coverage. For a fixed number of pixels, increasing this

window will decrease the detail of the captured images. Unlike an antenna, however, these

parameters can be a dynamically changed through the use of zoom. A high optical zoom would

then allow for fairly high detail throughout the flight. Camera choice is further complicated by

the need to algorithmically adjust the zoom of the camera during flight. While this functionality

is built in to most digital cameras, it is seldom available to users programmatically. Models

supporting this functionality often do so at prohibitively high costs.

Figure 43: Canon Powershot SX260

The Canon Powershot SX260 seems to satisfy all of these requirements. The camera is capable

of recording video at 24FPS with an image size of 1920x1080 pixel. The camera also offers and

20x optical zoom. Assuming a 30° vertical field of view or a 60° horizontal field of view, these

parameters mean that at its furthest point, each pixel would correspond to 1.7inches of the launch

vehicle. This camera also offers access to a user-supported firmware known as the Canon Hack

Development Kit which provides direct access to camera operations not offered by factory

firmware. This will considerably simplify gaining direct electronic control of zoom.

5.11.3.3. Motor Sizing




The ability of the platform to track the launch vehicle is inherently limited by the speed and

accuracy at which it can rotate. The rotational speed necessary will be dependent on the angular

velocity of the launch vehicle from the station’s reference frame. Assuming the launch vehicle’s

path is completely vertical from its Launchpad, the angular velocity of the launch vehicle is

given by:

𝑑𝜃𝑑𝑡

=𝑥𝑦′

𝑦2 + 𝑥2 (6)

Where x is the distance from the base station to the launch pad and y is the altitude of the launch

vehicle.

The maximum angular velocity of the launch vehicle will occur during the burn of the motor,

which will occur over the first two seconds of flight. At the end of this acceleration the launch

vehicle will be travelling at 177m/s.

This design will be used at events where participants will likely use at most class M motors. For

this size motor NAR requires a minimum personnel distance of 500 feet 5F

6, or approximately 150

meters. Assuming this distance for x and constant acceleration over the motor burn yields the

following equation:

𝑑𝜃𝑑𝑡

=88.5 ∗ 150𝑡

44.252𝑡4 + 15021𝑠

(7)

This function takes a value of approximately 0.62radians/s at t=1.4seconds. The motor must then

be capable of rotating the mount at a minimum of this speed. Once the moment of inertia for the

mounted camera and antenna has been decided, this value can be used to find the required torque

for the motor.

6 http://www.nar.org/NARhpsc.html

http://www.nar.org/NARhpsc.html




5.11.3.4. Software Maturity

The software operation of the ground station can be broken into a number of logical

components. The process begins by configuring the unit, which consists of initializing contact

with the rocket and initializing the state variables for the rocket and ground station. Once this is

done, the station will then enter normal operation. This state consists of a loop which processes

incoming telemetry information, updating state information for the rocket, deciding whether to

update servo position, and deciding whether to update the zoom of the camera. During this

process, the station will also characterize the state using two Boolean variables, LAUNCHED

and LANDED. Once both of these variables become true, the loop will break, and the station

will transmit the resting coordinates of the rocket to the GPS Pendant. The figures below show

this process and the sub-processes involved in each of these steps.




Figure 44: High-Level Software Process




Figure 45: Updating Rocket State




Figure 46: Updating Servo Position




Figure 47: Updating Camera Zoom




Figure 48:Transmit Rocket Location

5.11.3.5. Effects of Excess RF Radiation on the Recovery Avionics

A simple testing procedure was implemented to ensure the safety of using the e-matches in

proximity to the transmitter. An Xbee transmitter operating at 100mW, with an omnidirectional

antenna was placed next to an e-match at several points of high transmission power along the

antenna and in the near field. The transmitter then sent a variety of packets varying in length

from a single byte to the entire ASCII alphabet. At no point during transmission did the e-match

ignite. This result was expected given the low output power of the transmitter.




Avionics Mechanical Integration 5.12.

The avionics and other electronic systems will be fixed to the rocket through the use of a wooden

sled fixed vertically to wooden ribs on the interior of the booster section of the rocket. The

circuit boards will be bolted to the sled and their batteries will be zip-tied to the sled. The same

method is used to mount the supporting electronics for the scientific payload.




6. General Safety

Vehicle Safety and Environment 6.1.

6.1.1. Overview

Ensuring the safety of our members during building, testing and implementation of the payload

experiment is an ideal condition. Procedures have been created and implemented in all of our

build environments to ensure safety requirements are met and exceeded. A key way the Ramblin'

Rocketeers ensure team safety is to always work in teams of at least two when using equipment

or during construction. This guarantees that should an incident occur with a device the other

member could provide immediate assistance or quickly get addition help if required. The

Invention Studio where the team does a majority of its work is equipped with safety glasses, fire

extinguishers, first aid kits, and expert personnel in the use of each of the machines in the area.

All the members of the payload and flight systems teams have been briefed on the proper

procedures and proper handling of machines in the labs.




Table 34: Risk Identification and Mitigation Steps

Step Name Step Definition

Hazard Identification The first step is to correctly identify potential

hazards that could cause serious injury or death.

Hazard identification will be achieved through

team safety sessions and brainstorming.

Risk and Hazard Assessment Every hazard will undergo extensive analysis to

determine how serious the issue is and the best way

to approach the issue.

Risk Control and Elimination After the hazards are identified and assessed a

method is produced to avoid the issue.

Reviewing Assessments As new information becomes available the

assessments will be reviewed and revised as

necessary.

The steps outlined above in Table 34 are being used to develop a set of standard operating

procedures for launch vehicle construction, payload construction, ground testing, and on all

launch day safety checklists.

6.1.2. Mission Assurance

The top priority of the Ramblin' Rocketeers is the completion of a safe, successful mission with

minimal risk and in-flight anomalies. For this reason, a comprehensive review of all risks

associated with the flight of the launch vehicle must be undertaken to gain a fuller understanding

what can go wrong from the ground preparation stage to vehicle recovery. Risks associated with

the mission may be classified by the probability of occurrence and the severity of a failure.

Table 35 provides a risk assessment matrix with color coding for composite risk severity and risk

class identification for easy reference at a later time.




Table 35: Risk Assessment Matrix with Risk Class

Probability

Seve

rity

Frequent Likely Occasional Seldom Unlikely

Catastrophic I II III IV V

Critical VI VII VIII IX X

Moderate XI XII XIII XIV XV

Negligible XVI XVII XVIII XIX XX

Table 36. Launch vehicle failure modes.

Failure mode Risk class Cause Mitigation

Motor CATO

V

Defective grains

Improper installation

Use proper equipment for motor

assembly

Use instruction manual during

assembly

Only certified Level 2 HPR fliers

should assemble motors

Recovery

separation

failure IV

Insufficient black

powder

Improper venting

Improper wiring

No GPS/data

downlink XIX

Deficient battery

Shorted circuit

Rocket out of range




Payload Safety 6.2.

As already mentioned in General Safety, the same methodology to identify and assess risks for

vehicle safety will be used to identify hazards for the payload. The entire payload and flight

systems teams have been briefed on the possible hazards they may encounter while working with

the payload and how to go about avoiding them. Some of these hazards include inhaling small

iron powder, ingesting inedible substances, and touching harmful materials. Mitigation steps

have been identified for these potential threats. Other hazards that relate specifically to the

payload are listed in Table 37. Payload failure modes are outlined in `Table 38.

Table 37. Payload hazards and mitigation

Hazard Risk Assessment Control & Mitigation

Electrocution Serious Injury/death Do not touch wires that are hot and not

insulated. Wear rubber gloves when

the device is in operation. Handle

leads to the power supply with care.

Use low voltage settings whenever

possible.

Electromagnetic Fields Interfere with electronic

devices inside the body

Ground test equipment, keep people

with electronic components in them

away from the coil when the

electromagnetic coil is in use.

Epoxy/glue Toxic fumes, skin

irritation, eye irritation

Work in well ventilated areas to

prevent a buildup of fumes. Gloves

face masks, and safety glasses will be

worn at all times to prevent irritation.

Fire Burns, serious injury and

death

Keep a fire extinguisher in the lab. If

an object becomes too hot or starts to

burn, cut power and be prepared to use





a fire extinguisher.

Soldering Iron Burns, solder splashing

into eyes

Wear safety glasses to prevent damage

to eyes. Do not handle the soldering

lead directly only touch handle. Do not

directly hold an object being soldered.

Drills Serious injury, cuts,

punctures, and scrapes

Only operate tools under supervision

of team mates. Only use tools in the

appropriate manner. Wear safety

glasses to prevent debris from entering

the eyes

Dremel Serious injury, cuts, and

scrapes

Only operate tools under supervision

of team mates. Only use tools in the



the eyes

Hand Saws Cuts, serious injury Only use saws under supervision of

team mates. Only use tools in the



the eyes. Do not cut in the direction of

yourself or others.

Exacto Knives Cuts, serious injury,

death

Only use knives under supervision of


appropriate manner. Do not cut in the

direction of yourself or others.

Hammers Bruises, broken bones, Be careful to avoid hitting your hand





and serious injury while using a hammer.

Power Supply Electrocution, serious

injury and death

Only operate power supply under

supervision of team mates. Turn of

power supply when interacting with

circuitry.

Batteries Explode Eye irritation, skin

irritation, burns

Wear safety glasses and gloves. Make

sure there are no shorts in the circuit.

If a battery gets too hot stop using it an

remove any connections to it.

Improper Dress during

construction

Serious injury, broken

bones

Wear closed toe shoes, clothing that is

not baggy, and keep long hair tied

back.

Exposed construction metal Punctures, scrapes, cuts,

or serious injury

Put all tools band materials away after

use.

Neodymium Magnets Pinching, bruising, and

snapping through fingers.

Do not allow magnets to fly together

from a distance, do not play with

powerful magnets, keep free magnets

away from powered solenoids.

Iron Powders Inhaling, skin irritation Wear masks at all time, wear clothing

that protects sensitive skin areas. Keep

away from oxidizing agents.

Mineral Oil Toxic to inhale, ingest,

and irritable to skin

Label product, wear gloves while

working, keep body parts as protected

as possible.

Oleic Acid Eye irritation, skin

irritation, slight hazard

for inhaling ang ingesting

Wear safety glasses, wear gloves, label

product to remove confusion.





Magnetorheological Fluid Dangerous for inhaling,

ingesting.

Label mixture, keep sealed, keep

magnets away unless it is being used

for testing.

Table 38. Payload safety failure modes

Potential Failure Effects of Failure Failure Prevention

No power Experiment cannot be

performed

Check batteries, connections, and

switches

Data doesn't record No experimental data Ensure power is connected to the

payload computer and that all

connections are firmly secured

Magnetic field

interferes with flight

computer

No experimental data Shield the flight computer from

any EMF interference

Accelerometers/

Sensors

Record erroneous data Calibrate and test accelerometers

and all sensors

Water/Fluid damages

the camera

Stop operating, no

images, no data

Shield the camera from the fluid.

Magnetorheologial fluid

under an applied

magnetic force mixes

with water

Erroneous data. Create different compositions of

MR fluid and ensure that MR

fluid is sturdy.

Solenoids Experiment cannot be

performed, wires melt

Check connections, ensure over

heating will not occur during

testing




Potential Failure Effects of Failure Failure Prevention

Too much current goes

into the solenoids

The wires in the

solenoids get very hot

Make sure current is only pulsed

into the solenoids

Improper dress during

construction

Maiming, cuts,

scrapes, serious

injury.

Do not wear open toed shoes in

the build lab. Keep long hair tied

back. Do not wear baggy

clothing.

Avionics Chips or boards are

manufactured

incorrectly causing

equipment failures

and misfires

Test avionics operations, and

perform a flight test.

Personnel and Environmental Hazards 6.3.

As already mentioned in Section 6.1.1, the same methodology to identify and assess risks for

vehicle and payload safety will be used to identify hazards for constructing various flight and

testing components. A Material Safety Data Sheet (MSDS) is on hand for all materials used in

the construction of components, and team members have been briefed on best practices for

creating a safe workplace. Table 39 lists possible environmental safety concerns.




Table 39: Environmental Hazards, Risks, and Mitigation


Electrocution Serious Injury/death Do not touch wires that are hot and

not insulated. Wear rubber gloves

when the device is in operation.

Handle leads to the power supply

with care. Use low voltage settings

whenever possible.

Electromagnetic Fields Interfere with

electronic devices

inside the body

Ground test equipment, keep

people with electronic components

in them away from the coil when

the electromagnetic coil is in use.

Epoxy/glue Toxic fumes, skin

irritation, eye irritation

Work in well ventilated areas to

prevent a buildup of fumes. Gloves

face masks, and safety glasses will

be worn at all times to prevent

irritation.

Fire Burns, serious injury

and death

Keep a fire extinguisher in the lab.

If an object becomes too hot or

starts to burn, cut power and be

prepared to use a fire extinguisher.





Soldering Iron Burns, solder splashing

into eyes

Wear safety glasses to prevent

damage to eyes. Do not handle the

soldering lead directly only touch

handle. Do not directly hold an

object being soldered.

Drills Serious injury, cuts,

punctures, and scrapes

Only operate tools under

supervision of team mates. Only

use tools in the appropriate manner.


debris from entering the eyes

Dremel Serious injury, cuts,

and scrapes

Only operate tools under

supervision of team mates. Only

use tools in the appropriate manner.


debris from entering the eyes

Hand Saws Cuts, serious injury Only use saws under supervision of



glasses to prevent debris from

entering the eyes. Do not cut in the

direction of yourself or others.





Exacto Knives Cuts, serious injury,

death

Only use knives under supervision

of team mates. Only use tools in

the appropriate manner. Do not cut

in the direction of yourself or

others.

Hammers Bruises, broken bones,

and serious injury

Be careful to avoid hitting your

hand while using a hammer.

Power Supply Electrocution, serious

injury and death

Only operate power supply under

supervision of team mates. Turn of

power supply when interacting

with circuitry.

Batteries Explode Eye irritation, skin

irritation, burns

Wear safety glasses and gloves.

Make sure there are no shorts in the

circuit. If a battery gets too hot stop

using it an remove any connections

to it.

Improper Dress during

construction

Serious injury, broken

bones

Wear closed toe shoes, clothing

that is not baggy, and keep long

hair tied back.

Exposed construction

metal

Punctures, scrapes,

cuts, or serious injury

Put all tools band materials away

after use.





Neodymium Magnets Pinching, bruising, and

snapping through

fingers.

Do not allow magnets to fly

together from a distance, do not

play with powerful magnets, keep

free magnets away from powered

solenoids.

RF Interference with the

Recovery System

Pre-mature firing of the

ejection charges

potential causing

significant damage to

the Launch Vehicle,

payload, and all

supporting systems

RF Testing has verified that, at

maximum power output, the on-

board XBee transmitter will not

unintentionally ignite our e-

matches from excess RF radiation.

Maximum output power is limited

to 100 mW




7. Project Budget

Funding Overview 7.1.

In order to fund the 2012-2013 Competition year, the Ramblin’ Rocketeers have sought

sponsorships from academic and industry sources. The current sponsors of the Ramblin’

Rocketeers and their contributions can be found in Table 40. As of CDR, the Ramblin’

Rocketeers have received $5,700 in funding. Furthermore, the Team has also received a

dedicated room in which the Team can construct and store their rocket and non-explosive

components. All explosive components (i.e. black power) are properly stored in Fire Lockers in

either the Ben T. Zinn Combustion Laboratory or the Center for Space Systems Flight Hardware

Laboratory.

Table 40. Summary of sponsors for the Ramblin. Rocketeers

Sponsor Contribution Date

Unused Funds from 2011-2012 $1,000 Aug 2012

Georgia Space Grant Consortium $2,500 Sept 2012

Georgia Space Grant Consortium $500 Sept 2012

Georgia Space Grant Consortium $1,000 Dec 2012

Generation Orbit $300 Dec 2012

Georgia Tech

Student Government Association

$1,000 Feb 2013

Georgia Tech

School of Aerospace Engineering

$2,000 Mar 2013

ATK Travel Stipend $400 (est) Apr 2011

ATK Motor Stipend $200 (est) Apr 2011

Total $8,900

The team is currently pursuing the following sponsors: Virgin Galactic, Georgia Tech – College

of Engineering, Georgia Tech SGA, as well as private donations.




Current Sponsors 7.2.

Table 41 lists the current sponsors of the Ramblin’ Rocketeers and their contributions.

Table 41. List of current sponsors of the Ramblin' Rocketeers.

Sponsor Contribution

Georgia Space

Grant Consortium

Financial contribution for general project expenses

Financial contribution for Outreach-specific expenses

Financial contribution for RGEFP-related activities.

Advanced Circuits Manufacturing of the SIDES boards throughout the design

process

Generation Orbit Financial contributions for general project expenses.

Huff Performance Discounts on motor and motor hardware

Georgia Tech

Invention Studio

Professional machines and tooling to fabricate the launch vehicle

and payload components

Actual Project Cost 7.3.

7.3.1. FRR Budget Summary

Table 42 illustrates the budget breakdown as of the CDR Milestone. The summary is broken

down into four (4) main categories: Launch Vehicle, Flight Systems, Operations, and Motors.

The Launch Vehicle and Flight Systems categories are further broken down into two (2) sub-

categories: Flight Hardware and Testing. Operational expenses are broken down into four (4)

sub-categories: Safety, Generic Supplies, Tooling, and Physical Capital. Lastly, while motors are

specific to the Launch Vehicle subsystem, they are critical component to the architecture and as

such are tracked separately from the Launch Vehicle subsystem.




Table 42. FRR Project Budget Summary.

Category Amt Spent Amt Remaining

Launch Vehicle $ 1,707.78 $ 92.22

Motors $ 300.00 $ 700.00

Flight Systems $ 84.09 $ 915.91

Operations $ 202.25 $ 1,272.75

Testing/Dev $ 1,702.73 $ 47.27

7.3.2. System-Level Budget Summary

Figure 49 illustrates the system-level expenditure summary for Project LSIM at the FRR

milestone. Cost reduction techniques, such as proper resource utilization has resulted in lower

Flight Systems costs. It is important to note that both the Launch Vehicle and Flight Systems

include both Flight Hardware costs in addition to Test/Development costs. Additionally, Figure

50 illustrates the breakdown.

System Expenditure Breakdown

Launch Vehicle $ 661.75

Flight Systems $ 362.32

Operations $ 202.25

Motors $ 300.00

Testing/Development $ 1,702.73

Outreach $ 20.00

Total $ 3,249.10

Figure 49. System expenditure summary at CDR.




Sub-system Testing/Development

Breakdown

Aerodynamics $ 346.42

Structures $ 639.61

Recovery $ 0.00

Avionics $ 142.32

Payload/Ground

Testing $ 514.38

Ground Station $ 0.00

Total $ 1,642.73 Figure 50. Sub-system Testing/Development Breakdown.

7.3.3. Flight Hardware Expenditures

7.3.3.1. Flight Hardware Expenditure Overview

Figure 51 summarizes the overall expenditures for all Flight Hardware purchased up to the CDR

milestone. In order to account for uncertainties in motor price, $300 has been allotted for the

purchase of the flight motor. As illustrated by Figure 51, only hardware for the Aerodynamics

and Mechanical Integration sub-systems has been purchased.




Sub-system Flight Hardware

Breakdown

Aerodynamics $140.73

Structures $473.31

Recovery $102.29

Motors/Motor

Hardware $ 300.0

Mechanical

Integration $ 84.09

Electrical

Integration $ 0.00

Flight

Avionics $ 121.84

Flight Payload $ 156.39

Total $ 1,378.65

Figure 51. Sub-System Flight Hardware Breakdown.

7.3.3.2. Flight Hardware Cost Breakdown

Figure 52 lists the flight hardware breakout for Flight Systems. It is important to note that the

materials purchased for the Launch Vehicle flight hardware has not been used to fabricate any

parts, therefore no breakout is available at this time for the launch vehicle.




Flight Experiment

Item Description Unit Price Qty Cost

L.S.I.M. Hardware $ 156.39 1 $ 156.39

Propellant Simulant

(Water) $ 0.01 2 $ 0.02

2” Diameter PVC

Pipe $ 9.24 2 $ 18.47

Balance Solenoids $ 0.83 36 $ 30.00

Camera Assembly $ 50.00 2 $ 100.00

Piezo Vibration

Sensor $ 2.95 2 $ 5.90

Base Plate $ 0.50 2 $ 1.00

Payload Bottom $ 0.50 1 $48.00

Total Flight Experiment Costs $ 156.39

Flight Avionics

Item Description Unit Price Qty Cost

SIDES Network $ 416.70 1 $ 416.70

SIDES Board $ 33.00 5 $ 165.00

Electrical Harness $ 45.00 1 $ 45.00

ClockDrive Board $ 15.00 1 $ 15.00

LSIM Board $ 25.00 1 $ 25.00

Telemetry Board $ 15.00 1 $ 15.00

MasterIMU $ 20.00 1 $ 20.00

Strain Gage Board $ 15.00 2 $ 30.00

SIDES Node Battery $ 8.95 6 $ 53.70

LSIM Battery $ 48.00 1 $ 48.00

Total Flight Avionics Cost $ 416.70

Figure 52. Flight Systems flight hardware breakout.




8. Project Schedule

Schedule Overview 8.1.

The Mile High Yellow Jacket’s project is driven by the design milestone’s set forth by the USLI

Program Office. The design milestones are listed in Table 43. The project Gantt Chart for Project

L.S.I.M. – located in Appendix I – contains only high-level activities due to the unique launch

vehicle and payload designs. A more detailed Critical Path chart is located in Section 8.2.

Table 43. Design milestones set by the USLI Program Office.

Milestone Date

Proposal 26 SEP Team Selection 17 OCT

Web Presence Established 4 NOV PDR Documentation 28 NOV

PDR VTC 6 DEC CDR Documentation 23 JAN

CDR VTC 2 FEB FRR Documentation 26 MAR

FRR VTC 2-11 APR Rocket Week 18-21 APR

PLAR Documentation 7 MAY




Critical Path Chart: CDR to PLAR 8.2.

The critical path chart illustrated by Figure 6 demonstrates the highly integrated nature of Project

L.S.I.M. The critical path chart identifies:

• High Risk Tasks • Low-Moderate Risk Tasks • Earned Value Management (EVM)

Goal Tasks

• Looping Tasks • Critical and Alternate Paths • Major Inputs to Tasks

2012 – 2013 Georgia Tech Ramblin’ Rocketeers Flight Readiness Review

Figure 53. Critical Path Chart from CDR to PLAR




Schedule Risk 8.3.

8.3.1. High Risk Items

Two (2) items have been identified as “High Risk Items.” These are:

• Launch Vehicle Structure Design • Recovery System Design

Table 44 lists the mitigations for these items.

Table 44. Identification and Mitigations for High-Risk Tasks.

High-Risk Task Potential Impact on

Project L.S.I.M.

Mitigation

Launch Vehicle

Design, Fabrication,

& Testing

1) Schedule Impact

2) Budgetary Impact

3) Not qualifying for Competition Launch

1) Ensure personnel have direct and free access to experienced personnel on and off of the team.

2) Ensure personnel have knowledge on to effectively utilize simulation and analysis tools.

3) Ensure personnel have direct and free access to the simulation and analysis tools.

4) Ensure personnel are familiar with relevant fabrication techniques.

Recovery System

Design, Fabrication,

& Testing

1) Excessive kinetic energy during landing resulting in damage to the rocket.

2) Failure to deploy the drogue and/or main parachute resulting in a high energy impact with the ground destroying the Launch Vehicle.

1) Ensure Recovery System Lead has direct and free access to experienced personnel on and off the team.

2) Provide real-time feedback of the design decisions to ensure all recovery-related requirements are meet with at least a 5% margin wherever possible.

3) Ensure proper manufacturing techniques are utilized during the fabrication of the recovery system.




High-Risk Task Potential Impact on

Project L.S.I.M.

Mitigation

Verification of Field

Equations & Control

Logic

1) Unsuccessful flight demonstration

2) Flight Experiment does not function properly during flight

3) Flight Experiment encounters a flight anomaly that results in excessive draw and damage to the Flight Avionics, Power Supply, and/or Launch Vehicle

1) Develop multiple paths to achieve the end goal of developing thee robust control logic that is required for the successful demonstration of the Flight Experiment.

2) Ensure Flight Systems personnel have direct and free access to experienced personnel on and off of the team.

4) Ensure personnel have direct and free access to the simulation and analysis tools necessary for the development (and subsequent verification) of the control logic.

8.3.2. Low-to-Moderate Risk Tasks

The “low-to-moderate risk tasks” are considered to be those risks that pose a risk to either the

project schedule and/or project budget but little to no risk of not meeting the Mission Success

Criteria in Table 5. The risks and mitigations are provided in Table 45.




Table 45. Low to Moderate Risk items and mitigiations.

Risk Risk Level Potential Impact on

Project A.P.E.S.

Mitigation

Fabrication of

Launch Vehicle

Sections

Moderate 1) Schedule Impact 2) Budgetary Impact 3) Not qualifying for

Competition Launch

1) Ensure Manufacturing and Fabrication Orders (MFO’s) are sufficiently detailed for the task prior to starting any fabrication.

2) Ensure proper manufacturing techniques are observed during fabrication.

Full-Scale Launch

Vehicle Test Flight

Moderate 1) Schedule Impact 2) Budgetary Impact 3) Not qualifying for

Competition Launch

1) Ensure Launch Procedures are established practiced prior to any launch opportunity.

2) Have a sufficient number of launch opportunities that are in different geographical areas as to minimize the effects of weather on the number of launch opportunities.

Flight Computer

Fabrication Low

1) Budgetary Impact 2) Not able to collect in-

flight data

1) Ensure proper manufacturing techniques are observed during fabrication.

2) Ensure Manufacturing and Fabrication Orders (MFO’s) are sufficiently detailed for the task.

3) Descope custom board to COTS hardware

Ground Testing &

Control Logic

Development

Moderate 1) Schedule Impact 2) No Experimental Flight

Data is recorded prior to the Competition Launch.

1) Ensure personnel have direct and free access to experienced personnel on and off of the team.




9. Educational Engagement Plan and Status

Overview 9.1.

The goal of Georgia Tech’s outreach program is

to promote interest in the Science, Technology,

Engineering, and Mathematics (STEM) fields.

The Ramblin’ Rocketeers intend to conduct

various outreach programs targeting middle

school students and educators. The Ramblin’

Rocketeers will also have an outreach request

form on their webpage for educators to request

presentations or hands-on activities for their classroom.

Atlanta Makers’ Faire 9.2.

Ramblin’ Rocketeers had a booth at the Atlanta Makers Fair,

a fair in which various craftsman from the community and

Georgia Tech assemble to show off their accomplishments.

The intent of this program is to give clubs, organizations, and

other hobbyists the opportunity to show others their unique

creations and skills. The event is open to the entire Atlanta

community and had a large attendance this year. The

Ramblin’ Rocketeers booth had a display of our various rockets, as well as a station for children

to make their own paper rockets. Our booth had 10-15 middle school aged children attend and

participate in the paper-rocket activity.

FIRST Lego League and Tech Challenge 9.3.

FIRST is a series of international robotics competitions for students from 3rd -12th grades.

FIRST Lego League is an engineering competition designed for middle school children in which

Figure 54. Participation at the Atlanta Makers' Faire.

Figure 55: Previous FIRST Lego League outreach event.




they build and compete with an autonomous MINDSTORMS robot. Every year there is a new

competition centered on a theme exploring a real-world problem. FIRST Tech Challenge is a

robotics competition designed for students in middle and high school where the robots can be

18x18x18 inches at the start of each match.

This year the Ramblin’ Rocketeers have had an educational booth at a FIRST Lego League

Regional Competition at Wheeler High School which occurred on Saturday, December 8th.At the

booth students ranging from 3rd-8th grade were exposed to how lift is generated and participated

in building a paper rocket with a straw launcher that they could take with them. The event

reached 373 students, 295 of which were in the 4th-9th grade range, and 31 educators. Below in

Figure 56 and Figure 57 are pictures from this event.

Figure 56: FLL Regional Event at Wheel High School




Figure 57: FLL Regional Straw Rocket Activity

In addition to the FIRST Lego League Regional at Wheeler High School, the Ramblin’

Rocketeers are scheduled to have a booth at both the FIRST Tech Challenge Regional at

Wheeler Middle School on Saturday, January 19th and the FLL State Tournament at Georgia

Tech on Saturday, January 26th.




References

Simon, T. M., Reitich, F., Jolly, M. R., Ito, K., & Banks, H. T. (1998). Estimation of the E ective

Permeability in Magnetorheological Fluids. CRSC Technical Report CRSC-TR98-35,

NC State Univ.

The Dynamic Behavior of Liquids in Moving Containers: with applications to space vehicle

technology. All articles. Ed. H. Norman Abramson. NASA, Washington, D.C., 1966.

"Apogee Paramagnetic Oxygen Gas Experimental Electromagnetic Separator: Preliminary Design Review." Comp. Georgia Tech University Student Launch Initiative. Atlanta: 2009. Print.

Cheng, David. Field and Wave Electromagnetics. 1st ed. Reading, MA: Addison-Wesley Publishing Company, 1985. Print.

Niskanen, Sampo. OpenRocket vehicle Technical Documentation. 18 July 2011. Web.

Apke, Ted. "Black Powder Usage." (2009). Print. <http://www.info-central.org/?article=303>.

PerfecFlite. StratoLogger SL100 Users Manual. Andover, NH: Print. <www.perfectflite.com>.

Roensch, S. (2010). "Finite Element Analysis: Introduction." 2011, from http://www.finiteelement.com/feawhite1.html.

“G-10 Fiberglass Epoxy Laminate Sheet.” MATWEB.com. http://www.matweb.com/ search/datasheet_print.aspx?matguid=8337b2d050d44da1b8a9a5e61b0d5f85

"Shape Effects on Drag." NASA Web. 19 Nov. 2011. <http://www.grc.nasa.gov/WWW/k- 12/airplane/shaped.html>.

Cavcar, Mustafa. "Compressibility Effects on Airfoil Aerodynamics." (2005). Print.

http://www.finiteelement.com/feawhite1.html




Appendix I: Gantt Chart

ID Task Name Duration Start Finish Predecessors

1 Project L.S.I.M. 225 days Wed 8/1/12 Thu 6/6/13

2 RFP Released by NASA 30 days Wed 8/1/12 Tue 9/11/12

3 Proposal 22 days Wed 8/1/12 Fri 8/31/12

4 Team Formation 5 days Mon 8/20/12 Fri 8/24/12

5 Initial Rocket Design 20 days Wed 8/1/12 Tue 8/28/12

6 Flight Experiment Definition 20 days Wed 8/1/12 Tue 8/28/12

7 Internal Proposal Review 0 days Tue 8/28/12 Tue 8/28/12 6

8 Proposal Submitted 0 days Fri 8/31/12 Fri 8/31/12

9

10 Prelimary Design Review 71 days Fri 8/31/12 Thu 12/6/12

11 Launch Vehicle 31 days Fri 8/31/12 Sat 10/13/12

16 Flight Systems 31 days Fri 8/31/12 Fri 10/12/12

23 Project Level 71 days Fri 8/31/12 Thu 12/6/12

30 PDR Documentation Submitted 0 days Mon 10/29/12 Mon 10/29/12

31

32 Critical Design Review 68 days Sat 10/13/12 Mon 1/14/13

33 Launch Vehicle 57 days Mon 10/29/12 Mon 1/14/13

34 Recovery Detailed Design 51 days Mon 10/29/12 Mon 1/7/13 30

35 Structure Hardware Testing 19 days Mon 10/29/12 Thu 11/22/12 14,30

40 Full-Scale Launch Vehicle Fabrication 38 days Fri 11/23/12 Mon 1/14/13 39

44 Recovery Ground Testing 1 day Sat 1/12/13 Sat 1/12/13 41

45 Stability Analysis 51 days Mon 10/29/12 Mon 1/7/13 30

46 CFD of Launch Vehicle & Fin Can 20 days Mon 10/29/12 Fri 11/23/12 30

47 Development of stability model 40 days Mon 10/29/12 Fri 12/21/12 30

48 Verification of Scaled Test Launch 11 days Mon 12/24/12 Mon 1/7/13 46,47

49 Flight Systems 67 days Sat 10/13/12 Sat 1/12/13

50 Control System Preliminary Design 56 days Mon 10/29/12 Sat 1/12/13 30

51 Detailed Experiment Modeling 36 days Sat 10/13/12 Fri 11/30/12 20

52 Ground Testing 31 days Sat 10/13/12 Fri 11/23/12 20

53 Flight Systems Integration Plan 36 days Mon 10/29/12 Mon 12/17/12 30

54 Initial Ground Station Development 35 days Mon 10/29/12 Fri 12/14/12 30

55 Project Level 56 days Mon 10/29/12 Mon 1/14/13

56 Website Updates 56 days Mon 10/29/12 Sat 1/12/13 30

57 Outreach Events 56 days Mon 10/29/12 Sat 1/12/13 30

58 Completed 1st Draft of CDR 5 days Mon 12/17/12 Fri 12/21/12

59 Completed 2nd Draft of CDR 8 days Tue 1/1/13 Thu 1/10/13 58

60 Final editing of CDR Package 2 days Fri 1/11/13 Sat 1/12/13 59

61 CDR Documentation Submitted 0 days Mon 1/14/13 Mon 1/14/13

62

63 Flight Readiness Review 46 days Mon 1/14/13 Mon 3/18/13

64 Rocket 36 days Mon 1/14/13 Mon 3/4/13

65 Launch Vehicle Final Assembly 15 days Mon 1/14/13 Fri 2/1/13 61

66 Full-Scale Test Flight(s) 21 days Mon 2/4/13 Mon 3/4/13 65

67 Flight Systems 46 days Mon 1/14/13 Mon 3/18/13

68 Experiment Refinement 30 days Mon 1/14/13 Fri 2/22/13 61

69 Control System Refinement 46 days Mon 1/14/13 Mon 3/18/13 61

70 Integration of Flight Experiment & Avionics 15 days Mon 1/14/13 Fri 2/1/13 61

71 Project Level 46 days Mon 1/14/13 Mon 3/18/13

72 Website Updates 46 days Mon 1/14/13 Mon 3/18/13 61

73 Outreach Events 46 days Mon 1/14/13 Mon 3/18/13 61

74 FRR Documentation Submitted 0 days Mon 3/18/13 Mon 3/18/13

75

76 Rocket Week 34 days Thu 3/7/13 Mon 4/22/13

77 Fabrication of Flight Experiment 20 days Tue 3/19/13 Mon 4/15/13 68,69

78 Competition Launch Preparation 28 days Thu 3/7/13 Mon 4/15/13 66

79 Arrive in Huntsville 1 day Wed 4/17/13 Wed 4/17/13

80 Tour of MSFC 1 day Thu 4/18/13 Thu 4/18/13

81 Rocket Fair 1 day Fri 4/19/13 Fri 4/19/13

82 Competition Launch 2 days Sat 4/20/13 Mon 4/22/13

83

84 Post-Launch Assument Review Submitted 24 days Mon 5/6/13 Thu 6/6/13 82

Project L.S.I.M.

RFP Released by NASA

Proposal

8/28

8/31Proposal Submitted

Prelimary Design Review

10/29PDR Documentation Submitted

Critical Design Review

1/14CDR Documentation Submitted

Flight Readiness Review

3/18FRR Documentation Submitted

4/22Rocket Week

Post-Launch

29 5 12 19 26 2 9 16 23 30 7 14 21 28 4 11 18 25 2 9 16 23 30 6 13 20 27 3 10 17 24 3 10 17 24 31 7 14 21 28 5 12 19 26Aug '12 Sep '12 Oct '12 Nov '12 Dec '12 Jan '13 Feb '13 Mar '13 Apr '13 May '13 J

Task Split Progress Milestone Summary Project Summary External Tasks External Milestone Deadline

Page 1

Project: 2012 - 2013 USLI Gnatt ChartDate: Mon 3/18/13




Appendix II: Launch Checklist

Pre-Launch

Performer Inspector

Packing The night before launch go through Launch Vehicle

Packing List and put all items in a designated spot.

The morning of launch go through Launch Vehicle Packing List and ensure all items are still there.

Load the vehicle(s)

Launch

Avionics On Prepare Payload Bay Ensure batteries and switches are wired to the altimeters

correctly.

Ensure batteries, power supply, switch, data recorder and pressure sensors are wired correctly.

Install fresh batteries into battery holders and secure with tape.

Test the altimeters.

Altimeter In Circuit Out of Circuit

Altimeter 1

Altimeter 2

Insert altimeter and payload into the payload bay. Connect appropriate wires. Verify payload powers on correctly and is working properly. If it is not, check all wires and connections.

Turn off payload power. Arm altimeters with output shorted to verify jumper settings. This is to check battery voltage and continuity.

Disarm altimeter, un-short outputs.




Assemble Charges Test e-match resistance and make sure it is within spec. Remove protective cover from e-matches.

Measure amount of black powder determined in testing. Put e-matches on tape with sticky side up.

E-match Resistance

E-match 1

E-match 2

E-match 3

E-match 4

Pour black powder over e-matches. Seal tape. Re-test e-matches.

Check Altimeters Ensure altimeter is disarmed.

Connect charges to altimeter bay. Turn on altimeter and verify continuity. Disarm altimeters.

Altimeter 1 Altimeter 2

OFF ON

Pack Parachutes




Connect drogue shock cord (long side) to booster section and altimeter bay (short side)

Fold excess shock cord so it does not tangle. Add Nomex cloth to ensure only the Kevlar shock chord is exposed to ejection charge.

Insert altimeter bay into drogue section and secure with shear pins.

Pack main chute. Attach main shock cord to payload bay (long side to nose

cone).

Fold excess shock cord so it does not tangle. Add Nomex cloth under main chute and shock cord ensuring that only the Kevlar part of the shock cord will be exposed to the ejection charge

Connect shock cord to nose cone, install nose cone and secure with shear pins.




Post Launch

Recovery Recover launch vehicle, document landing.

Disarm altimeter(s) if there are unfired charges. Disassemble launch vehicle, clean motor case, other parts, inspect for damage.

Record altimeter data. Download payload data.

Assemble Motor Follow manufacturer's instructions.

Put on safety glasses and gloves. Do not get grease on propellant or delay. Do not install igniter until at pad. Install gasket on top of motor. Install motor in launch vehicle. Secure positive motor retention.

Final Prep Turn on payload via a switch and start stopwatches.

Install skin. Inspect launch vehicle. Check CG to make sure it is in safe range; add nose weight if necessary.

Bring launch vehicle to the range safety officer (RSO) table for inspection.

Bring launch vehicle to pad, install on pad, verify that it can move freely (use a standoff if necessary).

Arm altimeters via switches and wait for continuity check for both.

Install igniter Touch igniter clips together to make sure they will not fire igniter when connected.

Make sure clips are not shorted to each other or blast deflector. Return to front line.

Launch Stop the stopwatches and record time from arming payload and

launch.

Watch flight so launch vehicle does not get lost.




Trouble Shooting

Test Problem Control & Mitigation

Power on payload Payload does not

power on

Check batteries have sufficient charge, check wires

are connected correctly

Check E-match

resistance

E-match resistance

does not match

required specifications

Replace e-match before use

Power on altimeters Altimeters do not

power on

Check batteries have sufficient charge, check wires

are connected correctly

Check for altimeter

continuity after

installing e-matches

No continuity Check wires are connected correctly

Launch Rocket Engine does not fire

Disconnect power, ensure igniter clips are not

touching, ensure power is reaching clips ,ensure

motor is assembled correctly




Appendix III: Science Overview

Ferromagnetism and MR fluid response

Scientific Background and Mathematical Modeling

To the end of accomplishing the goals of the LSIM experiment, some theoretical research and

work must be accomplished in tandem with experimentation. A passive or active control system

is to be developed in order to move the simulated propellant to its desired location with the

magnetorheological (MR) fluid. To model the behavior of the simulant-MR fluid system,

equations are being researched, modified, and developed in order to calculate magnetic fields

and forces, to govern the properties of MR fluids, and to model system dynamics. In addition to

equations, qualitative research has been done in the literature concerning MR fluids to suggest

approaches that may be taken during experimental testing.

Magnetic fields

The forces on the MR fluid that will be transmitted to the simulant will depend largely on the

magnetic fields that are applied to the fluid. Control of currents in a solenoid will allow for

precise control of the fields. Last year, it was derived and also confirmed in the literature that the

exact magnetic H field from a current loop in spherical coordinates, with the loop centered at the

origin in the xy -plane and counterclockwise current, is as below (θ denotes azimuth angle):

𝐻𝑟 =

𝐶𝑅2 cos 𝜃𝛼2𝛽

𝐸(𝑘2)

𝐻𝜃 =𝐶

2𝛼2𝛽 sin𝜃[(𝑟2 + 𝑅2 cos 2𝜃)𝐸(𝑘2) − 𝛼2𝐾(𝑘2)]

where K and E are complete elliptic integrals of the first and second kinds, respectively, and

𝛼2 = 𝑅2 + 𝑟2 − 2𝑅𝑟 sin𝜃, 𝛽2 = 𝑅2 + 𝑟2 + 2𝑅𝑟 sin𝜃, 𝑘2 = 1 − 𝛼2 𝛽2⁄ , and 𝐶 = 𝐼 𝜋⁄ . I is the

loop current, R is its radius, and r is the distance from the origin to the point of measurement. A

solenoid simply consists of several such current loops, with the fields adding vectorally. While

the above expressions are extremely nonlinear and difficult to analyze or work with, they may be

simplified as needed, or modeled using a computer.




Magnetic forces

After calculating the magnetic fields, in order to predict the motion of the MR fluid and simulant

in the container, the forces on the MR fluid due to the field must to be calculated. In any

material, the movement of atomic charges such as electrons causes the atoms to behave as

microscopic magnetic dipoles, experiencing forces in magnetic fields. The magnetization vector

M at a point in the material is defined as the volume “density” of magnetic dipole moment, i.e.

𝐌 = lim∆𝑣→0

∑𝐦𝑘

∆𝑣

Each 𝐦𝑘 is the magnetic moment of the kth atom in volume ∆𝑣, and the sum is over all atoms. M

depends on the magnetic field H at a point, and flux density B depends on the field, as follows:

𝐌 = χ𝑚𝐇

𝐁 = 𝜇0(𝐇 + 𝐌) = 𝜇0𝐇(1 + 𝜒𝑚) = 𝜇0𝜇𝑟𝐇 = 𝜇𝐇

where χ𝑚 is the material’s magnetic susceptibility, 𝜇𝑟 is its relative permeability, and 𝜇 is the

absolute permeability. It is assumed that χ𝑚, and hence 𝜇 and 𝜇𝑟, are approximately constant for

the MR fluid. This is a very valid assumption that greatly simplifies analysis, given that the fields

are not extremely large, as is evidenced in Figure 58 below taken from a paper by Simon et al.

Figure 58: Plot of B field magnitude in MR fluid versus magnitude of vector 𝝁𝟎𝑯, for iron volume concentrations of 10, 20, and 30 percent




The force on a magnetic material can be determined by summing the forces on the dipoles in the

material due to the field that it is placed in. The force on a magnetic dipole m in field B is

𝐅 = 𝛁(𝐦 ∙ 𝐁)

Let V be the volume of a very small region of the MR fluid in which M is approximately

constant. Then, letting 𝐦 = 𝐌𝑉 = χ𝑚𝑉𝐇 = χ𝑚𝑉µ𝐁, the force on the region is

𝐅 = 𝛁�χ𝑚𝑉µ

𝐁 ∙ 𝐁� =2χ𝑚𝑉µ

𝐁 ∙ 𝛁(𝐁)

Using equations (1), (2), and (5) for the H and B fields of a current loop, it can be seen that the

force on each small region, and hence on the whole fluid, should be directly proportional to the

square of the current. In addition, 𝐁 ∙ 𝛁(𝐁) may be calculated using equations (1) and (2). These

equations will be further developed to better understand response of the MR fluid and simulant.

MR fluid rheological properties

In addition to translational movement, which is governed by the preceding equations, MR fluids

experience large increases in yield strength in the presence of magnetic fields. This is desirable

for the LSIM system, as otherwise the sloshing propellant simulant would simply shear through

the MR fluid barriers with little resistance. It is desired to characterize the rheological properties

of MR fluid to understand how much resistance to movement the simulant will experience.

More precisely, MR can be modeled fairly closely as a Bingham plastic, a common example of

which is toothpaste. A Bingham plastic does not start flowing until a certain point of yield shear

stress, after which it behaves similarly to a viscous liquid. The equation governing the shear

stress of an ideal Bingham plastic, and so to model the MR fluid for future analysis, is

𝜏 = 𝜏𝑦𝑖𝑒𝑙𝑑(𝐇) + 𝜂

𝑑𝑣𝑑𝑛

for τ > 𝜏𝑦𝑖𝑒𝑙𝑑(𝐇)

𝜏𝑦𝑖𝑒𝑙𝑑(𝐇) is the yield shear stress of the MR fluid, and is larger for stronger H fields. η is the

flow viscosity after shear, and 𝑑𝑣𝑑𝑛

is the velocity gradient in the direction normal to the plane of

shear. This relation is shown on the next page in Figure 59, compared to a Newtonian fluid.




Figure 59: Shear stress of ideal Bingham plastic (and MR fluid model) versus shear rate 𝒅𝒗𝒅𝒏

, compared to ideal Newtonian liquid

Hence, if the simulant exerts such a force that MR fluid flow begins occurring, the shear stress

between layers of the MR fluid should increase, keeping the simulant comparatively restrained

until it settles again. If the need arises to decrease the yield shear stress for a given magnetic

field, such as to make the MR fluid flow more easily, replacing a percentage of microscale

ferroparticles with nanoscale particles can decrease the yield stress. Further research is still

required to find the relationship between the yield strength and magnetic field, which will allow

control of the yield stress acting against the simulant. However, the key observation is that there

is little to no MR fluid flow below some certain shear stress, for a given magnetic field H.

System Dynamics

While research on the physical properties and behavior of MR fluids is ongoing, basic system

dynamical modeling has already been started with variable parameters that will be determined

from theory and experimentation in the future. The fluid and MR fluid mixture is assumed to

operate roughly as a system with a spring, damper, and mass, where the driving force is the

solenoid. The fluid is considered the mass, whose motion is restrained by a spring and damper,

and driven by the MR fluid actuated by the solenoid. All system elements lie on the same x -

axis, with the solenoid axis coinciding. The dynamical equation of motion in this case is




𝑚�̈� = 𝐹𝑠𝑜𝑙𝑒𝑛𝑜𝑖𝑑 − 𝑘𝑥 − 𝑏�̇�

Where m is the mass of the fluid, k and b are unknown damping placeholder constants, and x is

the position of the simulant relative to some point. After some manipulation, the dynamical

equation for the response of the fluid becomes:




Understanding Slosh Damping

Fluid dynamics and hydrodynamic regimes of expected slosh

In considering the liquid slosh, the flight regime of the vehicle is extremely important. While the

experiment aims to approximate a spacecraft by manipulating MR fluid during microgravity to

dampen water slosh, the realities of atmospheric flight will limit the applicability of launch

vehicle test results. The extent of these flight regime limitations is revealed by three key

similarity parameters: the Weber (We) number, the Froude (Fr) number, and the Bond (Bo)

number. These three parameters measure the ratio of inertial to capillary forces, the effect of

gravitational body forces relative to inertial forces, and the relative magnitudes of gravitational

and capillary forces respectively. Finally, an understanding of the potential flow of sloshing

fluid is necessary to understand the motion of fluid inside a vehicle.

Flight regime

However, an estimate of the flight regime of the launch vehicle near apogee must first be known.

To better understand this flight regime and to confirm the microgravity requirements pulled from

previous team documents, a first-order analysis of the launch vehicle’s flight was computed.

Neglecting drag and assuming 2-D projectile motion with instantaneous acceleration from a

rocket motor, the flight profile of the launch vehicle was estimated and the characteristics of the

0.1-Gee requirement from the 2009 Georgia Tech team – 0.1-Gee being the definition of the

microgravity threshold for the purposes of the experiment – were examined. The results of this

simplified analysis are presented graphically in Figure 60.




Figure 60: Microgravity time as a function of launch angle from horizon

In Figure 60, the microgravity time, or ∆𝑡𝑚𝑖𝑐𝑟𝑜, was computed using equation (8).

∆𝑡𝑚𝑖𝑐𝑟𝑜 =

𝑉0sin (𝛼)𝑔

√0.05 + 1.5 (8)

In equation (8), the 1.5 s addition represents the time from apogee to chute deployment, which

by representation in Figure 60 is always less than the other half of the equation for launch angles

between 60 and 90 degrees. The drogue chute deployment therefore represents the bounding

time for the experiment operation in the mission profile. From the flight profile, a velocity

corresponding to 0.1-Gee and a maximum height can be calculated. These variables can be used

for the computation of similarity parameters, as well as comparison numbers to judge the validity

of the flight profile and microgravity estimates.

Two comparison measures will now be observed. Among the simplest environments for creating

microgravity is the free-fall drop test. This test provides microgravity times approximated by

equation (9) valid to heights of 20 m with atmospheric drag. Equation (9) is nonspecific with

regards to the accelerations achieved, however these are estimated by Reynolds and Satterlee (p.

435, Dynamic Behavior) to be between 10−7 and 0.2.




∆𝑡𝑚𝑖𝑐𝑟𝑜 =1

2.2√ℎ (9)

The predicted times from equation (9) and from the flight profile are given in Figure 60.

Table 46: Microgravity times for fall heights

Height (m) Microgravity time (s)

3733.96 (free-fall) 27.78

1609 (free-fall, target altitude) 18.23

3733.96 (90° launch angle) ~7.13

Adjusting for the 1.5 second chute deployment, the 90° launch microgravity time appears to be

about one half the time given for free-fall from the same maximum height – given that the max

Gee loading specified in the reference is 0.2, or twice the Ramblin’ Rocketeers’ requirement, this

difference appears to be acceptable for a bounding and ideal case. Of course, accelerations due

to aerodynamic forces will requirement additional modeling and adjustment.

Similarity parameters

Table 47 presents the similarity parameters relevant to the LSIM experiment calculated for the

propellant simulant, water (30 °C). The Weber, Bond, and Froude numbers are considered here.

These numbers provide an indication of the hydrodynamic regime – these regimes




Figure 61: Slosh regimes and similarity parameters 6F

7 for microgravity are illustrated in Figure 61. The Reynold’s number is also included for

comparison; although for this experiment the number itself is not as significant as long as the

regime described by different test configurations is similar, i.e. all turbulent, all laminar, etc. A

potential source of error in these computations is use of the launch vehicle velocity rather than

the relative velocity of the fluid in the tank. The Weber, Froude, and Reynold’s numbers are

affected by this choice, which is yet to be validated.

Table 47: Similarity parameters for simplified flight profile of the launch vehicle

Number Equation Value

Bo 𝜌𝑔𝐿2/𝜎 980

We 𝜌𝑢2𝐿/𝜎 1.37𝑥107

Fr We/Bo 1.4𝑥104

Re 𝜌𝑢𝐿/𝜇 2.023𝑥107

These parameters will allow verification and comparison of ground tests with the launch vehicle

test and RGEFP, vis-à-vis actual spacecraft and launch vehicles.

7 Reynolds, William C. and Hugh M. Satterlee. “Liquid Propellant Behavior at Low and Zero G”. The Dynamic Behavior of Liquids… p. 390. Ref Appendix XXX




Modeling Slosh

In order to approximately predict the behavior of the propellant in the sloshing modes expected

at apogee of our launch vehicle, a mathematical model has been developed. The aim of this

model is solely to predict sloshing behavior in the absence of dampening; this is an intermediate

step to modeling the effects of the MR fluid dampening. Predicting the exact distribution and

dynamics of all sloshing fluid within the container, however, would require a large amount of

complexity, which would only be exacerbated by attempting to add the effects of the MR fluid

later on. Therefore, for analysis to be feasible, a much simpler model is proposed.

Figure 62. Schematic and free-body diagram of slosh dynamic model

The fluid is modeled schematically as shown above in Figure 62. It is assumed that the center of

mass of the fluid in the tank behaves roughly as an object of mass m attached to a pendulum-

spring of spring constant k, with additional dampening effects represented by viscous dampers of

constants b1 and b2. Also, let L be the original length of the pendulum-spring with no forces

applied, and ∆L the the amount the spring is stretched from length L (so that negative ∆L implies

compression). It is anticipated that the majority of sloshing will be longitudinal, so significant

vertical motion can be expected of the fluid. Therefore, in our model, the spring may experience

appreciable compression. However, a much smaller amount of lateral sloshing is predicted, so in

the model, the angle θ may be assumed to be small. Therefore, throughout this analysis, it will be

assumed that sin θ ≈ θ sufficiently closely for the corresponding substitution to be justified.




The equations of motion can be written in the x and y directions, respectively, as follows:

−𝑘∆𝐿 sin𝜃 − 𝑏1�̇� ≈ 𝑘∆𝐿𝜃 − 𝑏1�̇� = 𝑚�̈� (1)

𝑘∆𝐿 cos𝜃 − 𝑏2�̇� − 𝑚𝑔 ≈ 𝑘∆𝐿 − 𝑏2�̇� − 𝑚𝑔 = 𝑚�̈� (2)

The equations to be developed will depend on state variables θ, �̇�, ∆L, and ∆�̇�. Let the origin of

the coordinate system be at the center of rotation of the pendulum. Then, it is known that

𝑥 = (𝐿 + ∆𝐿) sin𝜃 ≈ (𝐿 + ∆𝐿)𝜃 (3)

𝑦 = −(𝐿 + ∆𝐿) cos 𝜃 ≈ −(𝐿 + ∆𝐿) (4)

�̇� = �∆�̇��𝜃 + (𝐿 + ∆𝐿)�̇� (5)

�̇� = −∆�̇� (6)

�̈� = �∆�̈��𝜃 + 2�∆�̇��̇� + (𝐿 + ∆𝐿)�̈� (7)

�̈� = −∆�̈� (8)

where equations 5 through 8 are found by repeatedly differentiating equations 3 and 4. First of

all, substituting equations 6 and 8 into equation 2 and rearranging, it is readily found that

∆�̈� = 𝑔 −

𝑏2𝑚�Δ�̇�� −

𝑘𝑚∆𝐿 (9)

Note that this equation is independent of θ, and is the same as a one-dimensional spring-mass-

damper system.

The analysis and results from substituting equations 5 and 7 into equation 1 are significantly

more complicated. First of all, carrying out this substitution and rearranging,

�̈� = �

−𝑘∆𝐿 − 𝑏1�∆�̇��𝑚(𝐿 + ∆𝐿) −

�∆�̈��(𝐿 + ∆𝐿)� 𝜃 − �

𝑏1𝑚

+2�∆�̇��

(𝐿 + ∆𝐿)� �̇� (10)

Next, substituting equation 9 for ∆�̈� in equation 10, it is found that

�̈� = �

(𝑏2 − 𝑏1)�∆�̇��𝑚(𝐿 + ∆𝐿) −

𝑔(𝐿 + ∆𝐿)� 𝜃 − �

𝑏1𝑚

+2�∆�̇��

(𝐿 + ∆𝐿)� �̇� (11)

Therefore, as k, b1, b2, m, g, and L are constants, it is seen that �̈� is a function of the state

variables. Write �̈� = 𝑓�𝜃, �̇�,𝛥𝐿,𝛥�̇�� , where 𝑓:ℝ4 → ℝ is differentiable at any point




�𝜃, �̇�,𝛥𝐿,𝛥�̇�� such that ∆L≠-L, due to continuity of the partial derivatives of f at those points.

The relation ∆L = -L only occurs for the spring being compressed into a flat piece, which

corresponds to all fluid molecules touching the top surface of the tank; both of these are

impossible scenarios. Therefore, it is possible to linearly approximate f close to any point of

interest, for easier analysis. Initially, the relation is linearized for points near the equilibrium

point of the system, which is (0, 0, mg/k, 0). In this case, defining 𝛿𝐿 = 𝛥𝐿 −𝑚𝑔/𝑘,

�̈� = 𝑓�𝜃, �̇�,𝛥𝐿,𝛥�̇��

≈ 𝑓 �0,0,𝑚𝑔𝑘

, 0� + ∇𝑓 �0,0,𝑚𝑔𝑘

, 0� ∙ �𝜃, �̇�, 𝛿𝐿,𝛥�̇�� (12)

�̈� = 0 +

𝜕𝑓𝜕𝜃�𝑒𝑞𝜃 +

𝜕𝑓𝜕�̇��𝑒𝑞�̇� +

𝜕𝑓𝜕𝛥𝐿

�𝑒𝑞𝛿𝐿 +

𝜕𝑓𝜕𝛥�̇�

�𝑒𝑞𝛥�̇� (13)

Finally, evaluating the partial derivatives, it is found that close to (0, 0, mg/k, 0),

�̈� = �−

𝑔𝑘(𝑘𝐿 + 𝑚𝑔)� 𝜃 − �

𝑏1𝑚� �̇� (14)

Therefore, relations for both ∆�̈� and �̈� have been found only in terms of the four state variables

𝜃, �̇�,𝛥𝐿, and 𝛥�̇� , assuming that the values of 𝜃, �̇�, 𝛿𝐿, and 𝛥�̇� are small. Using the above

equations, whether the linear approximations (9) and (14) or the more precise but complicated

form (11), further analysis by hand or by computer should yield information as to how the

system should approximately behave in the absence of the MR fluid baffles. Further

development should also approximate system dynamics in the presence of the baffles, allowing

rough predictions to be made as to how the MR fluid baffles may impact fluid slosh.

Unifying the LSIM theories

Finally, it is necessary to connect the response of MR fluid to a magnetic field with the damping

of slosh. According to Dynamic Behavior the damping of slosh for a ring baffle is dependent on

the baffle cross-section7F

8. For containers of constant geometry and liquid at constant rest height

in gravity dominated slosh, empirical relationships have been illustrated between the geometry of

8 Abramson, H. Norman and Sandor Silverman. “Damping of Liquid Motions and Lateral Sloshing”. The Dynamic Behavior of Liquids… p. 109. Ref Appendix XXX.




a rigid baffle and the damping ratio vs. wave amplitude of slosh 8F

9. However, an effect not present

in the dynamic analysis above is the container geometry. In the case of LSIM where the liquid

height will be greater than the container radius, the damping coefficient in lateral slosh is given

by:

𝛿 = 4.98𝜈1/2𝑅−3/4𝑔−1/4 (1)

Where ν is the kinematic viscosity, R the container radius, g the acceleration of gravity. This

equation9F

10, along with curve fitting with the help of tables and plots given in Dynamic Behavior,

provides a means to experimentally determine the damping coefficients needed with given MR

fluid response to calculate the slosh dynamics.

9 P. 109 10 P. 110




Appendix IV: Ground Test Plan

Goals

The LSIM ground test data will provide the basis for empirical modeling of magnetorheological

fluid as a damper for liquid sloshing. All actions will be incremented to allow for a detailed

model for extrapolation and interpolation of the data for future flight control systems.

Ground Test Goal Ground Test Goal Definition

1 Create MR Fluid

2 Calibrate Sensors

3 Determine force of MR Fluid

4 Develop model for solenoid control

5 1-G slosh dampening

Test Sequence 1 - Creating MR Fluid

MR fluid will be created using different compositions of iron powder, mineral oil, and surfactant.

The iron powder will make up about 74-76% of the mixture's mass. Mineral oil will make up 20-

22% of the total mass, and the surfactant will make up the remaining 1-4%. Water is then added

to test the time to mixture separation and solenoids. Each mixture will be preliminarily tested by

neodymium magnets. The mixture will qualify as a successful batch if MR fluid under the

influence of an applied magnetic field prevents the leakage of water.

Test Sequence 2 - Characterize the shear stress of MR fluid

In characterizing MR fluid, the team will utilize a two-plate test for measuring the MR fluid's

force and viscosity with and without a magnetic field acting upon the MR fluid. This test was

chosen because of its simplicity; other tests such as a barometer test were considered for

measuring the MR fluid's viscosity and force, but they turned out too complicated to realize.




The two-plate test consists of two plates: a bottom plate, which is fixed to the ground and a top

plate, which is free to move. A load sensor will be placed on the top plate to measure the reaction

force that is generated. The plates used must not be strongly magnetic; thus, the two current

choices are wood or aluminum.

A control test will be performed by just having two plates together with a load sensor on the top,

moving plate to calculate the force by the plates themselves. For accurate and consistent results,

an automated pulling device will be used to pull the top plate. Once a control has been measured,

MR fluid will be placed between the two plates and the same procedure will repeat with and

without the MR fluid under a magnetic field. These tests will characterize the force that MR fluid

will generate when it is under a magnetic field and when it is free of a magnetic field

Test Sequence 4 - Developing solenoid control

Knowing the MR fluid shear stress properties will help determine the size and strength of the

solenoid used for flight testing. This will also enable the group to decide on what type of control

can be used on the solenoid. At the moment, an open loop control is considered.

If better coupling can be achieved between sensors and actuators, closed loop control may be

considered.

Test Sequence 5 – 1-G Slosh dampening

A vibration rig will be constructed such that several frequencies of vibration approximating those

experienced by the launch vehicle will be exerted on the ground test rig. Using similarity

parameters, the data gained from this experiment will allow predictions for dampening

performance of the controlled MR fluid during the microgravity period.




Appendix V: Science MFOs and Drawings

Bench Stand

Take one 1/16’’ 12’’ x 12’’ acrylic plate and laser cut out a square that is 20cm x 20cm.

Figure 63. Base Plate




Take the remaining two 12’’ x 12’’ plates and cut a rectangle that is 20cm x 22cm. Laser cut a

5.3cm diameter centered in the middle of the 20cm x 22cm acrylic plates as seen in Figure 2.

Figure 64. Second and Top Plate




Take four 30cm Maker Bars and place them vertically 20 cm apart in the shape of a square.

Attach horizontally a 20cm Maker Bar 2 cm above each 30cm Maker Bar base and secure them

with 90 Degree Maker L Brackets as shown in Figure 3 and 4. This will be done on each side of

the square. Figure 3 shows a side view of two 30 cm Maker Bars attached together by a 20cm

Maker Bar horizontally.




Figure 65. Side view of main structure

Figure 66. Trimetric view of main structure




Attach four flat “L” Maker Beam brackets on the 20 cm horizontal bars as show in in Figure 5.

The short end of the “L” bracket will be placed 1cm from the end of the 20cm Maker Bar. This

will hold the 20cm x 20cm solid acrylic base plate in place.

Figure 67. Top view of structure with 90 Degree L Brackets




On each vertical 30cm Maker Bar, attach the base of a 90 degree bracket 2 cm above the 20cm

horizontal Maker Bar (5cm from the bottom of the vertical 30cm Maker Bar) as shown in Figure

6. Do this for each vertical 30 cm Maker Bar facing inward into the square. These four 90 degree

brackets will hold the second acrylic plate.

Figure 68. Side view of 90 Degree Brackets




Four cm above the top of each 90 degree bracket (9cm from the bottom of the vertical 30cm

Maker Bar), place the base of another 90 degree bracket as shown in Figure 6. These four 90

degree brackets will hold the third acrylic plate.

Place the 20cm x 20cm acrylic base plate that has no hole cut in it over the four, flat “L” Maker

Beam brackets. The place the other two 20cm x 22cm acrylic plates that have holes in them on

the 90 degree brackets creating three layers of acrylic plates as shown in Figure 7.

Figure 69. Test Structure with Base Plates

Place the 100 mL beaker through the acrylic plates and onto the base acrylic plate.




Solenoids

Take each 20cm Soft Iron Core rod and cut them up into 6cm long rods. There will be 24 of

these rods after cutting.

Wrap a piece of paper or two around the Iron cores to provide insulation.

Take 100 feet of 32 gauge magnet wire and wind as many even turns possible around each core,

counting the turns. Make sure both ends of the magnet wire are free to be used (i.e. don’t start

winding the beginning of the magnet wire).

Place 8 magnets radially around the beaker on the first, base plate. Do the same for the other two

plates.




Appendix VI: Altimeter Wiring Harness Schematic

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