2012 – 2013 georgia tech ramblin’ rocketeers preliminary … gt usli pdr.pdf · 2012. 10....

2011 2012 – 2013 Georgia Tech Ramblin’ Rocketeers

Preliminary Design Review

Georgia Institute of Technology 2 of 128 Ramblin’ Rocketeers

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Table of Contents

1. Introduction ........................................................................................................................... 10 School Information and NAR Section Contacts ............................................................ 10 1.1. Work Breakdown Structure ............................................................................................ 10 1.2. Launch Vehicle Summary .............................................................................................. 11 1.3.

Overview ................................................................................................................. 11 1.3.1. Changes since PDR ................................................................................................. 12 1.3.2.

Payload Summary .......................................................................................................... 12 1.4. Overview ................................................................................................................. 12 1.4.1. Changes Since PDR ................................................................................................ 12 1.4.2.

2. Project L.S.I.M. Overview ..................................................................................................... 13 Mission Statement .......................................................................................................... 13 2.1. Requirements Flow Down .............................................................................................. 13 2.2. Mission Objectives and Mission Success Criteria ......................................................... 14 2.3. System Requirements Verification Matrix (RVM) ........................................................ 14 2.4. Mission Profile ............................................................................................................... 22 2.5.

3. Launch Vehicle ...................................................................................................................... 24 Overview ........................................................................................................................ 24 3.1.

Mission Criteria ...................................................................................................... 25 3.1.1. System Design Overview ............................................................................................... 25 3.2. Recovery System ............................................................................................................ 30 3.3.

Parachute Dimensions ............................................................................................. 32 3.3.1. Drift Profile Analysis .............................................................................................. 33 3.3.2. Kinetic Energy of Launch Vehicle ......................................................................... 34 3.3.3. Ejection Charges ..................................................................................................... 35 3.3.4. Testing..................................................................................................................... 36 3.3.5.

Structure ......................................................................................................................... 36 3.4. Payload Section ....................................................................................................... 38 3.4.1. Booster Section ....................................................................................................... 43 3.4.2. Mass Breakdown ..................................................................................................... 45 3.4.3. Section Integration .................................................................................................. 45 3.4.4. Manufacturing ......................................................................................................... 45 3.4.5. Future Testing ......................................................................................................... 46 3.4.6.

Launch Vehicle Performance Analysis .......................................................................... 47 3.5. Fin Design ............................................................................................................... 47 3.5.1. Nose Cone ............................................................................................................... 52 3.5.2.




Motor Selection ....................................................................................................... 53 3.5.3. CP and CG .............................................................................................................. 54 3.5.4. Altitude Predictions ................................................................................................ 56 3.5.5. Fabrication and Materials ....................................................................................... 56 3.5.6. Future Testing and Analysis ................................................................................... 57 3.5.7.

Vespula Mk II Mass Breakdown .................................................................................... 57 3.6. Interfaces and Integration ............................................................................................... 59 3.7.

Interface with the Ground ....................................................................................... 59 3.7.1. Interface with the Ground Launch System ............................................................. 59 3.7.2.

Launch Vehicle Operations ............................................................................................ 60 3.8. Launch Checklist .................................................................................................... 60 3.8.1.

4. Flight Experiment .................................................................................................................. 61 Introduction to the Experiment and Payload Concept Features & Definition ............... 61 4.1. Overview of the Experiment .......................................................................................... 62 4.2.

Hypothesis and Premise .......................................................................................... 62 4.2.1. Experimental Method and Relevance of Data ........................................................ 62 4.2.2. Accomplishments Since Proposal ........................................................................... 63 4.2.3. Ground Test Plan .................................................................................................... 63 4.2.1.

Overview ......................................................................................................... 63 4.2.1.1. MR Fluid Creation and Validation of Theory ................................................. 64 4.2.1.2. MR Fluid Shear Stress Characterization: Two Plate Test .............................. 65 4.2.1.3. Visible detection of motion: IR reflectance .................................................... 66 4.2.1.4. Working Ground Model .................................................................................. 67 4.2.1.5. Sensors ............................................................................................................. 67 4.2.1.6.

Payload Relevance and Science Merit .................................................................... 68 4.2.2. Experiment Requirements and Objectives ..................................................................... 69 4.3.

Success Criteria ....................................................................................................... 69 4.3.1. Requirements .......................................................................................................... 69 4.3.2.

LSIM and RGEFP .......................................................................................................... 70 4.4. RGEFP Motivation ................................................................................................. 70 4.4.1. RGEFP aircraft testing plan .................................................................................... 71 4.4.2.

RGEFP Requirements and Timeline ............................................................... 71 4.4.2.1. RGEFP Hardware Development ..................................................................... 72 4.4.2.2.

RGEFP Setup and Solution ..................................................................................... 73 4.4.3. Test Plan.................................................................................................................. 74 4.4.4.

Flight Experiment Integration ........................................................................................ 75 4.5.5. Avionics ................................................................................................................................. 79




Avionics Overview ......................................................................................................... 79 5.1. Avionics Success Criteria ............................................................................................... 80 5.2. SIDES Design Approach ................................................................................................ 81 5.3.

SIDESboard ............................................................................................................ 82 5.3.1. SIDES Electrical Harness ....................................................................................... 83 5.3.2. Master IMU ............................................................................................................. 84 5.3.3. Master Clock ........................................................................................................... 85 5.3.4. Science Experiment Computer ............................................................................... 85 5.3.5. Telemetry ................................................................................................................ 86 5.3.6. Sensors and Gauges ................................................................................................ 86 5.3.7.

Pressure Gauge ................................................................................................ 86 5.3.7.1. Temperature Gauge ......................................................................................... 86 5.3.7.2. Strain Gauge .................................................................................................... 87 5.3.7.3. On-board Camera for observing the Launch Vehicle ...................................... 87 5.3.7.4.

Ground Station ............................................................................................................... 87 5.4. Purpose .................................................................................................................... 88 5.4.1. Function .................................................................................................................. 88 5.4.2. Design Considerations ............................................................................................ 89 5.4.3.

Choice of Antenna ........................................................................................... 89 5.4.3.1. Choice of Camera ............................................................................................ 90 5.4.3.2. Motor Sizing .................................................................................................... 90 5.4.3.3.

6. Safety ..................................................................................................................................... 92 Overview ........................................................................................................................ 92 6.1. Launch Vehicle Safety ................................................................................................... 92 6.2. Payload Safety ................................................................................................................ 94 6.3. Environmental Concerns ................................................................................................ 99 6.4.

7. Project Plan .......................................................................................................................... 102 Project Schedule ........................................................................................................... 102 7.1. Schedule Risk ............................................................................................................... 103 7.2.

High Risk Items .................................................................................................... 103 7.2.1. Low-to-Moderate Risk Tasks ............................................................................... 104 7.2.2.

8. Project Budget ..................................................................................................................... 105 Funding Plan ................................................................................................................ 105 8.1. Current Sponsors .......................................................................................................... 105 8.2. Projected Project Costs ................................................................................................. 106 8.3. Actual Project Costs ..................................................................................................... 107 8.4.

PDR Budget Summary .......................................................................................... 107 8.4.1. Flight Vehicle Costs .............................................................................................. 108 8.4.2.




Actual Cost vs. Projected Cost.............................................................................. 108 8.4.3.Educational Engagement Plan and Status ................................................................................... 110

Overview ...................................................................................................................... 110 8.5. Atlanta Makers’ Faire ................................................................................................... 110 8.6. Civil Air Patrol ............................................................................................................. 110 8.7. FIRST Lego League ..................................................................................................... 111 8.8. Atlanta Middle School Outreach .................................................................................. 111 8.9.

Appendix 1: Project Timeline ..................................................................................................... 112 Appendix II: Launch Checklist ................................................................................................... 114 Appendix III: Science Overview ................................................................................................ 119 Appendix IV: Ground Test Plan ................................................................................................. 127




Table of Figures

Figure 1. 2012 – 2013 project work breakdown structure. .......................................................... 11 Figure 2. Flow down of requirements. .......................................................................................... 13 Figure 3. Project L.S.I.M. mission profile. ................................................................................... 23 Figure 4: Vespula Mk. II Overall Dimensions.............................................................................. 24 Figure 5: Internal Layout of the Launch Vehicle .......................................................................... 30 Figure 6: Electronic Altimeter Schematic ..................................................................................... 32 Figure 7: Main Parachute Sizing................................................................................................... 32 Figure 8: Staggering of the Stringers ............................................................................................ 38 Figure 9: Payload Section Structure ............................................................................................. 39 Figure 10: Stringer with Connection Hole .................................................................................... 40 Figure 11: Payload Section Rib .................................................................................................... 41 Figure 12: Connector for the Payload Section .............................................................................. 42 Figure 13: Connector for the Booster Section .............................................................................. 44 Figure 14: Testing rig.................................................................................................................. 46 Figure 15: Fin Sleeve .................................................................................................................... 48 Figure 16: Fin Sleeve Attached to Booster Section ...................................................................... 49 Figure 17: Open Rocket Aerodynamic Rocket ............................................................................. 49 Figure 18: Fin Approximation of Modeling in Simulation Software ........................................... 50 Figure 19: CP as a function of the Number of Fins ...................................................................... 50 Figure 20: 45% Scale Test Rocket and Flight .............................................................................. 51 Figure 21: Nose Cone Profile ....................................................................................................... 53 Figure 22: L789 Thrust (N) vs. Time (s) ...................................................................................... 54 Figure 23: Stability Factors during Flight ..................................................................................... 55 Figure 24: Angle of Attack vs. Time ............................................................................................ 55 Figure 25: Altitude vs. Time with an L789 Motor........................................................................ 56 Figure 26: Mass Breakdown by Section ....................................................................................... 59 Figure 27: Preliminary static testing of MR fluid mixtures in magnetic fields ........................... 65 Figure 28: Shear stress of a fluid using the two-plate test (Source: Wikipedia) ......................... 66 Figure 29: Electromagnetic absorption of water (Source: Wikipedia) ......................................... 67 Figure 30: RGEFP timeline for development .............................................................................. 72 Figure 31: Possible design of tank and pipe circuit for second phase testing of LSIM ............... 74 Figure 32: Payload Assembly ...................................................................................................... 75 Figure 33: Payload Base with 150N of loading ............................................................................ 76 Figure 34: Factor of Safety vs. Total Load from SolidWorks SimulationXpress and generated trend line equation......................................................................................................................... 78 Figure 35: SIDES system layout .................................................................................................. 82 Figure 36: SIDESboard bottom side view ................................................................................... 83 Figure 37: SIDESboard top side view ......................................................................................... 83 Figure 38: Generic star topology diagram ................................................................................... 84 Figure 39: Example of an electrical harness using zip ties and connectors ................................. 84 Figure 40: Maple board used in the Master IMU ........................................................................ 85




Figure 41: The connections of the clock to different nodes. ........................................................ 85 Figure 42: A possible camera used to analyze the payload experiment ...................................... 86 Figure 43: Xbee transceiver unit .................................................................................................. 86 Figure 44: An example of a pitot tube shown attached in flight, not the one to be used for Ramblin’ Rocketeers ..................................................................................................................... 87 Figure 45: Diagram of a helical antenna ...................................................................................... 89 Figure 46: Typical radiation pattern for a helical antenna ........................................................... 90 Figure 47: Canon Powershot SX260 ........................................................................................... 90 Figure 48. Estimated budget for the 2011-2012 USLI competition. .......................................... 106 Figure 49: Projected project cost per milestone. ......................................................................... 107 Figure 50: Actual vs. Projected Project Costs as of the PDR Milestone ................................... 109 Figure 51. Participation at the Atlanta Makers' Faire. ................................................................ 110 Figure 52. Previous FIRST Lego League outreach event. .......................................................... 111 Figure 53: Microgravity time as a function of launch angle from horizon ................................ 120 Figure 54: Slosh regimes and similarity parameters .................................................................. 121 Figure 55: Plot of B field magnitude in MR fluid versus magnitude of vector 𝝁𝟎𝑯, for iron volume concentrations of 10, 20, and 30 percent ....................................................................... 124 Figure 56: Shear stress of ideal Bingham plastic (and MR fluid model) versus shear rate 𝒅𝒗𝒅𝒏, compared to ideal Newtonian liquid ........................................................................................... 125

Table of Tables

Table 1. Mission Objectives and Mission Success Criteria for the L.S.I.M. mission .................. 14 Table 2. Launch Vehicle RVM ..................................................................................................... 14 Table 3. Flight Systems RVM ...................................................................................................... 19 Table 4. Flight Avionics RVM ..................................................................................................... 21 Table 5: Mission Success Criteria................................................................................................. 25 Table 6: Launch Vehicle Requirements ........................................................................................ 26 Table 7: Launch Vehicles Properties ............................................................................................ 33 Table 8: Recovery System Properties ........................................................................................... 33 Table 9: Recovery Characteristics ................................................................................................ 34 Table 10: Landing Kinetic Energies ............................................................................................. 34 Table 11: Black Powder Properties............................................................................................... 35 Table 12: Black Powder Masses ................................................................................................... 35 Table 13: Success Criteria ............................................................................................................ 36 Table 14: Failure Modes ............................................................................................................... 36 Table 15: Variable Definition ....................................................................................................... 37 Table 16. Payload section stringer material ................................................................................. 39 Table 17. Payload section rib material ......................................................................................... 41 Table 18: Connector material ...................................................................................................... 43 Table 19: Nose Cone Symbol Definitions .................................................................................... 52 Table 20: Altitude and Velocity of Selected Motors .................................................................... 53




Table 21: Payload Section Weight Budget ................................................................................... 57 Table 22: Booster Section Weight Budget .................................................................................... 58 Table 23: Overall Weight Budget ................................................................................................. 58 Table 24: Methods currently available for damping slosh. .......................................................... 61 Table 25: Scientific method fulfillment for LSIM ....................................................................... 63 Table 26: Accomplishments since proposal ................................................................................ 63 Table 27: List of MR fluid ingredients ........................................................................................ 64 Table 28: LSIM success criteria from the Requirements Verification Matrix ............................ 69 Table 29: LSIM Requirements ..................................................................................................... 70 Table 30: RGEFP requirements for LSIM ................................................................................... 71 Table 31: Payload Assembly Dimensions ................................................................................... 75 Table 32: Data from SolidWorks SimulationXpress, highlighting the data from assumptions... 77 Table 33: Avionics requirements ................................................................................................. 79 Table 34: Avionics Success Criteria ............................................................................................ 81 Table 35: Ground station requirements ........................................................................................ 88 Table 36. Risk Identification and Mitigation Steps ..................................................................... 92 Table 37. Launch vehicle failure modes and mitigation .............................................................. 93 Table 38. Payload hazards and mitigation ................................................................................... 95 Table 39. Payload safety failure modes ....................................................................................... 98 Table 40. Environmental Hazards, Risks, and Mitigation ........................................................... 99 Table 41. Design milestones set by the USLI Program Office. .................................................. 102 Table 42. Identification and Mitigations for High-Risk Tasks. .................................................. 103 Table 43. Low to Moderate Risk items and mitigiations. ........................................................... 104 Table 44. Summary of sponsors for the Mile High Yellow Jackets. .......................................... 105 Table 45. List of current sponsors of the Ramblin' Rocketeers. ................................................. 105 Table 46. Estimated budget distribution for the 2012-2013project. ........................................... 106 Table 47. PDR Milestone Budget Summary ............................................................................... 108 Table 48: Microgravity times for fall heights ............................................................................ 121 Table 49: Similarity parameters for simplified flight profile of the launch vehicle .................. 122




1. Introduction

School Information and NAR Section Contacts 1.1.

Team Summary Sc

hool

Info

&

Proj

ect T

itle

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 0F

1

Team

Info

rmat

ion Project Lead / Team

Official Richard

Safety Officer Tony, Joseph

Team Advisors Dr. Eric Feron Dr. Marilyn Wolf

NA

R In

form

atio

n NAR Section Primary: Southern Area Rocketry (SoAR) #571

Secondary: GA Tech Ramblin’ Launch vehicle Club #701

NAR Contacts Primary: Matthew Vildzius Secondary: Jorge Blanco

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 inviduals’ areas of expertise as well as personal interest. Figure 1 shows the work breakdown structure.

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.




Launch Vehicle Summary 1.3.

Overview 1.3.1.

The Vespula Mk II launch vehicle has a gross-lift off weight of approximately 35 pounds and currently is designed around a L789 solid motor. The structure of the launch vehicle features a rib-and-stringer design covered by a thin cellulose polymer composite skin to minimize weight. The recovery system utilizes a 36” drogue parachute slowing the launch vehicle down to 60 ft/s and a 120” main parachute to slow the launch vehicle down to 12 ft/s.

Figure 1. 2012 – 2013 project work breakdown structure.




Changes since PDR 1.3.2.

The following changes have been made since the proposal:

• Skin down-selected to be a cellulose polymer composite

• The ribs and stringers of the structure has been changed to Aluminum 6061-T6

• The fins were changed from 3 ‘traditional’ fins to 5 a modified tube-fin design to increase stability

Payload Summary 1.4.

Overview 1.4.1.

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.

Changes Since PDR 1.4.2.

• The LSIM sensor selection has targeted vibration sensors and cameras in the IR spectrum as the most viable options for detecting slosh during the flight of the launch vehicle.

• Integration will be done by a standardized simplified bracket, adaptable to any rib, rather than built into the launch vehicle structure.




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.

Figure 2. Flow down of requirements.




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) 2.4.

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






Status

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

LV-2.1

The Launch Vehicle shall use a commerically available solid motor using ammonium perchlorate composite propellant (APCP).


Use of a commerically available solid motor

In Progress

LV-2.2 The total impulse provided by the Launch Vehicle shall not exceed 5,120 N-s.


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 commerically available barometric altimeter for recording of the offical altitude

USLI Handbook 1.2 Inspection Commerically available altimeter

In Progress

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.


Proper motor selection for gross lift-off weight of the launch vehicle.

In Progress






Status

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

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 whihch is accessible from the exterior of the vechile airframe when the vehicle is in the luanch configuration on the launch pad.


Recovery system design shall incorporate one (1) independent arimig switch for each altimeter

In Progress






Status

LV-3.5 Each altimeter shall have a dedicated power supply. USLI Handbook 2.7 Inspection

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 drouge and and main parachute

In Progress

LV-3.9

Removable shear pins shall be used for both the main and drouge parachute compartments


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 restricitions.

In Progress






Status

LV-3.11

Each section of the Launch Vehicle shall have a maximum landing kinetic energy of 75 ft-lbf.

USLI Hanbook 2.2 Analysis

Properly sized main parachute to ensure landing kinetic energies below 75 ft.-lbf

In Progress

LV-3.12

The recovery system electronics shall be shielded from all onboard transmitting devices.

LV-3 Testig

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 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 speacial ground support equipment to initial launch.

USLI Handbook 1.9 Testing Use of standard igniters.

In Porgress

LV-4.2

The Launch Vehicle shall not require any external circuitry or special ground support equipment to initiate the launch other than what is provided by the range.


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 recommendations for power

In Progress






Status

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 feet long 1 in. rail (1010), or an 8 feet 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



Status

FS-1 The flight systems team shall design and build the LSIM Payload

MO-3 Inspection LSIM payload In Progress

FS-2 The LSIM payload shall be designed to fly on a SLP rocket

USLI Handbook 3.1.1 Inspection LSIM payload In Progress

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

FS-5 The Flight Systems Team shall ensure that all avionics are properly shielded from the LSIM payload.

MSC-3 Testing TBD Not Started

FS-6 The Flight Systems 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 In Progress






Status

FS-7 The Flight Systems Team shall conform to all weight, power, and dimensional requirements as per the rocket design.

MSC-3 Analysis TBD In Progress

FS-7.1 The Experiment and Avionics, with mechanical supports, shall weight no more than 10 lbf.

LV-1.1 Analysis TBD In Progress

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

FS-9 The LSIM payload shall have a dedicated power supply.

MSC-3 Inspection SIDES node In Progress

FS-10 The Flight Systems Team shall ensure redundancy and reliability of all internal electrical hardware.

MSC-3 Inspection SIDES network In Progress

FS-11 The Flight Systems 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

FS-12 The Flight Systems Team shall provide for electrical operations to begin at the beginning of the the flight trajectory.

MSC-3 Inspection TBD In Progress

FS-13 The Flight Systems Team shall ensure that the LSIM payload is shut down safely during the deployment phase of the flight trajectory.

MSC-3 Inspection TBD In Progress

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






Status

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.


Appropriate mounting to the payload interface.

In Progress

Table 4. Flight Avionics RVM



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

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






Status

FA-7 The Recovery Avionics and Recovery System shall be in a separate compartment from the Flight Avionics.

USLI Handbook 2.12.1 Inspection Integration

Elements In Progress

Mission Profile 2.5.

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.

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. The unique rib and stringer design of the vehicle will incorporate a standardized payload interface within the primary structure of the rocket. This integration will allow for higher structural efficiency, a lower structural mass fraction, and an increased payload carrying capacity. In addition, the design will feature a 5.25 inch airframe and be 8 feet, 9 inches in length. The main structure of the launch vehicle is illustrated below in Figure 3. As with any unique aerospace design, extensive ground testing will be performed to verify structural integrity and successful integration of the payload into the fully assembled launch vehicle.

Figure 4: Vespula Mk. II Overall Dimensions

The chosen launch vehicle design is unique in many aspects including its rib and stringer internal structure and one-of-a-kind fin design. The rib and stringer design provides an abundance of space for a variety of payload designs while allowing for easy access and integration. Though a kit launch vehicle would be easier to construct, the rib and stringer internal structure will have a lower mass and cost. The modified tube fin design provides a highly stable flight even in high winds. The design features five equally spaced fins that aim to reduce drag while increasing stability by combining aspects from both tube fins and straight fins. In addition, the use of a fin sleeve to attach the five fins allows their distance from the nosecone to be adjusted – thereby allowing the location of the center of pressure to be adjusted prior to each flight and accommodating changes in payloads while maintaining the desired 1.5 calibers of stability.




Vespula Mk II 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.

Mission Criteria 3.1.1.

The criteria for mission success are shown in Table 5.

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.




Table 6 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 Requirements



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

LV-2.1

The Launch Vehicle shall use a commerically available solid motor using ammonium perchlorate composite propellant (APCP).


Use of a commerically available solid motor

In Progress

LV-2.2 The total impulse provided by the Launch Vehicle shall not exceed 5,120 N-s.


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 commerically available barometric altimeter for recording of the offical altitude

USLI Handbook 1.2 Inspection Commerically available altimeter

In Progress






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.


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

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 whihch is accessible from the exterior of the vechile airframe when the vehicle is in the luanch configuration on the launch pad.


Recovery system design shall incorporate one (1) independent arimig switch for each altimeter

In Progress






Status

LV-3.5 Each altimeter shall have a dedicated power supply. USLI Handbook 2.7 Inspection

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

LV-3.8 The Launch Vehicle shall utilize a dual deployment recovery system.


Utilization of a drouge and and main parachute

In Progress

LV-3.9

Removable shear pins shall be used for both the main and drouge parachute compartments


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 restricitions.

In Progress






Status

LV-3.11

Each section of the Launch Vehicle shall have a maximum landing kinetic energy of 75 ft-lbf.

USLI Hanbook 2.2 Analysis

Properly sized main parachute to ensure landing kinetic energies below 75 ft.-lbf

In Progress

LV-3.12

The recovery system electronics shall be shielded from all onboard transmitting devices.

LV-3 Testig

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 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 speacial ground support equipment to initial launch.

USLI Handbook 1.9 Testing Use of standard igniters.

In Porgress

LV-4.2

The Launch Vehicle shall not require any external circuitry or special ground support equipment to initiate the launch other than what is provided by the range.


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 recommendations for power

In Progress






Status

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 feet long 1 in. rail (1010), or an 8 feet 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 in Figure 4. For the purpose of simulation, the launch vehicle has been modeled using the Open Rocket Software, with both parachutes made of rip-stop nylon. As a linearized assumption in the limitations of the software, the five modified tube fins are modeled as ten flat fins for drag profiles during descent.

Figure 5: Internal Layout of the Launch Vehicle




During descent, 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. 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.

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 22 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 parachute casings will be made of G10 fiberglass, and the bulkhead under the main chute will be made of aluminum. Two-inch stainless steel U-Bolts will be drilled into the bulkheads, and will be used to attach the shock cords. Five Nylon 2-56 screws will be used as shear pins to keep the rocket 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.

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 system setup for each altimeter is shown below in Figure 5.




Figure 6: Electronic Altimeter Schematic

Parachute Dimensions 3.3.1.

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 aforementioned 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 12.0 ft/s. The main parachute can have a diameter of 10 to 13ft depending on the weight of the payload section. Figure 6 below shows the main parachute sizing.

Figure 7: Main Parachute Sizing




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 35 lb

Sea level density 0.00237 slugs/ft3 CDof Launch vehicle 0.75 Max Kinetic Energy 75.0 ft-lbf

Table 8: Recovery System Properties

Properties Main Parachute Drogue Parachute

Dimensions 10 ft diameter 3.0 ft diameter Surface Area 136 ft2 7.07 ft2 Estimated CD 1.4 1.2

Target Descent Rate 12.0 ft/s 60 ft/s

Drift Profile Analysis 3.3.2.

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 for drift, time in flight, and decent velocity the drift of the launch vehicle under the main and drogue parachutes can be determined.

𝐷𝑟𝑖𝑓𝑡 = 𝑇𝑖𝑚𝑒𝑖𝑛 𝑓𝑙𝑖𝑔ℎ𝑡 ∗ 𝑉𝑤𝑖𝑛𝑑 (1)

𝑇𝑖𝑚𝑒𝑖𝑛 𝑓𝑙𝑖𝑔ℎ𝑡 =

𝐴𝑙𝑡𝑚𝑎𝑥

𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑠𝑝𝑒𝑒𝑑 (2)




𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦 = �

2𝑚𝑔𝜌𝐴𝐶𝑑

(3)

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 9.

Table 9: Recovery Characteristics

Recovery Systems Properties

Drogue Parachute Main Parachute Dimensions

(d) 3ft Dimensions

(d) 10ft

Flight Time 86.0 Sec

Flight Time 42.0 Sec

Horizontal Drift

1900 ft Horizontal Drift

925.8 ft

Kinetic Energy of Launch Vehicle 3.3.3.

The kinetic energy at landing of each independent section of the launch vehicle was calculated using the equation below. The results are summarized below in Table 10.

𝐾𝐸 =12𝑚𝑣2 (4)

Table 10: Landing Kinetic Energies

Launch Vehicle Section Weight (lb.) Velocity (ft/s) Kinetic Energy (ft-lbf) Nose Cone 1.7 12 4.476

Payload 17.6 12 39.386 Booster 10.5 12 23.497




Ejection Charges 3.3.4.

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 the equation 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 11.

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 12. The masses used will depend on the final container dimensions, which were estimated at 5.25 inches in radius and 22 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 11: Black Powder Properties

Constant Value

Combustion Gas Constant 22.16 ft lbf/ lbm °R

Combustion Temperature 3307 °R

Table 12: Black Powder Masses

Main Drogue

Total Pressurization 10 psig 9 psig

Ejection force 446lbf 393lbf Black Powder 5 grams 2 grams




Testing 3.3.5.

In order to ensure the safety and viability of the calculations made in determining the black powder masses, ground testing will be done before flying the launch vehicle recovery system. The rocket will be placed horizontally on the ground, on a relatively smooth surface to minimize unwanted static friction irrelevant to a flight environment. Padding will surround the test area to protect participants and the rocket from debris. Table 13 and Table 14 illustrate the conditions for test success and failure.

Table 13: Success Criteria

Success Criteria

Ejection charge ignites Shear pins break

Launch vehicle moves half the distance of shock cord

Table 14: 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 structural subsystem of the launch vehicle must provide a strong, reliable frame for carrying the payload throughout the duration of the mission, allow for the easy integration of aerodynamic components, and have minimal weight. Since the first two criteria must always be met, the structure should be optimized to have a minimal weight while still satisfying the first two constraints. To achieve this minimal weight, a rib and stringer architecture was chosen over the traditional thick wall architecture.

For the sake of clarity in the following discussion, the notation and variables used in the course of this section are summarized below in Table 15.




Table 15: Variable Definition

Variable Definition

t Minimum thickness (in inches) for a stringer wall or a rib to prevent yield, buckling, or general failure.

d Outer diameter (in inches) of each stringer. s A coefficient (dimensionless) related to the number of stringers in each section. n Factor of safety (dimensionless).

Pcr Critical loading (in pounds-force). For stringers, the loading is in compression. For ribs, the loading is in shear. The critical loading is determined by the maximum expected force in flight.

L The length (in inches) of each stringer. E Young's modulus (in psi).

ro-ri The difference in the outer radius and inner radius (in inches) of each rib. τultimate The ultimate shear stress (in psi) of the rib material.

In the rib and stringer architecture, long, thin stringers run parallel to the longitudinal axis of the launch vehicle and bear flight loads in compression. To minimize the weight of each stringer, cylindrical tube geometry was chosen. The specific dimensions required for structurally competent strings are given by the following equations:

𝑡 =

12𝑑 −

�𝑑4 − 𝑠𝑛𝑃𝑐𝑟𝐿24

2𝜋3𝐸

(6)

𝑠 =16

(𝑁𝑢𝑚𝑏𝑒𝑟 𝑜𝑓 𝑠𝑡𝑟𝑖𝑛𝑔𝑒𝑟𝑠 𝑝𝑒𝑟 𝑠𝑒𝑔𝑚𝑒𝑛𝑡)

(7)

Equation (6) gives the minimum wall thickness, t, of a given stringer assuming a hollow cylindrical stringer design. In Equation 1, the factor of safety, n, is taken to be 2.5, and the outer diameter of the stringers, d, the critical loading of the stringers, Pcr, the stringer length, L, the stringer coefficient, s, given by Equation (7), and the modulus of elasticity, E, are design decisions based on the requirements for each structural section.

Another component of the rib and stringer architecture is the thin ribs running along the central axis of the launch vehicle. To avoid large column lengths in the stringers and thicker stringers,




the stringers are staggered such that a stringer on one surface of the rib is not coaxial with a stringer on the opposite side of the rib, as seen below in Figure 7.

Figure 8: Staggering of the Stringers

Because of this staggering, the ribs experience very few forces in compression; the primary force on the rib is in shearing. The minimum thickness of each rib is governed by its strength in shearing. Equation (8) gives the minimum thickness of the rib as a function of the critical loading, Pcr, the factor of safety, n, the radial thickness, ro-ri, and the shear strength, τultimate.

𝑡𝑚𝑖𝑛 =𝑃𝑐𝑟𝑛

2(𝑟𝑜 − 𝑟𝑖)𝜏𝑢𝑙𝑡𝑖𝑚𝑎𝑡𝑒 (8)

Payload Section 3.4.1.

In order to have sufficient space to contain the scientific payload and to allow for easy access to said payload, the payload section will be comprised of three segments of four stringers as illustrated below in Figure 8. Each segment will be 12 inches in length. The outer diameter of




the stringers is set at 0.375 inches. These design parameters yield a critical loading of the stringers of 87.5 lbf.

Figure 9: Payload Section Structure

Initially, carbon fiber was being considered the most likely material to be used for the stringers due, to its high strength to weight ratio. However, more materials were also considered and compared with the carbon fiber, as listed in Table 16.

Table 16. Payload section stringer material

Material Minimum thickness (in) Actual thickness (in) Stringer mass (lbm) Aluminum 6061 0.00398 0.049 0.0592

Grade 4130 Alloy Steel 0.00134 0.028 0.104 G-11/Garolite 0.375 0.375 0.0788 Carbon Fiber 0.375 0.375 0.0775

Copper Alloys 0.0023 0.032 0.134

Some concerns with carbon fiber include delamination and tolerance issues. As the stringers will have holes on each end to provide connection points with the ribs, as seen below Figure 9, there is a possibility of delamination. Delamination could be a structural liability, and could result in failure of structural integrity. In addition, the straightness tolerance of carbon fiber from the material supplier was not rated. Given the internal requirement that the launch vehicle should be within 1 degree of vertical, a low straightness tolerance for the stringers is essential.




Figure 10: Stringer with Connection Hole

Delamination could cause a structural liability and possible failure of structural integrity. In addition, the straightness tolerance of carbon fiber from the material supplier was not rated. Given the internal requirement that the launch vehicle should be within 1 degree of vertical, a low straightness tolerance for the stringers is essential.

After researching and comparing alternative materials, Aluminum 6061 was chosen as the stringer material in lieu of carbon fiber. Aluminum 6061 has a straightness tolerance of ±0.010 inch per foot and an outer diameter and inner diameter tolerance of ±0.004 inches for a tube. Since aluminum is not a composite material delamination around the holes in the stringers is not a concern. In addition, aluminum’s reasonable strength to weight ratio, price, and availability from the supplier made aluminum the ideal stringer material.

In designing the ribs for the payload section, the shear loading was taken to be 87.5 lbf to match the loading experienced by each stringer. That is, given the predicted mass of the launch vehicle at lift-off and the maximum acceleration due to thrust, each stringer in the payload section will transmit 87.5 lbf to the surface of the rib. To reduce mass, the center section of the rib was removed, as most of the forces collect on the outer surface of the rib, rendering the core a non-load-bearing element. Furthermore, eight points along the outer diameter were notched, as seen in Figure 10, to allow the avionics group easy accessfor wiring the interior of the launch vehicle.




Figure 11: Payload Section Rib

This design feature eliminates the tangling of wires on the interior of the launch vehicle and allowes for easy removal of avionics components. Several materials were examined in determining the optimal material and thickness for the ribs. The results of this study are found below in Table 17.

Table 17. Payload section rib material

Material Minimum thickness (in) Actual thickness (in) Rib mass (lbm) Aluminum 6061 0.0056 0.016 0.02

Carbon steel 0.0037 0.012 0.04 Copper 0.0040 0.016 0.06

G-10 / Garolite* 0.0089 0.031 0.03 Plywood* 0.0015 0.125 0.02

Stainless steel 0.0009 0.015 0.05 Titanium 6Al-4V 0.0021 0.016 0.03




Despite the light weights of composite materials such as G-10 fiberglass (garolite) and plywood, composites tend to perform poorly in shear along layer boundaries and tend to have poor bearing strengths. The failure modes of composite materials include delamination, which is difficult to detect but can have devastating consequences, and shattering, which compromises the rest of the launch vehicle. Metals, on the other hand, experience a plastic deformation before ultimately failing. While this deformation ultimately destroys the structural integrity of the rib, the deformation does not cause the immediate failure of the rest of the launch vehicle structure. . Due to these failure modes, Aluminum 6061, a ductile metal which was also chosen for the stringers, has been chosen for the rib material. Initial studies in the SolidWorks simulation environment indicate that the chosen thickness is not sufficient to prevent plastic deformation, so the thickness of the rib was increased to 0.08 inches. In the interest of verifying that the plastic deformation was not the result of a simple mathematical error, the same simulations were run on the rib component, but using a variety of the materials previously considered. These trials confirmed that an increase in material thickness would be required.

Due to the thin dimensions of the ribs, stringers cannot be bolted directly to the ribs without some sort of adapter. To overcome the problem of attaching the ribs and the stringer, a connector was designed to securely grasp the stringer and the rib such that there are no degrees of freedom between the two primary structural components. The connector design is illustrated below in Figure 11.

Figure 12: Connector for the Payload Section




After designing the connector a material study was conducted and the results are listed below in Table 18.

Table 18: Connector material

Material Manufacturing method Component mass (lbm)

ABS plastic 3D printing 0.02

Aluminum 6061 CNC milling 0.06

Delrin plastic Injection molding 0.03

Polyurethane resin Casting 0.05

Titanium 6Al-4V CNC milling 0.10

While ABS plastic allowed for the lightest material, the extra strength afforded by molding the connector out of Delrin plastic was the optimal solution. Metals were avoided due to the difficulty of machining the complex shape, particularly the deep, narrow channels found in the connector.

Booster Section 3.4.2.

The structure for the booster section of the launch vehicle is based on the structure of the payload section. Similar to the payload section structure, the structure of the booster section will be a skeleton composed of long, thin stringers connected to thin ribs.

Due to concerns of integrating aerodynamic subsystems with the booster section structure, the booster section structure has a specific set of requirements to adapt to aerodynamic design. Because of the launch vehicle’s unique five-fin design and the fact that the positioning of the fin sleeve can be changed with respect to the longitudinal axis of the launch vehicle, the connection points in the structure must match those of the fin sleeve. Since the fin sleeve connection points are directly above and below each fin, connection points in the fin sleeve are 72° apart. This geometry requires that the connection points in the booster section structure are also 72° apart. The connection separation angle requires each subsection of the booster skeleton to contain five equally-spaced stringers as opposed to the four required by the payload section skeleton, hence the coefficient s from Equation (7) is 3.2 for the booster section. The coefficient s relates the distribution of the critical loading to the number of stringers in a structural segment.




Since each booster section stringer has a length of 9.5 inches, the minimum wall thickness for the booster section is smaller than for the payload section. However, this difference in wall thickness between the payload and booster sections is not significant enough to obtain new material stock specifically for the booster stringers. For this reason, the booster section stringers will use the same material, and have the same radial dimensions, as the payload section stringers, with an outer diameter of 0.375 inches and an inner diameter of 0.277 inches.

The radial dimensions of the booster section are not significantly different from those of the payload section, and because booster section components will experience the same loading as payload section components, ribs and connectors will maintain nearly the same dimensions and mass properties as their payload section counterparts; thus, it is not necessary to reexamine the material selections for each component.

The largest deviation from the payload section components in the booster section is the rib-to-stringer connector. Since each rib must accommodate ten stringers instead of eight in the payload section, the maximum angular width between each stringer is 36°. Additionally, since the fin sleeve must bolt onto the connectors, the flange on each connector must have an appropriate adaptor for the fin sleeve. Similar to the connectors for the payload section, the connectors for the booster section will be made from injection molded Delrin plastic. The booster section connector design is illustrated below in Figure 12.

Figure 13: Connector for the Booster Section




Mass Breakdown 3.4.3.

Section Part Mass(lbs.) Number Payload Rib .08 4 Payload Stringer .06 12 Payload Connector .03 24 Payload Bulkhead .15 2 Booster Rib .08 4 Booster Stinger .05 15 Booster Connector .03 30

Total Mass(lbs.) 4.03

Section Integration 3.4.4.

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 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.

Manufacturing 3.4.5.

There are three main parts that make-up the internal structure of the launch vehicle; ribs, stringers, and connectors. Each of the ribs and stringers will be made of aluminum and the connector will be made of Delrin plastic.

The ribs have the simplest geometry of all the components, as a shape cut from a flat plate. Because of this simple geometry, the ribs can be manufactured as a two-dimensional object. Using the two-dimensional drawing of the rib, a part can be cut out from a flat plate of Aluminum 6061 using a water jet or a laser cutter. Using either of these two machines, which both cut a material using a computer-controlled cutting tool, precision parts can be mass-produced with minimum wasted material and maximum accuracy and precision.

The stringers will be cut from a rod of aluminum and each stringer will be 10-12 inches long. In order to securely fit each stringer to a connector, a thru hole will be drilled ¼ of an inch from the top and bottom ends of each stringer. This hole is in a plane perpendicular to a plane flush with the top or bottom flat surface and both holes are in the same plane. The purpose of the holes is to insert a #2-56 McMaster screw to hold the stringer securely in the connecter.




The connector is a piece designed solely for the purpose of mechanically attaching the ribs and stringers. The connectors are more complex, but necessary for the stability of the internal structure. Each connector will be manufactured using injection molding. The bottom of the connector contains a space the same thickness as the rib for the rib to slide into. The top of the connector contains a hole for the stringer to fit into and a hole for a screw that will hold the rib in place. There is also a hole for a #2-56 McMaster screw to fit into for the purpose of holding the stringer in place.

.

Future Testing 3.4.6.

Ground testing will be performed to ensure structural integrity while loads up to a factor of safety of 2.5 are applied. The test rig used is designed to perform dynamic and static loading tests and is illustrated below in Figure 13. The testing rig features a rail-mounted impact machine capable of holding different mass that can be lifted to a max height of five feet. The ability to hold different mass at specific heights simulates different impact energies.

Figure 14: Testing rig

For each test a known mass will be dropped at a certain height that correlates to a specific design impulse. The impulse value will estimated from the maximum acceleration of the launch vehicle in the program simulation Open Rocket. The height for the mass in Equation (11) will be found




using the Conservation of Energy and Conservation of Momentum principles shown in equation (9) and equation (10), where

mt is the drop mass.

E = mtgh +12

mtV2 (9)

Momentum = mtV2 − mtV1 (10)

h =1

2gI

M

2

(11)

In addition to the dynamic loading testing static loading testing will be conducted on the ribs and thrust plates to determine their maximum allowable loading and yield stresses. The static loading testing will be conducted utilizing an Instron loading machine and Vic-3D stress/strain analysis image capturing software.

In addition, future testing will possibly include a vibration test to check for resonance frequencies. This test is important in that if the vibrations oscillate with the components natural frequency structural failure will occur.

Launch Vehicle Performance Analysis 3.5.

Fin Design 3.5.1.

Performance characteristics as related to the aerodynamics of the launch vehicle primarily involve the management of stability factors throughout the ascent phase. Knowledge of the center of gravity (CG) location with respect to the location of the center of pressure (CP) is critical in designing for static stability. Because the location of the CG and CP change throughout flight, evaluation of these changes was completed through software simulation and analysis. Vespula Mk II utilizes a new stabilizing fin concept, as shown in Figure 14.




Figure 15: Fin Sleeve

The new fin concept is a modified tube fin design intended to provide more stability with less drag than a traditional tube fin. Additionally, the fins will be mounted to an adjustable sleeve for the purpose of altering fin position on the launch vehicle with respect to the nosecone. The fin sleeve design allows for the CP location to be altered before flight.




Figure 16: Fin Sleeve Attached to Booster Section

For such a design, flight simulation tools can not accurately predict performance with regard to stability factors. Evaluation of the CP location for the fin configuration was completed in Open Rocket. The Open Rocket aerodynamic model used for these approximations is illustrated below in Figure 16.

Figure 17: Open Rocket Aerodynamic Rocket

Approximations were completed to account for the unique shape of the fins through modeling each fin as having a span equal the surface length from the fin root chord to a tip chord lying at the mid-point of the semi-circle, as shown in Figure 17.




Figure 18: Fin Approximation of Modeling in Simulation Software

In addition, each Vespula Mk II fin was modeled as two fins in the Open Rocket software program. Since the maximum number of fins available for modeling purposes is eight in Open Rocket, the CP location was plotted as a function of number of fins, as displayed in Figure 18.

Figure 19: CP as a function of the Number of Fins




A trend-line was formed from the data plot in order to develop equation X, which provides the CP location, in inches, with respect to the nose cone tip with x representing the number of fins.

𝐶𝑃 𝑙𝑜𝑐𝑎𝑡𝑖𝑜𝑛 = . 062𝑥3 − 1.2933𝑥2 + 9.4455𝑥 + 54.252 (12) From equation (12), the CP location is calculated to be 81.4 inches from the nose cone tip. The CP and CG locations are separated by a designed distance to allow for 1.75 calibers of static stability at lift-off as illustrated previously in Figure 16.

Verification of the evaluated results from flight simulation was completed through the development of a 45% scale test vehicle. The test flight served to verify the driving concept behind the experimental fins and visually inspect the vehicle’s response to flight conditions in order to validate CP location approximations. Figure 19 displays the 45% scale launch vehicle on the launch rail and during ascent.

Figure 20: 45% Scale Test Rocket and Flight




In addition to verifying simulation software analysis and the fin design concept, flight data from the test will be used to develop an aerodynamic drag model and altitude approximation method for the full scale model based on performance variables.

Nose Cone 3.5.2.

The nose cone style selected in termed a Von Karman nose cone. Von Karman nose cones are designed for a theoretical minimum drag, and described mathematically by the following equations:

𝜃 = cos−1 �1 −2𝑥𝐿 �

(13)

𝑦 =

𝐷/2√𝜋

�𝜃 − sin 2𝜃

2 (14)

The variables are defined below in Table 19.

Table 19: Nose Cone Symbol Definitions

Symbol Definition 𝜃 Surface Turning Angle

𝑥 Incremental Length from Nose Cone Tip

𝐿 Overall Nose Cone Length

𝑦 Incremental Distance from Nose Cone Centerline 𝐷 Maximum Nose Cone Diameter

These equations yield a nose cone 25 inches in length with an outer diameter of 5 inches. The shape of the nose cone is illustrated below in Figure 20.




Figure 21: Nose Cone Profile

Motor Selection 3.5.3.

Table 20 below provides a range of motors and simulated altitudes and maximum velocities.

Table 20: Altitude and Velocity of Selected Motors

Motor Simulated Altitude (ft) Vmax (MPH) L1150 5823 462 L111 5786 460 L935 5293 424 L789 5280 412 L425 5118 329

Based on these calculations, the current motor chosen for Vespula Mk II is the L789 motor. The plot of thrust versus time for the selected L789 motor is shown below in Figure 21. As the thrust from the motor cuts off, the rocket’s rate of climb slowly declines. The bulk of the launch vehicles’s altitude gain takes place after the motor burnout.




Figure 22: L789 Thrust (N) vs. Time (s)

CP and CG 3.5.4.

Windy flight conditions can cause the launch vehicle’s angle of attack, and hence CP location, to change during flight. An increase in angle of attack causes the CP location on the rocket to move forward. Simultaneously, in response to the change in angle of attack, the launch vehicle changes pitch in attempt to return to a zero angle of attack, returning the CP to its original location. The pitch and pitch rate of the launch vehicle in response to angle of attack perturbations are determined by the distance between the CG and CP locations. For Vespula Mk II, a caliber of stability equal to 1.75 is desired for launch. Figure 22 displays the anticipated CP and CG changes during flight in response to a 13.5 ft/s horizontal wind and mass changes based on an L789 solid rocket motor and gross lift-off weight of 32.4 lbs. The weather conditions for this simulation were taken from last year’s competition.




Figure 23: Stability Factors during Flight

Figure 23, below, displays the expected angle of attack changes for Vespula Mk II in response to a 13.5 ft/s horizontal wind during launch. As with Figure 22 the wind speed was selected due to flight conditions found during previous competition flights. As Figure 22 and Figure 23 suggest, approaching one second, the angle of attack is approaching zero while the CP is reaching its equilibrium point. At three seconds into flight, the launch vehicle reaches a desired equilibrium point of 3 calibers of stability for the remainder of the ascent phase.

Figure 24: Angle of Attack vs. Time




Altitude Predictions 3.5.5.

Vespula Mk II will have a dry mass of 25 lbs. prior to the installation of the rocket motor. Currently, the selected rocket motor is rated as L789. Based on this motor selection, the gross lift-off weight is 32.4 lbs. Flight weather conditions based on previous competitions were used as inputs for flight simulations completed in Open Rocket. Wind at launch was approximated at 13.5 ft/s with a standard atmospheric model using the elevation above sea level at Toney, Alabama. Flight simulation approximates an apogee at 5,280 feet above ground level, as shown in Figure 24. When comparing Figure 23 to Figure 24, it can be seen that the motor burnout corresponds to an inflection point in the altitude curve.

Figure 25: Altitude vs. Time with an L789 Motor

Fabrication and Materials 3.5.6.

The modified tube fins will be composed of a carbon fiber/Kevlar hybrid. These fins will be fabricated using a vacuum bagging technique over a mold made of an appropriately-sized steel rectangular block and a steel half-cylinder welded together. The adjustable sleeve on which the fins will be mounted will be a hollow carbon fiber tube fabricated in a similar manner. The sleeve will have holes drilled into it on either end, which will be used to fasten the sleeve to the




rocket body. The fastening points on the rocket body will be placed one half inch apart, allowing the CP to be adjusted by moving the fin sleeve over a range of four inches.

Future Testing and Analysis 3.5.7.

Following is a list of future plans for aerodynamic testing of the rocket:

1. Perform wind tunnel testing to obtain experimental CD for comparison with test flight.

2. Perform CFD analysis on rocket to obtain CD, CP location, and pressure distribution for

comparison with experimental values.

3. Create altitude vs. time model using wind tunnel results.

4. Determine boundary layer profiles and pressure distributions inside of tube fins.

5. Use strain gauges to determine flutter and vibrations on tube fins.

6. Analytical comparison between modified tube fin concept and traditional fin design.

Vespula Mk II Mass Breakdown 3.6.

Mass breakdown for the booster and payload sections are summarized in Table 21 and Table 22 with systems level summary shown in Table 23. The values obtained for the booster section were mostly estimated utilizing Solidworks. However, since material properties were entered manually, the estimated weight should be fairly accurate. The values for nose cone, drogue chute, main chute, shock cords, iMPS structures, and the motor case are actual weights obtained from a scale. Note that the ribs make up the most mass for the iMPS structure. This component is overbuilt structurally in order to have a satisfactory fastener edge clearance. Additionally, the mass breakdown is also presented in terms of mass fractions, as illustrated in Figure 25.

Table 21: Payload Section Weight Budget

Payload Section Weight (lb.) Quantity Total Weight (lb.) Rib 0.13 4 0.52

Stringer 0.06 12 0.72




Payload Section Weight (lb.) Quantity Total Weight (lb.) Connector 0.03 24 0.72 Bulkhead 0.23 2 0.46

Skin 0.1 1 0.1 Epoxy/Paint 0.13 1 0.13

Total 2.65

Table 22: Booster Section Weight Budget

Booster Section Weight (lb.) Quantity Total Weight (lb.) Rib 0.13 4 0.52

Stinger 0.05 15 0.75 Connector 0.03 30 0.9

Epoxy/Paint 0.3 1 0.3 Fin Sleeve 1.75 1 1.75

Total 4.22

Table 23: Overall Weight Budget

Component Weight (lbs)

Nose Cone 1.7 Drogue Chute + Shock Cords 1.43 Main Chute + Shock Cords 2.76

Avionics System 5 Payload 10

Payload & Recovery Structure 2.65 Booster Structure 4.22 AeroTech L1390 7.45

Total 35.21




Figure 26: Mass Breakdown by Section

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.

Interface with the Ground 3.7.1.

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.

Interface with the Ground Launch System 3.7.2.

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 willbe 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

1.7

17.6

14.5 Nosecone

Payload

Booster




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.

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.

Launch Checklist 3.8.1.

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 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, 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 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 24.

Table 24: 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.




Overview of the Experiment 4.2.

Hypothesis and Premise 4.2.1.

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.

Experimental Method and Relevance of Data 4.2.2.

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 25.




Table 25: 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.

Accomplishments Since Proposal 4.2.3.

The science section of Flight Systems has accomplished several milestones since proposal, and these accomplishments are listed in Table 26.

Table 26: Accomplishments since proposal

Accomplishments since Proposal 1 Preliminary ground testing for understanding MR fluid behavior in magnetic fields 2 Preliminary sensor selection 3 Preliminary modeling and simulation of MR fluid behavior for the purpose of sizing control

solenoids 4 Ground test plan and testing matrix 5 Requirements definition and verification matrix 6 Literature review of slosh 7 Preliminary design of ground and flight test hardware

Ground Test Plan 4.2.1.

Overview 4.2.1.1.

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. Further testing for infrared reflectance of MR fluid relative to water may be conducted so that position information may be derived from a camera installed within the flight model. Finally, a working ground model will be developed using the results from (1), (2), and (3) with constraints for flight experimentation.

MR Fluid Creation and Validation of Theory 4.2.1.2.

MR fluid can be created from three ingredients: carrier oil, magnetic particles, and surfactant. Table 27 provides example MR fluid ingredients in the design space.

Table 27: 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 27: 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.

MR Fluid Shear Stress Characterization: Two Plate Test 4.2.1.3.

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 plates used must not be strongly magnetic; thus, the two current plate choices are wood or aluminum.

Figure 28: 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 Appendix IV.

Visible detection of motion: IR reflectance 4.2.1.4.

The team will utilize IR reflectance because it may show the slosh of the propellant simulant relative to MR fluid when using an appropriate recording device. Makeshift testing using an 850nm IR LED emitter and a cell phone CMOS camera showed clear waves in the propellant simulant as it sloshed. Figure 3 shows that water’s absorption rate at around 850nm is relatively low; thus, there will be high reflectance when IR light is projected onto the propellant simulant.




Figure 29: Electromagnetic absorption of water (Source: Wikipedia)

Even in the case that the MR fluid is entirely opaque to IR light, differential lighting of the water should present much information of the slosh pattern.

Certain CMOS cameras have the ability to detect IR light without the need of a specifically IR rated camera. The team will need to do more research on which camera to use for recording.

Working Ground Model 4.2.1.5.

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.

Sensors 4.2.1.6.

The team will review two sensors for slosh measurement: vibration sensors and IR emitters and detectors. Vibration sensors will be placed on the control and on the experiment to detect a difference in vibration between the two. This will indicate the effectiveness of baffling the




propellant simulant. A CMOS camera will be used to detect the propellant simulant’s IR reflectance further indicating the effectiveness of baffling the simulant.

Payload Relevance and Science Merit 4.2.2.

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 Appendix III




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.

Experiment Requirements and Objectives 4.3.

Success Criteria 4.3.1.

Minimum and maximum success criteria have been defined for the LSIM payload. Table 28 lists these criteria.

Table 28: 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%.

Requirements 4.3.2.

The requirements set for the LSIM experiment to satisfy both the goals of Ramblin’ Rocketeers and the USLI requirements are listed in Table 29.




Table 29: LSIM Requirements

Req. Number Requirement Description FS-1 The flight systems team shall design and build the LSIM Payload FS-2 The LSIM payload shall be designed to fly on a SLP rocket FS-4 The Flight Systems Team shall produce a working system for manipulating MR fluid

in LSIM. FS-5 The Flight Systems Team shall ensure that all avionics are properly shielded from the

LSIM payload. FS-6 The Flight Systems 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.

FS-7 The Flight Systems Team shall conform to all weight, power, and dimensional requirements as per the rocket design.

FS-7.1 The Experiment and Avionics, with mechanical supports, shall weight no more than 10 lbf.

FS-8 The flight computer shall execute all tasks necessary to the operation of the LSIM payload and avionics.

FS-9 The LSIM payload shall have a dedicated power supply. FS-10 The Flight Systems Team shall ensure redundancy and reliability of all internal

electrical hardware. FS-11 The Flight Systems 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. FS-12 The Flight Systems Team shall provide for electrical operations to begin at the

beginning of the flight trajectory. FS-13 The Flight Systems 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. 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.

LSIM and RGEFP 4.4.

RGEFP Motivation 4.4.1.

The experimental simulation of low-G environments is a notoriously difficult problem in the development of space and launch vehicle systems. Whether in the case of satellite deployables or the LSIM goal of damping slosh, a low-gravity environment analogous to the free fall




environment found on orbit is difficult to simulate and to maintain. Ground testing at 1- G is currently planned for LSIM; this testing may give insights into the general performance of the MR fluid and as well as provide more insight into various active slosh damping techniques. However, ground testing alone is insufficient; in order to provide a simulated environment and gain more understanding of the behavior of slosh and various active slosh damping techniques, the LSIM experiment will also be flown as the flight experiment aboard the Ramblin’ Rocketeers’ Student Launch Project (SLP) launch vehicle. While high launch accelerations and perturbations from aerodynamic forces make the flight experiment less-than-ideal for simulating a microgravity environment, RGEFP promises to offer at least several 30 second periods of high-precision low-G environment for the testing of microgravity experiments. For sustained durations of microgravity, the true performance of an MR fluid baffle can be measured, as well as the whole-architecture (simulated) propellant cleaning system.

RGEFP aircraft testing plan 4.4.2.

Ideally, the expanded level 2 and level 3 deliverables would be activated simply. Data logging would be started by crew members before microgravity, and accelerometer data would trigger payload operations during microgravity. The larger system for stabilizing the simulant and steady state filtering of simulant and MR fluid through a system of pipes and electromagnets would then operate for the duration of microgravity. The experiment would cease operations after the end of the preset microgravity settings and await the next microgravity period.

RGEFP Requirements and Timeline 4.4.2.1.

For convenience, the requirements for the LSIM experiment are re-stated in Table 30.

Table 30: RGEFP requirements for LSIM

Requirement Number Requirement Definition 1.1 The LSIM payload shall be designed to be scalable to fly on the reduced gravity

aircraft. 1.2 The LSIM payload shall weigh no more than three hundred (300) pounds when

scaled for the reduced gravity aircraft. 1.3 The LSIM payload shall not exceed the size of 24 in. by 60 in. by 60 in. when

scaled for the reduced gravity aircraft. 1.4 When scaled, the LSIM payload shall be designed to perform the experiment for

two flights on the reduced gravity aircraft. 1.5 Only students will participate in the organization, design and operation of the

LSIM payload.




RGEFP Hardware Development 4.4.2.2.

RGEFP hardware will be developed along with the level 1 deliverable LSIM flight payload. Control systems may need to be developed empirically as systems are built to minimize lag time due to development of theoretical models of system operation. The timeline for RGEFP development is given in Figure 29.

Figure 30: RGEFP timeline for development

If selected for the RGEFP, the second level deliverable would be a demonstration of a magnetic siphon; the third level deliverable would be to demonstrate an effective circuit of MR fluid using only electromagnets to propel the fluid through plumbing during the microgravity phase – this level in combination with the previous levels may be sufficient for investigation of a more complete microgravity system for controlling and moving a MR substance in the context of anti-slosh systems.




RGEFP Setup and Solution 4.4.3.

To take advantage of the much larger space offered for the experiment by the RGEFP program, the original payload from the USLI launch vehicle will be significantly expanded. One of the primary foreseeable problems with the MR fluid system described in previous sections would be MR fluid flowing out of the propellant tank into subsequent feed lines and other parts of a launch vehicle along with fuel, which would be inefficient and dangerous. Hence, the final RGEFP deliverable will consist of an experimental design to demonstrate a filtering setup that would solve this problem. Figure 1 below shows one possibility for the proposed experiment.

The payload for the launch vehicle comprises of a tank containing propellant simulant and MR fluid surrounded by solenoids. For RGEFP, in order to simulate propellant and possibly MR fluid exiting the tank flowing into feed lines, a pipe will be attached to one end of the tank through which both simulant and MR fluid may flow. This pipe will fork into two sections. At this point, there are two primary options for the rest of the setup.

In an actual launch vehicle, the propellant would be expelled into another chamber. Therefore, one option is to send one section of the pipe into another tank, which would carry simulant. The other pipe section’s purpose is to siphon the MR fluid out of the simulant flow and send it to a different tank. The MR fluid pipe would have magnets (most likely solenoids) attached in positions and controlled to enable the siphoning and movement of the MR fluid through the pipe, hence preventing it from going with the simulant into the other pipe (or, in an actual launch vehicle, into the next chamber). Measuring the amount of MR fluid in the simulant tank would allow the efficacy of the system to be tested. However, this setup would require that the entire experiment be reset when the solenoids are shut down after each microgravity period. All fluids in the collection tanks would have to be put back into the main tank so that the experiment can be rerun during the next microgravity period.

A second option would be rather than expelling simulant and MR fluid into separate tanks from the main tank, to send both pipes back into the main tank. This is unlike what is done with the propellant in an actual launch vehicle; however, this enables the experiment to keep running continuously, rather than draining of the simulant or MR fluid. In addition, there would be no need to reset the experiment between microgravity periods. The MR fluid pipe would act the same way in this option as in the option with separate tanks. Samples can removed and set aside after each test for analysis later regarding the effectiveness of the solid-state filter.

A third and final option would be to expel simulant into a separate tank, while sending MR fluid back into the main tank. This is perhaps most similar to how an actual launch vehicle might implement such a system, with propellant moving into the next chamber and MR fluid being re-




sent for use in the fuel tank, where it is needed for anti-slosh damping. In this case, the amount of MR fluid in the tank would be measured after each test to observe the system’s efficacy, and the simulant tank contents would be poured back into the main tank between experiment runs.

In addition to the tank and primary experimental setup, there are two possible methods of restraining it in the aircraft that are under consideration. The first is to simply secure the setup to an object or wall in the aircraft. The second is to mount the tanks and pipes inside another container, which is then secured to the aircraft. The second method would have a method of manually or automatically moving the tank inside the container, so that if desired, the tank may be perturbed by forces and accelerations other than only that of the aircraft. None of the possible setups are anticipated to be larger than 24 inches by 60 inches by 60 inches, nor are any of them expected o weigh more than 300 pounds.

In order to be fairly sure of the filter efficacy before RGEFP testing, the magnetic siphon system would first have been tested on the ground. In these tests (not conducted on the aircraft), a system consisting only of a pipe containing MR fluid will be surrounded by magnets, after which the magnets will be controlled in order to move the MR fluid through the tank. No simulant will be used during these tests. Once a working setup and control method for the solenoids is developed from this testing, the system will be combined with the main tank in order to create the RGEFP hardware described above.

Figure 31: Possible design of tank and pipe circuit for second phase testing of LSIM

Test Plan 4.4.4.

In any of the three options, crew members would start data logging before the beginning of the microgravity period. Data logging would likely include a video recording of the system, as well

Magnetic Siphon

Propellant Tank




as any sensor data described in the sections on the USLI payload for the launch vehicle. The movement of the aircraft would induce sloshing in the main tank. Experiment operations would be triggered manually or automatically during microgravity, and the entire anti-slosh system, as well as the filtering system described in the previous section, would both operate for most of the rest of the duration of microgravity. If the setup is inside the container, as described in the previous section, the crew may perturb the tank in additional predetermined ways in order to induce inertia-dominated slosh.

The effects of the MR fluid damping and filtering systems will be observed and recorded. The data logging will be stopped at the end of the microgravity period, and measurements will be taken if necessary. The final mass of MR fluid in the separate tanks, if those options are used, would be the primary anticipated possible measurement to be taken. The payload will then be reset, so any fluids that were sent to separate tanks will be placed back into the main tank, and all solenoids will be turned off. The process will then be repeated during each microgravity period.

Flight Experiment Integration 4.5.

The payload includes all experimental components. A possible configuration for the payload is shown in Figure 31. 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 31. .

Figure 32: Payload Assembly

Table 31: Payload Assembly Dimensions Parameter Value

Base Diameter 4.97” Total Height 10.58”

Payload Height 8.95”

Base Bolt

Payload Plug

Payload

Base




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.

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 32 shows the stress distribution through the base, using SolidWorks SimulationXpress Wizard.




Table 32: 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 32 shows data taken from SolidWorks SimulationXpress for the payload base. This data was interpolated to find the maximum load of the payload base. Figure 33 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. 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. To accomplish this goal, SIDES (Simultaneous Independent Data Logging & Experiment System) is being 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 33: 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.

Testing In Progress

2. Key elements of the flight systems shall

operate on independent power supplies.

Inspection, Analysis

In Progress




Requirement Number


Design Feature


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.

Analysis, Testing

In Progress

4. The flight avionics shall be capable of being

attached to the launch vehicle structure.

Inspection, Analysis

In Progress

5. GPS coordinates of the launch vehicle shall be transmitted to a ground

station.

Analysis, Testing

In Progress

6. Each avionics node shall be capable of data logging with or without

a clock pulse.

Testing In Progress

7. Each avionics node shall operate at some

equal or reduced functionality during

RS485 communication failure

Testing

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 34: Avionics Success Criteria

Requirement Numbers


Design Feature


1. The avionics system is functional throughout

the flight and if failures do occur the entire system does not go

down.

Analysis, Testing

In Progress

2. The ground station should be capable of

receiving supplementary data transmitted from the

launch vehicle.

Analysis, Testing

In Progress

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

In Progress

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 and the use of a synchronized clock signal will reduce the risk associated with centralized communication while maintaining the weight advantages of a multi-drop bus.

SIDESboard 5.3.1.

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 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 35 and Figure 36 depict the SIDESboard PCB (Printed Circuit Board) design, supporting the features listed above.

Figure 35: SIDES system layout




Figure 36: SIDESboard bottom side view

Figure 37: SIDESboard top side view

SIDES Electrical Harness 5.3.2.

The SIDES electrical harness connects all the SIDES nodes together, but improper harness design risks significant weight penalties. Architecture decisions mentioned previously have already helped to reduce the number of wires in any particular harness run, leaving distribution of the communication lines, clock synchronization lines, and any other signals as the main concern. The two main busses could either be distributed as a daisy-chain configuration to save weight, or using as star topology as illustrated in Figure 37. Considering increased risk has already been allowed by the choice of RS485 bus configuration, the trade-off for less weight is increased risk of failure. With daisy-chain topology, a broken wire in the harness affects all the nodes past the breakage, an unacceptable additional risk to the system. This leaves the star topology as the ideal configuration.

Each wire run will have up to 9 wires and use zip ties to keep the wires together. The zip ties placed approximately every 6 inches along the wire will keep the runs bundled and organized as well as anchor the runs to the structure. Figure 38 depicts a similar idea to the electrical harness for SIDES.




Figure 38: Generic star topology diagram

Figure 39: Example of an electrical harness using zip ties and connectors

Master IMU 5.3.3.

The master IMU utilizes a Maple board for increased computing power over the SIDESboard triple axis accelerometer, gyro and magnetometer IMU, and RS485 hardware. The Maple board, depicted below in Figure 39, is the communication bus master and allows for the different nodes to communicate with each other if desired. In particular, the Master IMU will facilitate sending data of interest to the Telemetry node to be forwarded to the ground station.




Figure 40: Maple board used in the Master IMU

Master Clock 5.3.4.

The Master Clock will send a synchronization pulse to all the other nodes. Each microcontroller has a clock input and return so that the clock can run to each node. The clock has a strong drive circuit to ensure that the nodes are consistently receiving clock signals. Figure 40 shows the connection layout of the Master Clock to all the nodes.

Figure 41: The connections of the clock to different nodes.

Science Experiment Computer 5.3.5.

The LSIM payload requires several LEDs (light emitting diodes), multiple vibration sensors, and a camera in order to record data, such as the camera shown in Figure 41. The LEDs will be necessary to light up the Magnetorheological fluid (MR fluid) so that the camera can show the fluid slosh. The microcontroller at this payload node will also actuate solenoids that will be used to control the MR fluid during the flight.




Figure 42: A possible camera used to analyze the payload experiment

Telemetry 5.3.6.

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 tranciever and a SIDEDboard to log the GPS data while the Xbee is transmitting the data. An example of the Xbee is depicted in Figure 42.

Figure 43: Xbee transceiver unit

Sensors and Gauges 5.3.7.

There will be three separate nodes that will be collecting various types of data from the launch vehicle flight. The types of data that will be collected are the airspeed the launch vehicle feels using a pitot tube, the temperature experienced by different parts of the launch vehicle, and the deflection of key structural elements of the launch vehicle.

Pressure Gauge 5.3.7.1.

The pitot tube, depicted in Figure 43, will be attached to the nose cone of the launch vehicle that estimates the dynamic pressure the launch vehicle is experiencing. It will be attached to its own SIDESboard that will record the pressure data.

Temperature Gauge 5.3.7.2.




The temperature node will contain temperature sensors for monitoring the behavior of materials in the booster section. Pending appropriate trade studies, these temperature sensors would provide information on differential expansion to due temperature in the booster section. The temperature experienced will be recorded on another SD card on the temperature node microcontroller.

Strain Gauge 5.3.7.3.

The strain gauge node makes use of a Wheatstone bridge, to determine the deflections the launch vehicle is experiencing at various points in the launch vehicle structure.

On-board Camera for observing the Launch Vehicle 5.3.7.4.

There will be a camera on the inside of the launch vehicle nose cone, which will be used to monitor any phenomena arising from anon-rigid aerodynamic structure. This camera will point to a mirror held in a shroud attached to the nose cone. The use of a mirror creates a smaller profile outside of the rocket, reducing drag, and increases mounting options for the camera interface within the nose cone.

Figure 44: An example of a pitot tube shown attached in flight, not the one to be used for Ramblin’ Rocketeers

Ground Station 5.4.

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.

Purpose 5.4.1.

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.

Function 5.4.2.

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 35: Ground station requirements

Requirement Design Feature Satisfying Requirement

Requirement Verification Success Criteria




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 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

Design Considerations 5.4.3.

Choice of Antenna 5.4.3.1.

Figure 45: 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 46: Typical radiation pattern for a helical antenna

Choice of Camera 5.4.3.2.

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 47: 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.

Motor Sizing 5.4.3.3.




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 (15)

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 1F

2, 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

(16)

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.

2 http://www.nar.org/NARhpsc.html

http://www.nar.org/NARhpsc.html




6. Safety

Overview 6.1.

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 36. Risk Identification and Mitigation Steps

Step Name Step Definition 1. 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. 2. 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.

3. Risk Control and Elimination After the hazards are identified and assessed a method is produced to avoid the issue.

4. Reviewing Assessments As new information becomes available the assessments will be reviewed and revised as

necessary.

The steps outlined above in Table 36 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.

Launch Vehicle Safety 6.2.

Failure modes for the launch vehicle were developed to better ensure success of the entire project. Possible modes, resultant problem, and mitigation procedures are given for each failure mode. These modes will continue to evolve and expand in scope as the project progresses. The mitigation methods will be continuously incorporated into preflight checklists. The mitigation




items detailed therein will be incorporated into the preflight checklist. Launch vehicle failure modes and mitigation are listed in Table 37.

Table 37. Launch vehicle failure modes and mitigation

Potential Failure Effects of Failure Failure Prevention

Fins Launch vehicle flight path becomes unstable

Test fin failure modes at connection to launch vehicle to ensure sufficient strength

Structural ribs buckle on take off

Launch failure, launch vehicle destroyed, possible injury from shrapnel

Wear eye wear protection, test the internal structure to ensure a factor of safety against buckling

Thrust retention plate

Motor casing falls out Test reliability of thrust retention plate

Skin zippering Internal components are exposed to flowing air currents, launch vehicle becomes unstable

Test skin adhesion reliability

Launch buttons Launch vehicle becomes fixed to launch rail, or buttons shear off

Ensure buttons slide easily in launch rail, ensure rail is of the proper size

Drogue separation Main shoot takes full brunt of launch vehicle inertia, launch vehicle becomes ballistic

Do a ground test of drogue separation as well as a flight test

Main shoot Launch vehicle becomes ballistic, severe injury, irrecoverable launch vehicle

Do a ground test of main shoot deployment, as well as a flight test.




Potential Failure Effects of Failure Failure Prevention

Land directly on fins

Fins break, and launch vehicle cannot be flow twice without fixing

Test fin failure modes at connection to launch vehicle to ensure sufficient strength

Ignition failure Launch vehicle does not launch

Follow proper procedure when setting up launch vehicle ignition system

Motor failure Motor explodes, possibly compromising launch vehicle and payload

Install motors properly according to manufacturer instructions.

Payload Safety 6.3.

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 38. Payload failure modes are outlined in




Table 39.

Table 38. 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



the eyes





Hand Saws Cuts, serious injury Only use saws under supervision of team mates. Only use tools in the appropriate manner. Wear safety

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.

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 39. 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

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.




Potential Failure Effects of Failure Failure Prevention Avionics Chips or boards are

manufactured incorrectly causing equipment failures

and misfires

Test avionics operations, and perform a flight test.

Environmental Concerns 6.4.

As already mentioned in Section 6.1, 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 40 lists possible environmental safety concerns.

Table 40. 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



the eyes

Dremel Serious injury, cuts, and scrapes



the eyes

Hand Saws Cuts, serious injury Only use saws under supervision of team mates. Only use tools in the appropriate manner. Wear safety

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 the e-matches

from excess RF radiation. Maximum output power is limited to

100 mW




7. Project Plan

Project Schedule 7.1.

The Ramblin’ Rocketeers’ project is driven by the design milestone’s set forth by the USLI Program Office. These milestones – and their dates – are listed in Table 41.Additionally, a preliminary Gantt Chart is provided in Appendix 1 It is important to note that due to the complexities of both the launch vehicle and payload designs, the Gantt chart will contain only high-level activities. In order to visualize the major tasks/steps in our design, the Team will utilize a PERT Chart/Network Diagram. This will allow for the identification of the critical path(s), alternative paths, indicate the risk associated with any particular task and will ensure a successful launch.

Table 41. Design milestones set by the USLI Program Office.

Milestone Date

Proposal 31 AUG Team Selection 27 SEP

Web Presence Established 22 OCT PDR Documentation 29 OCT

PDR Telecon 8 NOV CDR Documentation 14 JAN

CDR Telecon 23 JAN – 1 FEB FRR Documention 18 MAR

FRR Telecon MAR 25 – APR 3 Rocket Week 17 – 21 APR

PLAR Documentation 6 MAY




Schedule Risk 7.2.

High Risk Items 7.2.1.

Two (2) items have been identified as “High Risk Items.” These are: • Launch Vehicle Structure Design • Recovery System Design Table 42 lists the mitigations for these two (2) items.

Table 42. 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.




Low-to-Moderate Risk Tasks 7.2.2.

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 43.

Table 43. 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.




8. Project Budget

Funding Plan 8.1.

In order to fund the 2012-2013 competition cycle, the Ramblin’ Rocketeers have sought sponsorships from academic and industry sources. The current sponsors of the Ramblin’ Rocketeers and their predicted contributions can be found in Table 44. Additionally, the Team has also received a dedicated room in which the Team can construct and store their launch vehicle, payload, and other 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. Furthermore, the Georgia Tech Invention Studio will support all fabrication needs of the Team.

Table 44. Summary of sponsors for the Mile High Yellow Jackets.

Sponsor Contribution Date 2011-2012 Unused Funds $1,000 --

Georgia Space Grant Consortium $2,500 Sept. 2012 Georgia Tech

School of Aerospace Engineering (est.) $1,000 Nov 2012

Georgia Tech Student Government Association

(est.) $1,000 Nov. 2012

SCITOR Corp. (est.) $500 Nov. 2012 SpaceX (est.) $1,000 Dec 2012

Corporate Donations (est.) $1,500 Jan 2012 ATK Travel Stipend (est.) $400 Apr 2013 ATK Motor Stipend (est.) $200 Apr 2013

Total $9,100

Current Sponsors 8.2.

Table 45 lists the current sponsors of the Ramblin’ Rocketeers and their contributions.

Table 45. List of current sponsors of the Ramblin' Rocketeers.

Sponsor Contribution Georgia Space Financial contribution for general project expenses




Grant Consortium Financial contribution for Outreach-specific expenses Advanced Circuits Manufacturing of the SIDES boards throughout the design

process

Projected Project Costs 8.3.

The projected project budget is approximately $6,582.06 – below the projected fundraising goal by just under 40%. This cost was derived using the actual project costs from the 2011-2012 USLI competition cycle and a 15% margin was added to the Launch Vehicle and Flight Systems costs during the previous project cycle. While the Mk I and Mk II designs utilize different material for the ribs and stringer design, the previous project cycle also included start-up costs that will not be present during this project cycle. As a result, between the 15% and the inclusive start-up costs, the probability is believed to be significantly high that the projected project cost of $6,582.06 is an upper bound and should not be exceeded. The projected budget breakdown is listed numerically in Table 46 and graphically in Figure 47. The projected project cost per design milestone is illustrated in Figure 48.

Table 46. Estimated budget distribution for the 2012-2013project.

Section Cost Launch Vehicle $1,189.83

Flight Systems $617.73

Testing $1,736.96 Motor $1,000

Outreach $500.00 Operations $1,037.53

Travel $500.00 Total

Budget: $6,582.06

Figure 48. Estimated budget for the 2011-2012 USLI competition.




Actual Project Costs 8.4.

PDR Budget Summary 8.4.1.

Table 47 lists the expenses as of the PDR Milestone. The summary is broken down into four (4) main categories: Launch Vehicle, Flight Systems, Operations, and Motors. It is important to note that since separate funding for purchasing only motors is being pursued, it is has been determined to track this as a separate line item. The Launch Vehicle and Flight Systems categories are further broken down into two (2) subcategories: Flight Hardware and Testing. Operational expenses include: non-system specific test equipment, Team supplies, non-system specific fabrication supplies, as well as any travel and outreach expenses. Any system-specific equipment bought for testing is charged against that specific system, whereas generic equipment. 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.

Figure 49: Projected project cost per milestone.




Table 47. PDR Milestone Budget Summary

Subsystem Amt. Launch Vehicle $ 189.11 Flight Systems $ 0.00

Operations $185.00 Motors $ 60.00 Total: $ 434.11

Flight Vehicle Costs 8.4.2.

As of PDR, no purchases have been made for Flight Hardware for the Launch Vehicle, Flight Experiment, Flight Avionics, or the Ground Station.

Actual Cost vs. Projected Cost 8.4.3.

Figure 49 compares the actual and projected total project costs. Since testing of the Mk II structure has been delayed, the predicted project costs are projected to come in higher at CDR. However, with more time to thoroughly analyze the structure, more efficient testing batteries, and improved fabrication techniques can be developed that will minimize costs further down the project cycle. Additionally, decreased project costs will be achieved through the following:

• Internal design reviews at regular intervals • Creation and review of Manufacturing and Fabrication Orders (MFOs) prior to ordering

material • Communication, proper analysis, and continuous constructive criticism by peers all

throughout the design, fabrication, and testing processes.




Figure 50: Actual vs. Projected Project Costs as of the PDR Milestone




Educational Engagement Plan and Status

Overview 8.5.

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 8.6.

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.

Civil Air Patrol 8.7.

In the Aerospace Education program, Cadets have the opportunity to earn a Model Rocketry Badge by furthering their knowledge in the history and physics of rocketry as well as building five separate rockets ranging from non-solid fuel rockets to scale models of historic rockets. These rockets must meet specific altitude and payload requirements. The Ramblin’ Rocketeers will be working again this year with a local Atlanta-based squadron to help 20 children earn their own badge this spring.

Figure 51. Participation at the Atlanta Makers' Faire.




FIRST Lego League 8.8.

FIRST Lego League is an engineering competition designed for middle school children in which they build and compete with an autonomous MINDSTORMS robot. Every year this is a new competition centered on a theme exploring a real-world problem. The Ramblin’ Rocketeers plan to have a booth at the Georgia State FIRST Lego League Tournament and illustrate how the skills and ideas utilized in the competition translate to real world applications, such as a launch vehicle with autonomous capabilities. In addition the Ramblin’ Rocketeers plan to both help set up and judge the tournament, which is tentatively scheduled for January 25, 2013.

Atlanta Middle School Outreach 8.9.

The Ramblin’ Rocketeers also plan to go to various middle schools in the Atlanta area to make presentations and demonstrations about science and physics. The intent of these programs is to teach students in the 6th to 8th grade range about science and rockets, as well as to spark their interest in STEM fields. Specific topics that the Ramblin’ Rocketeers plan to cover in the multiple demonstrations are electricity, basic concepts of flight, and general topics relating to engineering. The Ramblin’ Rocketeers will be working in conjunction with organizations like the Society of Women Engineers to reach even more middle school students, exposing them to various topics regarding space exploration and inspiring them to pursue careers in the STEM fields. Drew Charter Middle School has been contacted, and a STEM demonstration will occur there on November 27.

Figure 52. Previous FIRST Lego League outreach event.




Appendix 1: Project Timeline

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 10/29/12




Appendix II: Launch Checklist

Pre-Launch

Performer Inspector

Packing The night before launch go through Launch Packing List

and put all items in a designated spot.

The morning of launch go through Launch 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.

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).

Install igniter in launch vehicle. 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. Arm altimeters via switches and wait for continuity check for both.

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

Modeling Slosh and MR Fluid

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 52.




Figure 53: Microgravity time as a function of launch angle from horizon

In Figure 52, the microgravity time, or ∆𝑡𝑚𝑖𝑐𝑟𝑜, was computed using equation (17).

∆𝑡𝑚𝑖𝑐𝑟𝑜 =

𝑉0sin (𝛼)𝑔

√0.05 + 1.5 (17)

In equation (17), the 1.5 s addition represents the time from apogee to chute deployment, which by representation in Figure 52 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 (18) valid to heights of 20 m with atmospheric drag. Equation (18) 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 (18)




The predicted times from equation (18) and from the flight profile are given in Figure 52.

Table 48: 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 49presents 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 54: Slosh regimes and similarity parameters

for microgravity are illustrated in Figure 53. 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 49: 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.

Fluid dynamics

Finally, a fluid dynamics analysis of the potential flow for sloshing water must be undertaken to properly simulate and interpret the results of testing. At the project’s current status, the simulation of MR fluid motion in magnetic fields is more mature than the slosh simulation itself. This delay is seen as acceptable, as due to the complex nature of liquid slosh, empirical results will be equally if not more important than simulation – and the understanding of MR fluid is key to beginning testing. However, generally the sloshing motion of a liquid in the lateral directions is given by equation (19).

∅1 = ∅2 + ∅3 + 𝐴𝑇𝑥 �

−𝜃𝑦𝑖𝜔 � + 𝐴𝑇𝑦 �

𝜃𝑥𝑖𝜔�

(19)

The potential variables ∅2and ∅3 must satisfy the free-surface condition of the liquid in the container given in equation (20).

𝜕2∅𝜕𝑡2

+ 𝐴𝜕∅𝜕𝑧

= 0 (20)

Mathematical and physical model of MR fluid in magnetic fields

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 54 below taken from a paper by Simon et al.

Figure 55: 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 = χ𝑚𝑉

µ𝐁, 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 55, compared to a Newtonian fluid.

Figure 56: 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 becomes:




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 - Calibrating Sensors

The team will be using two sensors for characterizing MR fluid and recording data: a load sensor and a CMOS camera that can detect light in the IR spectrum. The team will attach a known mass tothe load sensor and measure the reading. For the CMOS camera, the team will emit IR light onto water and note the brightness displayed by the camera. The relative brightness will indicate a fluid's IR reflectance.

Test Sequence 3- 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.

2012 – 2013 georgia tech ramblin’ rocketeers preliminary … gt usli pdr.pdf · 2012. 10....

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