llv propulsion system 100 pct design report

Lunar Lander Propulsion System: 100% Design Review

AE 445 Spacecraft Detail Design

Department of Aerospace Engineering Embry-Riddle Aeronautical University, Daytona Beach

Instructor: Eric Perrell, Ph.D.

23 April 2008

Lunar Lander Vehicle Propulsion System: 100% Design Review

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Team Members

Executive Staff

Nicholas H. Perry, Project Manager John C. Swatkowski, Research Director Mitchell Graves, Budget Manager Shailesh Kumar, System Safety Officer

Mission Assurance Division

Johann Schrell, Lead Eric McLaughlin Brendan McMahon

Configuration Management Division

Robert J. Scheid, Lead Jessica Chen Todd A. Snyder

Thermal Management Division

Jade Pomerleau, Lead Ben Klamm Chi Zhang

Manufacturing, Division

Chukwuma Akosionu, Lead Kristopher Greer Joe Tabor

Systems Integration & Testing Division

Eric J. Thompson, Lead Maggie Cordova Autumn Gee Gary Kelley Andrew Kreshek Nicholas Rehak


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Special Thanks

Team Cynthion would like to thank the following people who helped make this project a reality: to Bogdan Udrea, Ph.D., for collaborating with Team Cynthion to make the lunar lander propulsion system a reality, to Brian Ruby, for helping machine the rocket engine, to Geoffrey Kain, Ph.D., for funding the Lunar Lander project, to Richard Hedge, for all of his hard work machining the majority of the rocket engine, and finally, a very special thanks to Eric Perrell, Ph. D., for all of his hard work with the project as a mentor and instructor.


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Table of Contents

1. Introduction ................................................................................................................. 12. Main Rocket Thrusters Design Parameters ................................................................. 22.1. Performance Metrics ............................................................................................ 22.2. Engine Dimensions ............................................................................................ 122.3. Injector Design ................................................................................................... 132.4. Engine Throttling ............................................................................................... 142.5. Ignition System .................................................................................................. 152.6. Backup Ignition System ..................................................................................... 152.7. ERPL Liquid Rocket Engine Creator ................................................................. 162.8. NASA Chemical Equilibrium Analysis ............................................................. 19

3. Propulsion System Configuration ............................................................................. 203.1. CATIA ................................................................................................................ 203.2. Connections and Assembly ................................................................................ 273.3. NASTRAN Pressure Analysis ........................................................................... 293.4. Mass Budget ....................................................................................................... 39

5. Manufacturing Plan ................................................................................................... 685.1. Raw Materials and Hardware ............................................................................. 685.2. Fabrication Process ............................................................................................ 705.3. Assembly Process ............................................................................................... 825.4. Part Assembly .................................................................................................... 82

6. Systems Integration & Testing.................................................................................. 846.1. Feedline System ................................................................................................. 846.2. Electrical System for Solenoid Valves ............................................................... 856.3. Test Stand ........................................................................................................... 866.4. Data Acquisition (DAQ) .................................................................................... 876.5. Test Preparation Procedures ............................................................................... 936.6. Fuel Tank Air Evacuation .................................................................................. 956.7. Fuel Tank Fill (Propane) .................................................................................... 976.8. Fuel Tank Fill (Water)........................................................................................ 996.9. Assembly Procedure ......................................................................................... 1006.10. Test Area Overview ...................................................................................... 1036.11. Feed Line System Testing Overview ............................................................ 1146.12. Instrumentation ............................................................................................. 1206.13. Igniter Testing Procedure ............................................................................. 1226.14. Ignition Test Results ..................................................................................... 1236.15. Engine Firing Test ........................................................................................ 125

7. Project Economics .................................................................................................. 1288. System Safety Program Plan ................................................................................... 1319. Appendix ................................................................................................................. 1589.1. NASA CEA Output .......................................................................................... 158

10. References ............................................................................................................ 198


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Table of Figures, Tables, and Equations

Figure 1: Determination of Nozzle Expansion Ratio .......................................................... 2Figure 2: Nozzle Performance for Major Thrust Levels ..................................................... 3Figure 3: Specific Impulse for Varied O/F Ratio ................................................................ 4Figure 4: Characteristic Velocity for Varied O/F Ratio...................................................... 4Figure 5: Chamber Temperature for Varied O/F Ratio ...................................................... 5Figure 6: Exit Mach Number for Varied O/F Ratio ............................................................ 5Figure 7: Nozzle Exit Pressure for Varied O/F Ratio ......................................................... 6Figure 8: Coefficient of Thrust for Varied O/F Ratio ......................................................... 6Figure 9: Chamber Mass Flow Rate for Varied O/F Ratio ................................................. 7Figure 10: Exit Pressure Over Thrust Range ...................................................................... 8Figure 11: Mass Flow Rate Over Thrust Range ................................................................. 9Figure 12: Specific Impulse Over Thrust Range ................................................................ 9Figure 13: Chamber Temperature Over Thrust Range ..................................................... 10Figure 14: Assembled Engine ........................................................................................... 21Figure 15: Supersonic Nozzle ........................................................................................... 22Figure 16: Combustion Chamber ...................................................................................... 22Figure 17: Cooling Jacket ................................................................................................. 23Figure 18: Fuel Inlet Manifold .......................................................................................... 23Figure 19: Nitrous Injection Plate ..................................................................................... 24Figure 20: Nitrous Injection Dome ................................................................................... 25Figure 21: Steel Connection Ring ..................................................................................... 25Figure 22: Copper Connection Ring ................................................................................. 25Figure 23: Exploded Assembly drawing........................................................................... 27Figure 24: Section Cut: Deformation of Dome and Plates ............................................... 30Figure 25: Section Cut: Mean Pressure Distribution of Dome and Plates ........................ 31Figure 26: Section Cut: Von Mises Stress of Dome and Plates ........................................ 32Figure 27: Isometric Cut: Von Mises Stress of Plates ..................................................... 33Figure 28: Section Cut: Deformation of Jacket, Nozzle, Chamber and Manifold ............ 34Figure 29: Section Cut: Mean Pressure Distribution of Jacket, Nozzle, Chamber and Manifold ............................................................................................................................ 35Figure 30: Section Cut: Von Mises Stress of Jacket Nozzle Chamber and Manifold ...... 36Figure 31: Section Cut of Jacket and Chamber: Von Mises Stress of Assembly ............. 37Figure 32: Section Cut of Jacket and Chamber: Von Mises Stress of Jacket and Chamber........................................................................................................................................... 38Figure 33: Coordinate Axis System .................................................................................. 41Figure 34: Engine Wall Temperature from MATLAB ..................................................... 46Figure 35: Coolant Temperature from MATLAB ............................................................ 47Figure 36: Coolant Pressure from MATLAB ................................................................... 48Figure 37: Gas Wall Temperature from Excel .................................................................. 49Figure 38: Coolant Temperature from Excel .................................................................... 50Figure 39: Coolant Pressure from Excel ........................................................................... 50Figure 40: Transient Analysis Diagram ............................................................................ 51Figure 41: Time Step and Stability Data for Each Disk ................................................... 54


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Figure 42: Transient Gas/Wall Interface Temperatures for Disk 1 .................................. 55Figure 43: CFD Simulation Configuration ....................................................................... 56Figure 44: Gas Streamlines through Engine, Axial View................................................. 57Figure 45: Gas Streamlines Through Section 2 Region, Top View ................................. 58Figure 46: Conic vs. Cylindrical Surface Area ................................................................. 59Figure 47: Jacket Thickness from MATLAB ................................................................... 60Figure 48: Jacket Thickness from Excel ........................................................................... 60Figure 49: Nitrous Oxide Thermal Conductivity .............................................................. 62Figure 50: Nitrous Oxide Cp............................................................................................. 63Figure 51: Nitrous Oxide Density ..................................................................................... 63Figure 52: Time to Reach Critical Wall Temperature for Each Disk ............................... 65Figure 53: Steady State and Transient Wall Temperature Comparison ........................... 66Figure 54: C14500 Tellurium Copper Round Solid Bar ................................................... 68Figure 55: Stainless Steel Round Solid Bar ...................................................................... 69Figure 56: Brass Round Solid Bar .................................................................................... 69Figure 57: Steel Round Solid Bar ..................................................................................... 69Figure 58: End Mills ......................................................................................................... 71Figure 59: Boring Bars...................................................................................................... 71Figure 60: Rough and Finish Bar ...................................................................................... 71Figure 61: Mandrel Piece (Lower) .................................................................................... 72Figure 62: Mandrel Piece (Upper) .................................................................................... 72Figure 63: Mandrel Assembly .......................................................................................... 73Figure 64: Supersonic Nozzle (front elevation) ................................................................ 74Figure 65: Supersonic Nozzle (looking down at top) ....................................................... 74Figure 66: Nozzle and Chamber ....................................................................................... 75Figure 67: Cooling Jacket (nozzle portion) ...................................................................... 77Figure 68: Injection Dome (while being machined) ......................................................... 79Figure 69: Manifold Ring ................................................................................................. 80Figure 70: Propane Manifold Assembly ........................................................................... 80Figure 71: Test Stand ........................................................................................................ 81Figure 72: Connector Blocks ............................................................................................ 81Figure 73: Electrical System for Solenoid Valve and Indicator Light.............................. 85Figure 74: Steel Mounting Block ...................................................................................... 87Figure 75: Oxidizer Stand with Two Tanks ...................................................................... 89Figure 76: Half of the Symmetrical Load Cell Electrical Circuit ..................................... 89Figure 77: Testing Software Front Panel GUI .................................................................. 91Figure 78: LabVIEW Code. .............................................................................................. 92Figure 79: Top View of Housing Compartments ........................................................... 103Figure 80: Housing Illustration, Top View ..................................................................... 104Figure 81: Testing container as viewed in Catia ............................................................. 106Figure 82: Testing Container Illustration, Front View ................................................... 107Figure 83: Primary test site easily accessible from IC Auditorium parking lot.............. 109Figure 84: Primary test site easily accessible from Clyde Morris Blvd. ........................ 110Figure 85: Alternate test site easily accessible from ROTC parking lot......................... 112Figure 86: Alternate test site easily accessible from Richard Petty Blvd. ...................... 114Figure 87: Feed System .................................................................................................. 116


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Figure 88: Oxidizer System ............................................................................................ 117Figure 89: Pressurant System.......................................................................................... 118Figure 90: Fuel System ................................................................................................... 120Figure 91: Fault Tree Analysis ....................................................................................... 145

Table 1: Summary of Changes in Performance for Varied O/F ......................................... 7Table 2: Performance Metrics for Thrust Chamber at Nominal at O/F=7 ........................ 11Table 3: Performance Metrics for Thrust Chamber at 60% Nominal at O/F=7 ............... 11Table 4: Performance Metrics for Thrust Chamber at Nominal at O/F=3 ........................ 11Table 5: Performance Metrics for Thrust Chamber at 60% Nominal at O/F=3 ............... 12Table 6: Dimensions of the Thrust Chamber .................................................................... 12Table 7: Requirements for Injector Design at 60% Nominal 0/F=7 ................................. 13Table 8: Requirements for Injector Design at 60% Nominal 0/F=3 ................................. 13Table 9: Properties of Propellant Injectors O/F=7 ............................................................ 13Table 10: Determination of Minimum Stable Thrust at O/F=7 ........................................ 14Table 11: Determination of Minimum Stable Thrust at O/F=3 ........................................ 14Table 12: Single Engine Mass Budget (without feedline) ................................................ 39Table 13: Variable Gas Properties .................................................................................... 43Table 14: Analysis Inputs ................................................................................................. 46Table 15: Propane Thermodynamic Properties................................................................. 64Table 16: Nitrous Oxide Thermodynamic Properties ....................................................... 64Table 17: Fabricated Parts List ......................................................................................... 82Table 18: Hardware and COTS List ................................................................................. 82Table 19: Feedline System Parts and Components ........................................................... 84Table 20: Total Cost Estimate ......................................................................................... 129Table 21: Current Budget................................................................................................ 130Table 22: Propulsion System Total Cost ........................................................................ 130 Equation 1: Gas/Wall Heat Transfer Coefficient .............................................................. 40Equation 2: Coolant/Wall Heat Transfer Coefficient ....................................................... 40Equation 3: Mach-Area Relationship................................................................................ 41Equation 4: Section 1 y-position Equations ...................................................................... 42Equation 5: Section 2 y-position Equation ....................................................................... 42Equation 6: Section 3 y-position Equations ...................................................................... 42Equation 7: Section 4 y-position Equations ...................................................................... 42Equation 8: Section 5 y-position Equation ....................................................................... 42Equation 9: Gas Static Temperature ................................................................................. 42Equation 10: Gas Static Pressure ...................................................................................... 43Equation 11: Heat Transfer Between the Gas and Wall ................................................... 43Equation 12: Heat Transfer Through the Wall ................................................................. 43Equation 13: Heat Transfer Between the Wall and Coolant ............................................. 43Equation 14: Coolant Temperature Rise ........................................................................... 44Equation 15: Gas Wall Temperature................................................................................. 44Equation 16: Coolant Wall Temperature .......................................................................... 44Equation 17: Coolant Temperature at End of Disc ........................................................... 44Equation 18: Surface Area of a Body of Revolution ........................................................ 44


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Equation 19: Section 5 Surface Area ................................................................................ 45Equation 20: Coolant Pressure .......................................................................................... 45Equation 21: Differential Transient Heat Transfer Equation ............................................ 51Equation 22: Simplified Transient Heat Transfer Equation ............................................. 51Equation 23: Difference Quotient for Time Partial Derivative ........................................ 52Equation 24: Difference Quotient for Position 2nd Order Partial Derivative .................... 52Equation 25: Finite Difference Numerical Solution Through the Wall ............................ 52Equation 26: Modulus of the Finite Difference Formula.................................................. 52Equation 27: Differential Heat Transfer Boundary Equation ........................................... 52Equation 28: Difference Quotient for Heat Transfer Boundary Condition ...................... 52Equation 29: Coolant Boundary Finite Difference Numerical Solution ........................... 53Equation 30: Gas Boundary Finite Difference Numerical Solution ................................. 53Equation 31: Biot Number for Finite Difference Method ................................................ 53Equation 32: Wall Stability Criteria ................................................................................. 53Equation 33: Coolant Boundary Stability Criteria ............................................................ 53Equation 34: Gas Boundary Stability Criteria .................................................................. 53Equation 35: Roy and Thodos Estimation Technique ...................................................... 61Equation 36: Generalized Excess Conductivity Correlation ............................................ 61Equation 37: Specific Heat Equation ................................................................................ 62


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1. Introduction

Teams from AE 445 Sections 01 and 02 have collaborated together in order to design a viable entry in the Northrop Grumman X-Prize Lunar Lander Challenge. The goal of this challenge is to design and build a Lunar Lander Vehicle (henceforth referred to as LLV) which could potentially explore the Moon. The requirements of this vehicle are that it must carry a payload of at least 25 kg, rise to an altitude of at least 50 m, translate a distance of 120 m, land and perform a similar path but in the reverse direction. It is the responsibility of Team Cynthion to design the main rocket thrusters, the fuel system, and the attitude control thrusters that meet the LLV structure and control system team's requirements.

It is the intention of this report to provide a complete record of Team Cynthion's design project since the fusion of Team ALLSTAR and Team Medea. A full overview of the has been included.



2. Main Rocket Thrusters Design Parameters

2.1. Performance Metrics

The first priority was to quickly explore ideal average operating conditions for the given flight requirements where the minimum thrust is 1,633 N and nominal thrust is 4,000 N. The NASA Chemical Equilibrium with Applications (CEA) program was run, using an Oxidizer to Fuel (O/F) ratio of 7 and a chamber contraction ratio Ac/At of 8, to compare exit pressures and expansion ratios at given chamber pressures. These chamber pressures represent the range of engine thrust. It was decided from the data shown in Figure 1 to consider 60 percent nominal thrust to be the operating thrust level. This corresponds to a thrust of 3,053.2 N and a chamber pressure of 29 bar. This decision was based on the thrust level that would provide an even range of exit pressures over the flight profile.

0

0.5

1

1.5

2

2.5

3

0 1 2 3 4 5 6

Ae/At

Pc =10

Pc = 15

Pc = 20

Pc = 25

Pc = 30

Pc = 35

Pc = 38

Figure 1: Determination of Nozzle Expansion Ratio

NASA CEA was run again at this chamber pressure and the extreme thrust level chamber pressures with different expansion ratios. The resulting plotted data is shown in Figure 2. The regression equation for the average thrust was found and used to obtain the expansion ratio at which the exit pressure reached atmospheric conditions of 1 bar. The resulting expansion ratio was determined to be 4.7.



y = 0.0066x4 - 0.1435x

3 + 1.228x

2 - 5.0356x + 9.153

0

0.5

1

1.5

2

2.5

3

0 1 2 3 4 5 6

Ae/At

Tave

Tmin

Tmax

Figure 2: Nozzle Performance for Major Thrust Levels

The next priority was to explore the expected performance values for the given flight requirements where the minimum thrust is 1,635 N and nominal thrust is 4,000 N. NASA CEA was run, using O/F ratios of 3,4,5,6, and 7 and a chamber contraction ratio Ac/At of 8. This was done to obtain values needed for input into the Liquid Rocket Engine Creator (LREC) whose output was plotted versus the varying O/F ratio. The results of this are shown in Figure 3, Figure 4, Figure 5, Figure 6, Figure 7, Figure 8 andFigure 9.



200

205

210

215

220

225

230

235

240

245

250

0 1 2 3 4 5 6 7 8

O/F

Figure 3: Specific Impulse for Varied O/F Ratio

1350

1400

1450

1500

1550

1600

1650

0 1 2 3 4 5 6 7 8

O/F

Figure 4: Characteristic Velocity for Varied O/F Ratio



0

500

1000

1500

2000

2500

3000

3500

0 1 2 3 4 5 6 7 8

O/F

Figure 5: Chamber Temperature for Varied O/F Ratio

2.7

2.75

2.8

2.85

2.9

2.95

0 1 2 3 4 5 6 7 8

O/F

Figure 6: Exit Mach Number for Varied O/F Ratio



0

20000

40000

60000

80000

100000

120000

0 1 2 3 4 5 6 7 8

O/F

Figure 7: Nozzle Exit Pressure for Varied O/F Ratio

1.465

1.466

1.467

1.468

1.469

1.47

1.471

1.472

1.473

0 1 2 3 4 5 6 7 8

O/F

Figure 8: Coefficient of Thrust for Varied O/F Ratio



1.25

1.3

1.35

1.4

1.45

1.5

1.55

0 1 2 3 4 5 6 7 8

O/F

Figure 9: Chamber Mass Flow Rate for Varied O/F Ratio

It can be seen from Figure 3Figure 4, Figure 8 that the performance efficiency of the engine drops quite substantially as the O/F ratio drops from seven to three. The exit pressure, seen in Figure 7, at an O/F ratio of 3 is the same as that at the ratio of 7, so there is no drop in nozzle efficiency. Though the performance drops more than desired at the testing oxidizer to fuel ratio, the thermal effects of lowering the ratio are very desirable. Seen in Figure 5, the temperature of the combustion products drops very much at the testing O/F ratio of 3. Another desirable effect on the cooling is that the required total mass flow rate increases (Figure 9), thus allowing the propane coolant to remove heat conducted through the thrust chamber wall at a faster rate. Though the exit pressure increased slightly more towards 1 bar, this may not be a good thing. This may have the effect of creating undesirably low exit pressures when the engine is throttled down. A summary of the changes as the oxidizer to fuel ratio drops to three is shown in Table 1.

Table 1: Summary of Changes in Performance for Varied O/F

O/F Isp (sec) Cstar (m/sec) Tc (K) Me Pe (Pa) Cf M_dot (kg/s)

3 204.2288 1365.9 1772 2.718563 98587 1.466286 1.524465748

4 225.3369 1503.3 2389.76 2.903947 82626 1.469966 1.381663499

5 236.3192 1576.6 2819.35 2.871345 86783 1.469935 1.317454153

6 242.1622 1615.3 3093.78 2.834992 90922 1.470191 1.285666161

7 244.7421 1629.9 3237.06 2.79647 95900 1.472544 1.272113662

% Diff -18.0472 -17.62467454 -58.4964 -2.82526 2.763167 -0.42591 18.04719617



With the average thrust at 3,053 N and the expansion ratio at 4.7, some engine parameters were analyzed again using NASA CEA and the LREC tool. The exit pressure, mass flow rate, specific impulse, and chamber temperature were computed over the range of thrust and are shown in Figure 10, Figure 11,Figure 12, Figure 13, respectively. It can be seen here that the exit pressure and mass flow rate change proportionally to the thrust, while the chamber temperature increases proportionally to the specific impulse. Also, it can be seen how these values differ when the O/F ratio is dropped from the design value 7 to the expected testing value 3.

0

0.2

0.4

0.6

0.8

1

1.2

1.4

0 500 1000 1500 2000 2500 3000 3500 4000 4500

Thrust (N)

Figure 10: Exit Pressure Over Thrust Range



0.00000

0.20000

0.40000

0.60000

0.80000

1.00000

1.20000

1.40000

1.60000

1.80000

0 500 1000 1500 2000 2500 3000 3500 4000 4500

Thrust (N)

Figure 11: Mass Flow Rate Over Thrust Range

243.8

244

244.2

244.4

244.6

244.8

245

245.2

0 500 1000 1500 2000 2500 3000 3500 4000 4500

Thrust (N)

Figure 12: Specific Impulse Over Thrust Range



3180

3190

3200

3210

3220

3230

3240

3250

3260

3270

0 500 1000 1500 2000 2500 3000 3500 4000 4500

Thrust (N)

Figure 13: Chamber Temperature Over Thrust Range



Table 2 and Table 3 represent the performance metrics for full nominal thrust and for the operating thrust at sixty percent of nominal thrust for O/F ratios of seven and three.

Table 2: Performance Metrics for Thrust Chamber at Nominal at O/F=7 Nominal Performance Metrics

Thrust 4000 N

Chamber Pressure 38 bar

Specific Impulse 245.03 sec

Characteristic Velocity 1632.3 m/s

Characteristic Length 0.89 m

Mass Flow Rate 1.664 kg/s

Exit Pressure 1.25 bar

Exit Mach 2.799

Ae/At 4.7

Ac/At 8

Table 3: Performance Metrics for Thrust Chamber at 60% Nominal at O/F=7 Operating Performance Metrics

Thrust 3053







Exit Mach 2.796

Ae/At 4.7

Ac/At 8

Table 4: Performance Metrics for Thrust Chamber at Nominal at O/F=3 Nominal Performance Metrics

Thrust 4000 N







Exit Mach 2.799

Ae/At 4.7

Ac/At 8



Table 5: Performance Metrics for Thrust Chamber at 60% Nominal at O/F=3 Operating Performance Metrics

Thrust 3053







Exit Mach 2.719

Ae/At 4.7

Ac/At 8

2.2. Engine Dimensions

C tool and remain unchanged. These dimensions are shown in Table 6.

Table 6: Dimensions of the Thrust Chamber Thrust Chamber Dimensions (m)

Chamber ID 0.08460

Chamber OD 0.08740

Chamber Length 0.11905

Nozzle Entrance Length 0.01590

Total Chamber Length 0.13495

Nozzle Exit Length 0.06450

Nozzle Total Length 0.0804

Total Length 0.19945

Entrance Diameter 0.08460

Exit Diameter 0.06485

Throat Diameter 0.02991

Note that the nozzle dimensions are based on a 60/15 conical nozzle and thus they are different than the dimensions of the parabolic nozzle using these dimensions discussed later.



2.3. Injector Design

The propellant injector requirements have changed with the decision to test the engine at an O/F ratio of 3. The new requirements are shown in Table 7 and Table 8. Though the requirements did change, the design has been frozen. This does not pose a problem because fewer injectors would be required at the average thrust level. This means the required mass flow rate will still be able to be accommodated.

Table 7: Requirements for Injector Design at 60% Nominal 0/F=7

Injector Requirements

Propane Nitrous Oxide

Chamber Mass Flow 1.272 kg/s Chamber Mass Flow 1.272 kg/s

Pressure Drop 0.87 MPa (0.3 Pc) Pressure Drop 0.87 MPa (0.3 Pc)

Liquid Density 582 kg/m3 Liquid Density 1222.8 kg/m3

Cd 0.8 Cd 0.8

K 1.7 K 1.7

Table 8: Requirements for Injector Design at 60% Nominal 0/F=3

Injector Requirements


Chamber Mass Flow 1.524 kg/s Chamber Mass Flow 1.524 kg/s

Pressure Drop 0.87 MPa (0.3 Pc) Pressure Drop 0.87 MPa (0.3 Pc)

Liquid Density 582 kg/m3 Liquid Density 1222.8 kg/m3

Cd 0.8 Cd 0.8

K 1.7 K 1.7

The number of injectors at designed operating thrust is 132 for propane and 91 for nitrous oxide. A summary of the production injector information is shown in Table 9.

Table 9: Properties of Propellant Injectors O/F=7 Injector Properties


Material Copper Material Stainless Steel

Diameter Diameter 90o

Type Alternating Linear - Plain Type Shower Head - Countersunk

Number 147 3 Rows Number 102 5 Rows

Speed 38.4 m/s Speed 26.5 m/s



2.4. Engine Throttling

The throttled engine data was recomputed to confirm that the minimum stable thrust had not increased above the required value of 1,635 N with the variation of the O/F ratio. With the new CEA data, thrust data was computed using the LREC tool. The thrust did change, however it decrease which is good. This means that the engine may have the ability to throttle well below the requirement. The minimum stable thrusts were determined to be 1,633 N and 1,357 N, less than the desired 1,635 N. Also, the performance metrics changed with the drop in oxidizer similar to the changes at the operating thrust level. The resulting data is shown in Table 10 and Table 11.

Table 10: Determination of Minimum Stable Thrust at O/F=7

Throttle Capability

Thrust 1626.04 N

Exit Pressure 40759 Pa

Mass Flow 0.68 kg/s

Oxidizer Injector Velocity 10.82 m/s

Fuel Injector Velocity 15.68 m/s

Exit Velocity 2392.99 m/s

Exit Mach 2.734

Oxidizer Pressure Drop w/ % Pc 190000.56 5.00

Fuel Pressure Drop w/ % Pc 190000.56 5.00

% Required Minimum Thrust 100.6

Table 11: Determination of Minimum Stable Thrust at O/F=3

Throttle Capability

Thrust 1356.54 N

Exit Pressure 46313 Pa

Mass Flow 0.68 kg/s

Oxidizer Injector Velocity 10.82 m/s

Fuel Injector Velocity 15.68 m/s

Exit Velocity 1995.82 m/s

Exit Mach 2.837

Oxidizer Pressure Drop w/ % Pc 189995.31 5.00

Fuel Pressure Drop w/ % Pc 189995.31 5.00

% Required Minimum Thrust 120.5

At the lower oxidizer to fuel ratio, a lower minimum thrust can be attained, assuming the five percent pressure drop as the minimum stable thrust point.



2.5. Ignition System

An igniter is needed to start the combustion of the mixture of fuel and oxidizer entering the combustion chamber. Research conducted showed that the best igniter for the engine was a glow plug. A glow plug is a device that has a heating element or filament coiled up. The filament acts as a resistor, similar to a light bulb filament, and heats up when an electric current is passed through it, igniting the reactants. The glow plug resembles a short spark plug with a heating filament fitted into its tip but does not extrude out into the chamber like the spark plug would. It requires less hardware and power to operate than a spark plug, while being widely available. The glow plug that will be used is the R8 extra cold glow plug made by Axe Motor Rossi. A spark plug would produce a higher temperature which would result in an easier start of the combustion, but the voltage required to produce the spark is too high and would require cumbersome equipment. The R8 extra cold glow plug will still be able to reach a temperature for the mixture to combust while using a voltage that can be obtained through a variable power source. Assuming the glow plug must be white hot to ignite the reactants, a single glow plug requires around 2 volts and 5 amps. There will be a headlock connector on the glow plug serving as the electrical connection to the power source. Three igniters will be used to ensure consistent combustion and injection pattern, while also acting as a safeguard in the case that one or more igniters fail. The three glow plugs will be assembled in a parallel arrangement. Though this requires a larger current than a series connection, if one fails the other two will still be operable. The variable voltage source must be as close as possible to the igniters to maintain a low wire resistance while being protected by a steel barrier. The power source also has a port that will be connected to a computer to monitor the current and voltage and more importantly allow the user to control the igniters from a safe distance. Tests are to be conducted to determine a more accurate voltage and current needed to ignite an oxidizer rich propane mixture.

2.6. Backup Ignition System

The proposed backup ignition system that will be utilized will be a hybrid/solid rocket ignition system. This ignition system is composed of two wires covered with Pyrogen. There will be a current applied to the wires that will ignite the Pyrogen. The Pyrogen then ignites a piece of ammonium perchlorate from a hobby rocket motor to start the ignition. This is a crude method of igniting our engine but has proven to be effective as seen in hybrid and solid rockets. This would be taped lightly inside the engine. If the glow plugs begin the combustion, the backup system will be ejected out the nozzle. If the glow plugs do not do their job, the backup system may be used to continue the testing. This will provide ignition but will not be a restarting or reusable ignition system. This system shall only be used for the continuing of a test fire to enable the collection of performance data.



2.7. ERPL Liquid Rocket Engine Creator

Description: The Liquid Rocket Engine Creator (LREC) is a proprietary design spreadsheet. It was developed in Microsoft Excel by Embry-Propulsion Labs (ERPL). This spreadsheet is composed of eight sheets which take in some required user inputs and outputs nominal design data useful in the design of a liquid propellant rocket engine. The eight sheets are titled Inputs, Nozzle, Chamber, Throttle, Ox Injectors, Fuel Injectors, Heat Transfer, and Dimensions. Each sheet has color-coding. Yellow indicates a user input while light blue indicates a computed output. Please note that this valuable computational tool should only be used by a person skilled and knowledgeable in rocket engine fundamentals and design. Some of the data that is output may require some interpretation and many of the inputs require values from a chemical equilibrium analysis similar to NASA CEA. While any value may be entered, this does not mean it is a smart input and only personnel trained in rocketry should make these decisions. Also, this tool is still considered experimental and is being tested for the use of simple hybrid rockets as well as liquid rockets. The first sheet is the Inputs page of LREC and has six sections. General Inputs includes desired engine performance data, such as thrust and chamber pressure, and data obtained from NASA CEA Analyses. Other sections are Nozzle Inputs, Chamber Inputs, Fuel Injector Inputs, Ox Injector Inputs and Material Properties. The Nozzle sheet in LREC is where all the rocket calculations are performed to obtain dimensions, temperatures, pressures, velocities, and so on. This is divided into Throat and Exit sections. The Chamber sheet consists of outputs to rocket equations. The outputs of this sheet are similar data types as in the Nozzle sheet. Also included is a calculation of working hoop stress and wall thickness with from user inputs, including desired safety factor. The Throttle sheet is calculator for the throttling capability of the engine. This allows the user to find the thrust they can throttle down to based on CEA data input and desired pressure drop. It also outputs important design information such as flow rates, velocities, pressures, and temperature. The Ox Injectors and Fuel Injectors sheets are identical. These sheets take values from Input sheet and calculate injector data such as number of injectors, pressures, flow rates, and velocities. This sheet is useful in the dynamic design of injector configurations.



The Heat Transfer sheet is still in construction. This sheet requires a large amount of data input by the user, much of which can come from NASA CEA, and outputs the wall temperature and its difference from critical temperature at four main points through the engine. The final sheet of LREC is the Dimensions page. This sheet has two sections. In one, many useful required dimensions of the engine are calculated. In the other section, the volume and mass of propellant needed is shown as well as an estimate of the engine mass. Shown below is a list of the inputs and outputs within the LREC spreadsheet.

Inputs: General:

CEA Data: Chamber Temperature Chamber Pressure Specific Heat Ratios

Molar Masses Specific Impulse Desired Thrust Burn Time Fuel and Oxidizer Densities Oxidizer to Fuel Ratio Nozzle:

Nozzle Half Angles (i.e. 15/60) CEA Characteristic Velocity CEA Exit Pressure CEA Exit Temperature

Chamber: Characteristic Length

Injectors: Coefficient of Discharge Head Loss Coefficient Orifice Diameter Pressure Drop Material Properties: Throttling:

CEA Data: Temperatures Specific Heat Ratios

Molar Masses Specific Impulse Exit Mach Number Desired Thrust



Outputs: General: Specific Heat (constant pressure) Nozzle:

Specific Gas Constant Throat and Exit Area Exit to Throat Area Ratio Pressure at Throat Temperature at Throat Throat and Exit Diameters Velocity at Throat and Exit Ideal Exit Velocity Mach at Exit Throat and Exit Sonic Velocity Chamber, Throat, and Combined Specific Volumes Coefficient of Thrust

Chamber: Area Volume Length Mass Flow Rate Stagnation Temperature Stagnation Pressure Fuel, Oxidizer, and Combined Mass in Chamber Mach in Chamber Chamber to Throat Area Ratio Injectors: Orifice Area Individual and Total Mass Flow Rate Individual and Total Volume Flow Rate Change in Pressure Pressure into Injectors Injection Velocity Combined Injector Area Number of Injectors Throttling: Exit Pressure Mass Flow Rate Injector Velocities Injector Pressure Drops Exit Velocity Throat Velocity Stagnation Temperature Percent of Minimum Required Thrust Dimensions:



Chamber Inner Diameter Chamber Outer Diameter Chamber Length Nozzle Entrance Length Total Chamber Length

Nozzle Exit Length Total Length Nozzle Entrance, Throat, and Exit Diameter Wall Thickness

Quantities: Total Propellant Needed Fuel and Oxidizer Needed Estimated Mass of Engine Oxidizer to Fuel Ratio Check

2.8. NASA Chemical Equilibrium Analysis

the design of rocket engines. Though it can be used for many other applications, the Rocket problem is what has been used for this project. A new window appears when Rocket is selected. This allows the user to input chamber pressure(s), assign or estimate a combustion temperature, nozzle expansion ratio(s), and calculation type. The two types are infinite and finite area combustor. For preliminary engine design it may be best to use infinite area combustor and compare results for frozen and equilibrium calculations. When looking at the frozen calculations, the freezing point may be critical as it determines where CEA assumes the combustion process stops. The finite area combustor option is more appropriate for detail design as the user should have a good enough idea of the thrust chamber dimensions to use a contraction ratio. Once the user inputs their requirements of this page and saves it, the user enters desired oxidizer to fuel ratio(s) on the first page titled Problem. On the next page, titled Reactant, the user must input the desired propellants and their injection temperatures. The next three pages titled Only, Omit, and Insert allows the user to specify or ignore certain gas species to be included in calculations. These inputs are not required. The last page is Output and here the user chooses the data needed to be displayed after calculation.



3. Propulsion System Configuration

3.1. CATIA

A total of 15 parts have been modeled in CATIA, including eight original parts, three stock standard parts based off dimensions from the McMaster-Carr website, an igniter, two O-rings and one additional part for the test stand which is not included in the models below. The current assembly of the entire rocket, shown in Figure 14Error! Reference source not found., without the feed line system attachments and piping, is 296.0mm tall, 217.7mm wide at the widest part (the connection plates) and is predicted to mass about 9.3kg. In Figure 14, the combustion chamber, supersonic nozzle, four washers and the four nuts are not visible. The washers and nuts are below the bottom of the three plates near the top, and the nozzle and combustion chamber are hidden by the cooling jacket.



Figure 14: Assembled Engine



Nozzle

The supersonic nozzle, depicted in Figure 15, is 106.6mm tall; 90.2mm wide; has a thickness of 3.175mm at all points except the top flange; and a predicted mass of 0.625kg, assuming general material properties of tellurium copper. The bell is a fully expanded (100%) parabolic nozzle; the convergent region has a 60° half-angle linear convergent section. The throat has a diameter of 29.83mm, and the exit has a diameter of 68.69mm, giving an expansion ratio of 5.301. This expansion ratio was used to get the initial parabola angle of 20.178°. The top of the nozzle has a lip 1.588mm thick to allow for joining with the bottom of the combustion chamber. The lip is designed to provide a 6.35mm flat area that can be welded or brazed to the combustion chamber, reducing the probability of leaks without the need for threading (which would be difficult considering the thickness of the part, the material and the expected operating temperatures.

Combustion Chamber

As shown in Figure 16, the combustion chamber remains a hollow cylinder made of tellurium copper. The chamber has an inner diameter of 84.15mm for the whole of its 131.7mm length, but the thickness again changes for the bottom 6.35mm of the length from 3.175mm to 1.588mm. It masses approximately 0.994kg. The top of the combustion chamber contains three rows of 49 holes apiece. The holes are 0.397mm in diameter, equivalent to a 1/32 inch drill bit, and are evenly distributed around the circumference in staggered rows. These rows begin 15.875mm from the top edge of the combustion chamber and are separated by 1.588mm vertically. The volume from just above the top row of injectors to the top of the combustion chamber will be used for mounting other components.

Figure 15: Supersonic

Nozzle

Figure 16: Combustion Chamber



Cooling Jacket

Overall, the jacket is 226.2mm tall and 101.6mm wide. It incorporates a linear section to provide cooling past the combustion chamber and a tapered outline of the nozzle to regulate coolant flow speed past different parts of the nozzle. The cooling channel is just under 1.5mm wide at the base of the inlet, narrowing to 1mm near the throat and then gradually expanding to 2.27mm at the bottom of the combustion chamber. It remains 2.27mm up to the injection holes. Because this jacket will be carrying the full pressure of the fuel as it flows to the combustion chamber, 3.175mm thick steel was chosen for the material. This material should not pose problems because it will not be exposed to full combustion temperatures, but it will need to handle moderate temperatures and high pressures during operation and in the event of a catastrophic failure of the engine. Using steel of this thickness will add a measure of safety to the rmass will be about 1.57kg, making it the second most massive component on the rocket. After fabrication, the jacket will be cut axially into halves, which will be fitted around the combustion chamber and nozzle and then welded shut. 1mm tolerances represent approximately 10 times the maximum manufacturing machine tolerance in the Lehman manufacturing lab, but this narrow a channel may experience problems if the welding process produces debris in the channel.

Fuel Inlet Manifold

The fuel inlet manifold will be constructed out of red brass, which has a similar coefficient of thermal expansion as the copper nozzle (pure copper has an expansion coefficient of 16 to 17 parts per million per degree C, while red brasses are in the range of 18 to 20 and stainless steels can vary from nine to 17, depending on the alloy and amount of austenite. Considering the necessity of buying materials from scrapyards, the brass is a more consistent material without the need for identifying the steel

alloy). The manifold is 21.6mm tall and has an overall width of 109.9mm. The top and bottom are 3.175mm thick, while the side wall is 6.35mm thick. Mass is predicted to be about 0.50kg.

Figure 17: Cooling Jacket

Figure 18: Fuel Inlet Manifold



National Pipe Taper (NPT) is a standard that dictates how fluid-carrying pipes and holes must be constructed to create tight seals. Instead of holes using Unified

which slopes to close the hole the deeper it is drilled. When a pipe using NPT threading is screwed into the hole, the straight sides of the pipe are held more and more tightly by the closing hole, creating a seal which is more resistant to leaks than straight threading. Four 1/8 inch NPT holes will be drilled into the sides to accept connections to the feed line system. These holes have 27 threads per inch, are just under 10.3mm wide and require a depth of 6.35mm to form a good seal. These minimums dictate the outer dimensions and sidewall thickness of the part.

Nitrous Injection Plate

The nitrous injection plate, shown in Figure 19, will be constructed of stainless steel to prevent oxidation during exposure to the nitrous oxide. It has a diameter of 217.7mm, a thickness of 6.35mm at most points, and will mass about 1.81kg. The bottom of the plate has a 1.588mm deep, 3.175mm wide semicircular groove for an O-ring seal between it and the copper connection plate beneath. Nitrous oxide will flow through the 102 0.397mm diameter holes arrayed within the projected area above the combustion chamber. They have been spaced so as to maximize cooling through the plate while still being easily machinable. Four half-inch holes with 13 threads per inch are spaced around the disk at 83.5mm from the center, for connection with the other connection rings and the test stand. Three smaller holes for the igniters, with a diameter of ¼ inch and 32 threads per inch (Unified National Extra Fine threading), are spaced at 120 degree angles with their axes 38.5mm from the center of the disk.

Figure 19: Nitrous Injection Plate



Nitrous Injection Dome

The nitrous injection dome will be welded to the top of the nitrous injection plate, creating a hard seal to contain the nitrous oxide flowing in through the 3/4 inch NPT (21.3mm wide, 14.2mm deep) hole at the top and out through the small holes in the plate. The added tube-like section at this inlet, shown in Figure 20, is necessary to ensure proper thickness for threading and to provide a minimum depth for the NPT to make a good seal. Overall, the height will be 51.2mm and the outer diameter will be 100.1mm. The mass will be 0.321kg, using stainless steel as the material. However, the manufacturing lab may elect to increase the bottom ridge downward to make fabrication easier; this extra material at the bottom may be left on for testing purposes, since it does not impact engine operation, or it may be cut off after the rest of the part has been finished.

Copper and Steel Connection Rings

The connection rings are designed to provide the combustion chamber/nozzle assembly and the cooling jacket structural support, a point of connection between different metals at a location removed from heat (to avoid problems with thermal expansion), a point of connection and load bearing to the test stand, and help with assembly. The steel ring will connect to the outside of the top 6.35mm of the cooling jacket after it has been assembled around the combustion chamber and nozzle. The copper ring will be brazed to the top 6.35mm of the combustion chamber. Both have an outer diameter of 217.7mm, with ½ inch threaded holes to match with the holes in the nitrous injection plate. However, their inner diameters vary because of the different parts they must encompass; the copper ring has an inner diameter of 45.3mm, and the

steel ring has an inner diameter of 50.8mm. The copper ring is expected to mass 1.685kg, and the steel ring will mass 1.417kg.

Figure 22: Copper Connection Ring

Figure 21: Steel Connection Ring

Figure 20: Nitrous Injection Dome



The steel ring has a 1.588mm deep, 3.175mm wide semicircular groove at a radius of 55.9mm for an O-ring seal between it and the copper ring; the copper ring has a similar groove on its underside, and a groove with the same depth and width at 63.4mm from the center to house the O-ring between it and the nitrous injection plate.



3.2. Connections and Assembly

All 29 engine parts and a general assembly sequence are shown in Figure 23.

Figure 23: Exploded Assembly drawing

To fit the cooling jacket around the combustion chamber it will be necessary to cut the cooling jacket into two parts and then weld the two pieces back together around the combustion chamber. Centerlines can be seen showing the assembly order of the flanges (and accompanying nuts, washers and bolts) as well as the injector assembly. Part of the main engine contains several flanges that provide a mounting surface for attachment to the superstructure and also bridges between the combustion chamber/cooling jacket and the injector assembly. As these flanges will be fastened by nuts, washers and bolts, it is necessary for an O-ring or gasket to be between them, in order to prevent hot-gas leakage. For this purpose, Permatex High-Temperature Anaerobic Flange Sealant may be used. The use of this material will allow a custom gasket to be made for the current design, while no reliance on the design fitting the O-ring will be necessary. This gasket- -applied. It protects up to 400° F (205° C) in continuous exposure, and has no restrictions on use in highly oxidizing environments. Cure time from initial exposure to assembly is one hour, with full cure occurring after a period of 24 hours.



The nuts, bolts and washers to be used for fastening of the flanges may be purchased from McMaster-Carr or most hardware stores and will be of standard size and threading. The bolts (McMaster- -

inch. They will be hex-headed cap screw type of grade-5 zinc-plated steel with an ultimate strength of 120 ksi (827.37 MPa). This is preferred as a corrosion (in this case oxidation) inhibitor in case of direct contact with the hot-gases in use. Nuts used will be P/N 94805A224 or similar. Made from 316 steel, they are hex head

exterior and thus not subjected to any harsh oxidation environments. Washers to be



3.3. NASTRAN Pressure Analysis

The assembly of the entire engine was overlaid with materials for each section; the nozzle, combustion chamber and copper connection plate were overlaid with the material properties of tellurium copper; the cooling jacket, nitrous dome, injection plate and steel connection plate were overlaid with properties of steel; and the cooling manifold used the material properties of red brass. The assembly was then overlaid with pressures on both the inside, outside and in-between areas, simulating the flows of gaseous combustion at different points, liquid cooling and injection of the liquid propellants. The assembly was then separated into two sections: one section includes the plates and injector dome, and the other includes the combustion chamber, nozzle, cooling jacket and manifold.

Pressures overlaid on the assemblies are as follows: Injector Plate/Injector Dome: 716.486 psi (4.94 MPa) Inner Nozzle & Throat: 313.281 psi (2.16 MPa) Inner Combustion Chamber: 551.143 psi (3.8 MPa) Cooling: 744.044 psi (5.13 MPa)



First Assembly: Injector Plate, Injector Dome and Connection Plates

Figure 24: Section Cut: Deformation of Dome and Plates

As seen above in Figure 24, the exaggerated warping deformation that is produced from the injection pressure of the nitrous oxide will not cause any fluids to leak out, due to the O-rings placed in between the plates, along with the tight clamp produced by the bolts attached to the plates. This deformation has a maximum deflection of 0.005 in (0.127 mm) and will not affect the performance of the engine.



Figure 25: Section Cut: Mean Pressure Distribution of Dome and Plates

A mean pressure distribution, as show in Figure 25, shows the areas in which pressure from the fluid entering the assembly will be the greatest. Negative values are listed because of the opposite force and direction that the pressure is being distributed (up for positive and down for negative). This distribution map confirms the prediction that the topside of the injector plate will be subjugated to the most pressure, due to the direct, vertical injection of the nitrous oxide. The greatest amount of pressure, 79,000

process a fine enough mesh on the O-rings to produce a more accurate result.



Figure 26: Section Cut: Von Mises Stress of Dome and Plates

The structural integrity of the plates and dome, as shown in Figure 26, is consistent with predictions that the highest concentration of stress would be in the shower head area of the injector plate. High concentrations of stress would also be on the areas in which the bolts would clamp down onto the plates. These values range from 15,571 psi (107.38 MPa) to 31,148 psi (214.75 MPa). These values are well below the maximum yield stress of steel, 78,300 psi (540 MPa) and therefore, will be able to withstand the injection pressure.



Figure 27: Isometric Cut: Von Mises Stress of Plates

Figure 27 shows the stress produced on the injector plate by the nitrous oxide. The

around the O-rings. The highest concentration on the steel injector plate is significantly less than the maximum yield stress of steel and will not be structurally compromised.



Second Assembly: Cooling Jacket, Nozzle, Combustion Chamber and Manifold

Figure 28: Section Cut: Deformation of Jacket, Nozzle, Chamber and Manifold

Deformation of the second assembly, as shown in Figure 28, has a maximum of 0.00234 in (0.059 mm). It is shown in the angled section cut of the area joining the combustion chamber with the nozzle and is due to the high pressure combustion of the gases pushing towards the nozzle. The relative high pressure in the cooling jacket will help to mitigate this deformation.



Figure 29: Section Cut: Mean Pressure Distribution of Jacket, Nozzle, Chamber and Manifold

The mean pressure distribution of the assembly, as show in Figure 29, shows the areas in which pressure from the fluid and gases entering will be the greatest. Negative values are listed because of the opposite force and direction that the pressure is being distributed (up for positive and down for negative). The greatest pressure on the angled section cut of the assembly is at the throat of the cooling jacket. This can be due to the cooling fluid flowing through the jacket in such a small area, that pressing outwards will cause a great amount of force to be subjected to that part of the assembly, with ambient pressure on the outside of the assembly. The greatest pressure, 8,900 psi, can be attributed to NASTRAN being unable to process a finer mesh on certain areas of the assembly.



Figure 30: Section Cut: Von Mises Stress of Jacket Nozzle Chamber and Manifold

The structural stress that the assembly will be subjected to is shown in Figure 30. The high stress concentrations are the top area of the assembly and the area just before the throat of the nozzle. The stresses of these areas range from 10,051 psi (69.30 MPa) to 15,237 psi (105.05 MPa). For the top area of the assembly, this high concentration of stress is due to the fixed constraints. For the lower area just before the throat, this is possibly due to the pressure build-up of the cooling fluid expanding into a much larger lateral area. Please see Figure 31 and Figure 32 for details on this particular section.



Figure 31: Section Cut of Jacket and Chamber: Von Mises Stress of Assembly



Figure 32: Section Cut of Jacket and Chamber: Von Mises Stress of Jacket and Chamber

Figure 31 and Figure 32 highlight the throat area stress that the cooling jacket will be subjected to during engine firing. As listed earlier, the material that the cooling jacket will be manufactured out of will be able to withstand the stress placed on it. The stresses that the rest of the assembly will subjected to are also within acceptable levels of material yield stress: 44 ksi (303.37 MPa) for tellurium copper, and 14 ksi (96.53 MPa) for red brass. Therefore, the structural integrity of the assembly is not compromised.



3.4. Mass Budget

Table 12: Single Engine Mass Budget (without feedline)

Component Material Quantity Unit mass(kg) Mass (kg)

Combustion Chamber Tellurium Copper 1 0.994 0.994

Nozzle Tellurium Copper 1 0.625 0.625

Cooling Jacket Steel 1 1.565 1.565

Injector Plate Stainless Steel 1 1.813 1.813

Connection Ring (Copper) Tellurium Copper 1 1.685 1.685

Connection Ring (Steel) Steel 1 1.417 1.417

Fuel Inlet Manifold Red Brass 1 0.487 0.487

Injector Dome Stainless Steel 1 0.321 0.321

O-Rings Permatex Sealant 2 0.003 0.006

Bolts Steel 4 0.070 0.280

Igniter (several materials) 3 0.004 0.012

Washers Steel 8 0.008 0.064

Nuts Steel 4 0.015 0.060

Engine Total Mass 9.329

Table 12 describes the estimated masses of individual components and the total predicted mass of a single engine without any attachments or feed line components. Masses were estimated using CATIA and the application of material data to part files. Where a specific material was not available, the closest available approximation was used (e.g., tellurium copper was replaced with pure copper).

.



4. Thermal Analysis

4.1. Introduction

The cooling system consists of propane as the coolant in a jacket around the engine. There are no channels or spacing fins. There is an intake manifold at the bottom, and the propane exits the cooling jacket through the combustion chamber injection holes at the top.

Two methods were used to analyze the heat transfer through the wall: a steady state analysis and a transient analysis. Maximum wall temperature (at steady state) varied from one analysis to the other by 1.4%. This gives an average maximum wall temperature of 2600K.

Transient analysis was applied to the wall only. Using the steady state analysis, the maximum propane temperature is 782K. Nitrous oxide was also analyzed as a coolant, using only steady state analysis, which predicts a maximum wall temperature of 2898.68K, and a maximum coolant temperature of 311.58K.

4.2. Computational Tools Used

Two programs were used independently to analyze the heat flow: MATLAB and Microsoft Excel. Both programs used numerical integration techniques with an energy balance for the analysis. The engine was discretised and then analyzed. StarCCM+ was used to perform CFD analysis on the flow of the combusted gas to ensure that the geometry of the engine will not produce stagnant or extremely turbulent flow.

The heat transfer coefficients for the gas/wall interface and the coolant/wall interface were found using Equation 1 and Equation 2, respectively, for both steady state and transient analysis.

4.8.

gg

E

g

g

Equation 1: Gas/Wall Heat Transfer Coefficient

3/25/1

llll

Equation 2: Coolant/Wall Heat Transfer Coefficient



Discretization Method A coordinate system was set up as shown in Figure 33, where the r-axis is at the throat and the y-axis is co-linear with the center line. The nozzle exit is at point 5.

Figure 33: Coordinate Axis System

The engine was discretised in the y-direction and an energy balance between the combusted gas and the coolant was performed at each disc. The numerical integration starts at the nozzle exit, and a disc is defined to be the space between the inside engine wall and the inside jacket wall from yn to yn+1; yn is the beginning of the disc and yn+1 is the end of the disc. To discretise, the Mach number was set as the independent variable, varying by .001, and Equation 3 was used to determine the associated area of the engine. The

throat area, *A , is known and is assumed to be constant for this part of the

analysis.

Equation 3: Mach-Area Relationship



The radius, r, can be found directly from the area. The y-positions, which define the discs, can then be found in each section using the following equations:

1

1

1

1cos

sin

c

th

c

r

rr

ry

Equation 4: Section 1 y-position Equations

Equation 5: Section 2 y-position Equation

3

31

3

3cos

sin3

c

c

c

r

rrr

ryy


4

1

4

1cos

sin

c

th

c

r

rr

ry


Equation 8: Section 5 y-position Equation

Section 6 has a constant radius equal to the chamber radius, and the y-distance from the bottom to the top is divided into 831 discs. Sections 1, 3, and 4 are portions of a circle with radii of curvatures rc1, rc3, and rc4 respectively. rth is the throat radius, and r1 and r3 correspond to the points in Figure 33. The angle of 60 in Equation 5 is the angle between the engine wall and the y-axis.

Steady State Analysis

The gas temperature at each disc was found using Equation 9. The total temperature is the flame temperature. The gas pressure was found using Equation

10, with the total pressure being the chamber pressure. For these two equations,

was assumed to remain constant.

Equation 9: Gas Static Temperature



100

Equation 10: Gas Static Pressure

CEA was used to find the gas properties at the injectors, the bottom of the combustion chamber, the throat, and the nozzle exit as shown in Table 13.

Table 13: Variable Gas Properties

Temperature (K)

Density (kg/m^3)

Viscosity (millipoise)

Gamma Cp

(kJ/(kg*K))

Therm Conduct

(mW/(cm*K))

Injector 3317.83 4.45734 0.99321 1.1677 3.3285 7.0457

Combustor End 3309.22 4.3193 0.9914 1.1677 3.3237 7.0228

Throat 3082.11 2.7089 0.94345 1.1757 2.88 5.5975

Exit 1707.12 0.17735 0.62716 1.2534 1.6351 1.5219

The NIST (National Institute of Standards and Technology) website was used to find the properties for propane in 1K increments. The properties that were varied are density, viscosity, specific heat capacity, and thermal conductivity. Each property was found for the associated propane temperature, to the nearest 1K, rounded down. For instance, if the propane temperature is 400.35K, the properties would correspond to a temperature of 400K. The first set of properties is associated with the initial coolant temperature. The following equations describe the energy balance between the gas and the coolant.

Equation 11: Heat Transfer Between the Gas and Wall

wcwg

w

w TTt

q

Equation 12: Heat Transfer Through the Wall

Equation 13: Heat Transfer Between the Wall and Coolant

Equation 14 describes the temperature rise of the coolant across each disc for a given amount of heat transferred to it, where hg and hc are the heat transfer

coefficients for the gas and coolant, respectively, is the thermal conductivity of



the wall, tw is the thickness of the wall, Cpc is the specific heat capacity of the wall, and A is the coolant to wall surface area of the disc.

12 ccc TTCpmqA

Equation 14: Coolant Temperature Rise

There are three unknown quantities between Equation 11, Equation 12, and Equation 13; the gas wall temperature, the coolant wall temperature, and the heat flow, Twg, Twc, and q respectively. By combining the three equations, the following relationships are found to give the gas and coolant wall temperatures at the beginning of the disc, where Tc is the coolant temperature at the beginning of the disc:

gw

w

cw

w

gw

w

cw

w

gw

w

c

cw

w

g

wg

hthththt

htT

htT

T

11

1

Equation 15: Gas Wall Temperature

cw

w

c

cw

w

wg

wc

ht

Tht

T

T

1

Equation 16: Coolant Wall Temperature

By combining Equation 13 and Equation 14, the coolant temperature at the end of the disc is found:

111

2 c

c

cwcc

c T

Cpm

ATThT

Equation 17: Coolant Temperature at End of Disc

The surface area, A, of the coolant/wall interface is found by using Equation 18.

2

1

Equation 18: Surface Area of a Body of Revolution

The y=f(r) functions, from Equation 4 through Equation 8, were solved for r in order to get the derivative with respect to y. The right hand side of the r=f(y)



equation is substituted into Equation 18 for r, and the integral is carried out. The only section this was not conducted for was Section 5. A closed form solution for the integral could not be found. In that case, Equation 19 was used.

Equation 19: Section 5 Surface Area

Here, ds is the arc length of the disc and r is the average radius of the disc.

Numerical Integration The integration process started at the nozzle exit, where the coolant enters the cooling jacket, and the y-position is y1. Tc1 for this first iteration is the initial coolant temperature, assumed to be 310K. The heat transfer coefficients for the gas and coolant were calculated for y1. Gas and coolant wall temperatures were then computed, and then the coolant temperature at the end of the disc, or at y2, was calculated. The heat transfer coefficients were found for y2, and then the wall temperatures, and then the coolant temperature at y3. This process continued through all of the discs. Both MATLAB and Microsoft Excel were used to perform this integration. One main M-file was written in MATLAB, which called two separate M-file functions. The main function first calculated the y-positions and associated gas temperatures along the entire length of the engine. These were indexed in a master matrix, which was then sent to the two separate functions. One function read the gas properties of Table 13 from an Excel spreadsheet, performed linear interpolations to find the properties corresponding to each of the indexed gas temperatures, and then indexed these properties in the master matrix. The other function read the propane properties from another Excel spreadsheet, found the properties for each indexed temperature, and indexed them in the master matrix. The main file then looped through the calculations of heat transfer coefficients, wall temperatures, and coolant temperatures for one less iteration than the number of y-positions. Coolant pressure was also calculated, using Equation 20, and indexed.

gHD

VyyPP ccc

2

64.1Relog82.1 2

12

2

1012

Equation 20: Coolant Pressure

The required inputs for the MATLAB program are shown in Table 14. The outputs are the wall temperature, coolant temperature, and coolant pressure profiles along the engine, as shown in Figure 34, Figure 35, and Figure 36 respectively.



Table 14: Analysis Inputs

Gravity (m/s) 9.81 Throat Radius (m) 0.014955

Initial Coolant Press (MPa) 5 Exit Radius (m) 0.032425

Chamber Pressure (MPa) 2.9 Section 6 Length (m) 0.13495

Max Section 1 Arc (deg) 60 Nozzle Length (m) 0.0804

Max Section 3 Arc (deg) 60 Section 2 Angle w/ Vertical (deg) 60

Flame Temperature (K) 3317 y4 -0.00193

1.19 Total Chamber Length (m) 0.13476

Gas MW 24.8 Initial Coolant Temperature (K) 310

RU (J/kgK) 8314 Coolant Mass Flow (kg/s) 0.208

Sec. 1 Radius of Curvature (m) 0.022352 Wall Thermal Cond. (W/mK) 350

Sec. 3 Radius of Curvature (m) 0.003175 Wall Thickness (m) 0.003175

Sec. 4 Radius of Curvature (m) 0.005258 Jacket Thickness at Exit (m) 0.0015

Chamber Radius (m) 0.0423 Jacket Thickness at Throat (m) 0.001

Coolant Volumetric Heat Capacity (J/m3K) 3441900 Jacket Thickness at Chamber (m) 0.002273

Figure 34: Engine Wall Temperature from MATLAB



Figure 35: Coolant Temperature from MATLAB



Figure 36: Coolant Pressure from MATLAB

The Excel integration method used the same inputs as in Table 14. The outputs are the wall temperature, coolant temperature, and coolant pressure as shown in Figure 37, Figure 38, and Figure 39 respectively.



Wall Temperature vs. y-Position

0

500

1000

1500

2000

2500

3000

-0.1 -0.05 0 0.05 0.1 0.15

y-Position (m)

Figure 37: Gas Wall Temperature from Excel



Coolant Temperature vs. y-Position

0

100

200

300

400

500

600

700

-0.1 -0.05 0 0.05 0.1 0.15

y-Position (m)

Figure 38: Coolant Temperature from Excel

Coolant Pressure vs. y-Position

4.65E+06

4.70E+06

4.75E+06

4.80E+06

4.85E+06

4.90E+06

4.95E+06

5.00E+06

5.05E+06

-0.1 -0.07 -0.04 -0.01 0.02 0.05 0.08 0.11 0.14

y-Position (m)

Figure 39: Coolant Pressure from Excel



Transient Analysis The discretisation method, coolant properties, and gas properties of the steady state analysis were also used for the transient analysis. The wall in each disc was divided into 5 even sections, with each boundary of the wall coinciding with the end (or start) of a section, as shown in Figure 40.

Figure 40: Transient Analysis Diagram

The analysis began with the general transient heat transfer relationship for a material with temperature independent properties, Equation 21.

2T Wa T

t c

Equation 21: Differential Transient Heat Transfer Equation

From here, it was assumed that there were no heat sources ( =0), and heat is transferred only in the x-direction. This reduces the gradient operator to a single partial differential, simplifying Equation 21 to the form shown in Equation 22.

2

2

Equation 22: Simplified Transient Heat Transfer Equation

Here, a is the thermal diffusivity of the wall. From here the finite difference numerical analysis technique was applied. The partial differentials in Equation 22 were replaced with difference quotients where the generalization that a derivative is equivalent to the difference quotient plus a discretisation error was used. Equation 23 and Equation 24 describe the transformation of the two partial differentials, in Equation 22, into difference quotients.



Equation 23: Difference Quotient for Time Partial Derivative

2

2

11

2

2 k

i

k

i

k

i

k

i

Equation 24: Difference Quotient for Position 2nd Order Partial Derivative

f) function is the discretisation error. It approaches zero proportionally to f.

The position is noted by the subscript i while the time step is denoted by the superscript k. Substituting Equation 23 and Equation 24 into Equation 22, with some algebraic rearrangement, gives the finite difference numerical solution for transient heat transfer through the material, in this case the wall.

Equation 25: Finite Difference Numerical Solution Through the Wall

2x

taM

Equation 26: Modulus of the Finite Difference Formula

Equation 25 can only be used for x-positions that are entirely within the wall. For positions X1 and X6, the boundary conditions must be observed, which is heat transfer conditions for both. Equation 27 is the differential heat transfer boundary condition and Equation 28

sTTn

T

Equation 27: Differential Heat Transfer Boundary Equation

Equation 28: Difference Quotient for Heat Transfer Boundary Condition

Equation 28 is used with i = 1 or i = 6 and the (+/-) indicates the difference between the right and left hand boundaries since n is the outward normal to the surface. Referencing Figure 40 for the positive x-direction, the left hand boundary gives a negative n while the right hand boundary gives a positive n. Applying the heat transfer boundary conditions to Equation 25 yields the numerical solutions for the boundaries of the wall.



Equation 29: Coolant Boundary Finite Difference Numerical Solution

k

s

k

g

kk

65

1

6

Equation 30: Gas Boundary Finite Difference Numerical Solution

f

f

Equation 31: Biot Number for Finite Difference Method

Here, Ts is the temperature of the fluid for the specified boundary. The Biot number describes the heat transfer from the fluid to the wall, and hf is the heat transfer coefficient for the specific fluid, as described in Equation 1 or Equation 2thermal conductivity of the wall. A major consideration when using the finite difference method is the stability of the equations. It is considered stable if the associated errors for the equations reduce with each iteration. This is achieved by ensuring that all coefficients are non-negative. If a coefficient is zero, the equation could be stable or it could be unstable. Equation 25 is used for i = 2 to i = 5; Equation 29 is used at the coolant boundary; Equation 30 is used for the gas boundary. For these three conditions, the stability criteria are:

Equation 32: Wall Stability Criteria

Equation 33: Coolant Boundary Stability Criteria

g

Equation 34: Gas Boundary Stability Criteria

Due to these stability criteria, the time step cannot be arbitrarily selected. Rather, since the modulus M is a function of t, the stability criteria must be used to solve

t. Too large a time step will result in the stability criteria becoming negative. To avoid this, the time step for all three conditions must be solved for, and the smallest one used. However, for this analysis, it was found that the smallest time step was always associated with the gas boundary. It was also determined that a stability criteria equal to zero created an unstable equation. Therefore, the gas boundary time step was divided by two. Because the gas boundary stability criteria was used for each disk to find the time step, the gas stability remains constant while

Figure 41.



Figure 41: Time Step and Stability Data for Each Disk

The number of time steps required to reach steady state temperatures was determined experimentally, arriving at 12000. Figure 42 shows the gas/wall interface temperature calculations at the nozzle exit over the 12000 time steps. It can be seen that the temperature has not completely settled into a steady state. However, given the amount of time it takes to run a simulation at higher numbers of time steps, the level of steady state reached is acceptable. The total time analyzed was 11.8s.



Figure 42: Transient Gas/Wall Interface Temperatures for Disk 1

CFD StarCCM+ was used to perform CFD on the gas through the engine. This verified the assumption that the gas flow through the engine has no stagnation points and no extremely turbulent points. The model was built to simulate the approximate mass flow and flow velocities that are expected in the engine. Figure 43 shows the overall configuration of the model. The inlet is simply a circular section with the outward normal parallel to the y-axis, centered in and at the very top of the combustion chamber, and half the area of the combustion chamber. It was modeled as a mass flow inlet with a mass flow of 1kg/s. The exit was modeled as a pressure outlet at 0.0Pa (gage). The flow was modeled using K-Epsilon turbulence, ideal, multi-component, non-reacting gas, and coupled flow models. The composition of the gas was H2O, CO2, and N2 with the respective mass fractions 0.149, 0.272, and 0.579. Figure 44 and Figure 45 show streamlines through the engine. For both figures, the solid streamlines correspond to the Velocity: Magnitude scale while the smaller arrows correspond to the Cell Relative Velocity scale.



Figure 43: CFD Simulation Configuration



Figure 44: Gas Streamlines through Engine, Axial View



Figure 45: Gas Streamlines Through Section 2 Region, Top View

4.3. Discussion, Conclusions, and Limitations

Steady State Analysis The two methods give outputs that agree. MATLAB outputs the maximum wall temperature to be 2618K and the coolant temperature to be 782K at nominal, while the Excel spreadsheet outputs 2509K and 632K respectively. The Excel approach assumes the coolant/wall surface area of the discs to be cylindrical as in Figure 46 (1), but the surface area is actually parabolic or conic (except for Section 6, which is a cylinder) as in Figure 46 (2).



Figure 46: Conic vs. Cylindrical Surface Area

The surface area of a cylinder is less than that of the cone or parabola. The more surface area results in more heat transfer which would increase the temperature of the propane, which is reflected in the higher propane temperature of the MATLAB approach. Also, the MATLAB approach uses a linear interpolation to find all of the

Cpc, and c, an exponential equation for the density, and a linear equation for the viscosity. The jacket thickness at the nozzle exit is 1.5mm, 1mm at the throat, and 2.273mm at y2. The Excel approach uses a cubic equation to interpolate the jacket thicknesses between those points while the MATLAB approach uses a linear relationship. As Figure 47 and Figure 48 show, the jacket thickness around the throat area is larger for the Excel approach.



Figure 47: Jacket Thickness from MATLAB

Jacket Thickness for Excel

-0.08

-0.03

0.02

0.07

0.12

-0.1 -0.05 0 0.05 0.1 0.15

Radius (m)

Jacket Wall

Engine Wall

Figure 48: Jacket Thickness from Excel



The larger area around the throat would decrease the cooling in that region, whereas the thinner area around the throat area for the MATLAB approach would increase cooling. This is reflected in Figure 34 by the decrease in wall temperature just before the throat, and then increasing again. The reason for the zig-zag in Figure 34 is not known. However, the shape of the graph makes sense and the values agree with the Excel approach. The hump in the Figure 37 graph in the chamber y-position region is explained by the increase in jacket thickness at the top of the chamber as seen in Figure 48. The coolant pressure graphs are very similar. The greatest pressure drop is expected to happen where the velocity is greatest. In both Figure 36 and Figure 39, the pressure decreases most rapidly in the throat region, which is where the velocity is the greatest. The total drop in pressure from MATLAB and Excel are .2945MPa and .2957MPa respectively. Nitrous oxide was investigated as a coolant. It would be pressurized to 5.6MPa in the tank, which at an estimated ambient temperature of 310K results in the nitrous oxide being in the gaseous phase. The thermal conductivity of nitrous oxide was determined using Excess Thermal Conductivity Correlations. Using Roy and Thodos Estimation Technique, the inverse thermal conductivity of nitrous oxide is determined at the critical temperature and pressure.

1

3 6

4210 C

C

T M

P

Equation 35: Roy and Thodos Estimation Technique

where = inverse thermal conductivity (m*K/W)

CT = critical temperature (K)

= critical pressure (MPa)

M = molecular weight (kg/mol)

Using Stiel and Thodos generalized equations for excess conductivity correlations, the thermal conductivity is determined.

Equation 36: Generalized Excess Conductivity Correlation

where = Inverse Thermal Conductivity (m*K/W)

= nitrous oxide gas constant

= Thermal Conductivity (W/kg*m)

= Known Thermal Conductivity at a low pressure (W/kg*m)



0.03

0.04

0.05

0.06

0.07

0.08

0.09

300 400 500 600 700 800 900 1000

Temp (K)

Figure 49: Nitrous Oxide Thermal Conductivity

When nitrous oxide is at a temperature greater then 309.5 K and a pressure lower then 7.2 MPa it is a gas. After entering the gas phase the thermal conductivity is 0.033 W/m*K. The thermal conductivity of nitrous oxide at a constant pressure increases as the temperature is increased.

The specific heat capacity was calculated using Equation 37 .

2 3 40 1 2 3 4a a T a T a T a T

p R

Equation 37: Specific Heat Equation

where = specific heat J/(mol*K)

T = temperature (K) R = gas constant a = specific heat capacity temperature coefficients



39

41

43

45

47

49

51

53

55

300 400 500 600 700 800 900 1000

Temp (K)

Figure 50: Nitrous Oxide Cp

As temperature increases the specific heat capacity increases. At 310 K, the Cp is 38.9 J/(mol*K). As the temperature reaches 900 K, the Cp of the nitrous oxide is 52.9 K.

The density of nitrous oxide was determined using the ideal gas law.

30

40

50

60

70

80

90

100

300 400 500 600 700 800 900 1000

Temp (K)

Figure 51: Nitrous Oxide Density

As the temperature of nitrous oxide increases the density decreases. When the temperature reaches 900K, the density is 32.3 kg/m^3.



While the thermal properties of nitrous oxide in the gaseous phase are actually better than propane in the liquid phase, the low density compared to propane means that it will not absorb as much heat. Also due to the low density, and to accommodate the required mass flow, the cooling jacket radius was increased by 2mm in order to reduce the pressure drop to 1.43MPa. The analysis predicts a maximum wall temperature of 2898.68K and a maximum coolant temperature of 311.58K.

Table 15: Propane Thermodynamic Properties

Temperature (K)

Thermal Conductivity (W/m*K)

Density (kg/m^3) Cp

(J/(mol*K) Viscosity

300 0.097 502.61 115.9 104.1

500 0.053 70.8 126.4 15.6

700 0.082 22.37 128.54 18.3

Table 16: Nitrous Oxide Thermodynamic Properties

Temperature (K)

Thermal Conductivity (W/m*K)

Density (kg/m^3) Cp

(J/(mol*K)

300 0.033 97.1 38.9 133.2

500 0.047 58.2 46.1 152.2

700 0.064 41.5 50.5 182.7

Transient Analysis

The maximum wall temperature predicted by the transient analysis is 2581.15K. The wall temperature will reach the melting point for the copper engine in 0.2516s, at a y-position of 0.005905m. The wall temperature will reach the maximum service temperature of the copper engine in 0.0615s, at a y-position of 0.004918m. Figure 52 shows the time to reach the maximum service temperature for each disk.



Figure 52: Time to Reach Critical Wall Temperature for Each Disk

A comparison of the steady state analysis wall temperature and the transient analysis at steady state wall temperature profiles is shown in Figure 53.



Figure 53: Steady State and Transient Wall Temperature Comparison

The transient temperatures are slightly lower than the steady state temperatures. The difference between the maximum wall temperatures is 1.4%; 2581.15K from the transient analysis and 2618.27K from the steady state analysis. The shape and slopes throughout the two graphs are nearly identical. There is an anomaly in the transient temperature profile graph between the throat and 0.05m. The reason for this anomaly is unknown. However, the high level of agreement supports the validity of the analyses.

CFD CFD was conducted to inspect the flow characteristics of the gas through the engine. It is desired to have the flow steady throughout the engine. A turbulent flow would decrease engine performance and increase heat transfer to the wall. Stagnant flow would also decrease engine performance. Figure 44 and Figure 45 show that the streamlines are steady and there are no stagnation points. From CFD analysis, Figure 43, the Mach number at the throat is 1 while the Mach number at the nozzle exit is about 2.5. From the isentropic relationships used in the transient and steady state analyses, the Mach number at the throat is 1 and the Mach number at the nozzle exit is 2.8. The exit Mach number from the CFD analysis is 10.7% different from the heat transfer analysis. For the purpose of verifying that the engine geometry will not impede the gas flow, this percent difference is acceptable.



Conclusions The cooling jacket is not able to cool the engine to temperatures that the material can withstand. This is expected because the mass flow of coolant is very small. This makes the temperature of the coolant increase rapidly, and limits the velocity of the coolant to less than 7m/s everywhere. Propane properties become less desirable for cooling as the temperature increases, and the cooling effectiveness decreases as the flow velocity decreases. If pressurant pumps were used, the pressure of the coolant as it enters the cooling jacket could be increase substantially which would allow for a greater coolant pressure drop and hence greater coolant velocities. Also, the flame temperature for the combustion of propane and nitrous oxide at a 7:1 fuel-to-oxidizer ratio is 3317K. This requires the cooling system to cool the wall by 2567K, assuming a maximum service temperature of the material to be 700K. A lower fuel-to-oxidizer ratio would reduce the flame temperature which would reduce the maximum wall temperature. For an O/F ratio of 3:1, the maximum wall temperature falls to 1300K. This is approximately a 50% reduction in wall temperature; however it is still at the melting point and well above the maximum service temperature of the engine. The reduced O/F ratio would also decrease the ISP by 18%. One possible solution would be decrease the number of engines on the vehicle and increase the size of subsequent engines. This would increase the mass flow of coolant and increase the surface area for heat transfer. The use of nitrous oxide as a coolant is not a viable option. The maximum wall temperature is 300K higher than the propane cooled engine, and the size of the cooling jacket would have to be increased to maintain acceptable pressure losses which would increase the weight of the engine.

Limitations of the Analyses The cross sectional area that the coolant sees as it travels through the cooling jacket is defined to be perpendicular to the velocity vector. However, in both the MATLAB and Excel analyses, the cross sectional area used is perpendicular to the y-axis. This is correct for Section 6 only. To correct this, the surface area would have to be found using an equation similar to Equation 18, which was used to find the surface area of the coolant/wall interface. The r=f(y) function would have to be derived to define the line perpendicular to the tangent of the engine wall and extending out to the jacket wall; this would be the vector representing the jacket thickness. Similarly, the jacket radius for both analyses was found by adding the jacket thickness to the engine wall radius along the r-axis. The correct jacket radius is found by adding the jacket thickness in the direction of the jacket thickness vector described above.



5. Manufacturing Plan

The test and production engines for the Lunar Lander are being manufactured solely by the Embry-Riddle Aeronautical University CNC Manufacturing Lab. In-house production limits budgetary strain while expediting communication and feedback between fabrication and design revisions. The ERAU Manufacturing Lab has a lathe and 4-Axis mills; both of these machines are fully automated and most of the fabrication involves them. In an attempt to cut down on fabrication time, Brian Ruby has been given the task of fabricating the intake manifold dome, nitrous injection plate, and the copper and steel connection rings.

5.1. Raw Materials and Hardware

The raw materials and equipment used in the manufacturing process that have already been received include:

Manufacturing hardware

Figure 54: C14500 Tellurium Copper Round Solid Bar



Figure 55: Stainless Steel Round Solid Bar

Figure 56: Brass Round Solid Bar

Figure 57: Steel Round Solid Bar



As of this report, the following equipment has not yet been received:

O-ring sealant paste

5.2. Fabrication Process

Thrust Chamber Fabrication Process

Given the configuration of the supersonic nozzle, it is not possible to machine the entire thrust chamber from a single piece of copper stock. The bore bit used on the lathe in the ERAU Manufacturing lab cannot be controlled effectively at lengths exceeding 2.5 inches, the bore bit chatters uncontrollably after this depth. In the absence of a decision to shorten the combustion chamber length, a piece-wise fabrication is the best manufacturing approach available.

The ERAU Manufacturing Laboratory provided a tentative list of tools and fixtures needed during the fabrication of the various parts for the rocket engine. Most of the tools to be used on either the lathe or mill are already on site, but the fixtures will have to be manufactured in accordance to the part being fabricated and the manufacturing process adopted for that part.

Tools List

Boring bars

Rough and finish bits

Ball end mill

Square end mill

Drill bits

Welding equipment

Fixtures List

Mandrel, supersonic nozzle

Mandrel, cooling jacket



Figure 58: End Mills

Figure 59: Boring Bars

Figure 60: Rough and Finish Bar



Figure 61: Mandrel Piece (Lower)

Figure 62: Mandrel Piece (Upper)



Supersonic Nozzle Fabrication

The nozzle will be fabricated from the C14500 Tellurium copper round solid bar. The mandrel for this part will be fabricated from aluminum. All machining for this part will be performed on the lathe. The fabrication process for this part is as follows:

Lathe

Turn the aluminum on the lathe, cutting out the outer profile of lower mandrel piece

Cut hole in the center of the lower mandrel

Turn aluminum on the lathe, cutting out the outer profile of upper mandrel piece

Cut hole in the center of the upper mandrel

Turn out aluminum rod

Thread both ends of aluminum rod

Slide rod through both halves of the mandrel and bolt both ends

Bore copper (intake end), till half way

Make flange on the inlet side to enable chucking

Dismount piece and flip around to bore the exit side of the nozzle

Mount entire piece in mandrel

Turn the outer profile of the part

Cut out the flange used to chuck the part

Approximate fabrication time: 1 1/2 weeks

Status: Entire mandrel assembly/ supersonic nozzle completed

Figure 63: Mandrel Assembly



Figure 64: Supersonic Nozzle (front elevation)

Figure 65: Supersonic Nozzle (looking down at top)



Combustion Chamber Fabrication

The combustion chamber will also be fabricated from the solid round copper bar. Machining will be done on both the lathe and the 4-Axis mill. The fabrication process for the combustion chamber follows:

Lathe

Bore inside profile till halfway

Flip part around, bore other end

Chuck part and turn outside profile

Mill

Bore holes through from the outer wall

Approximate fabrication time: 1/2 week

Status: Combustion chamber completed

Figure 66: Nozzle and Chamber



Cooling Jacket (Chamber Portion) Fabrication

The cooling jacket will have to be fabricated in two separate pieces, the nozzle jacket and the chamber jacket. The cooling jacket assembly will be made out of steel. All work on the cooling jacket will be done on the lathe.

Lathe

Bore inside profile till halfway

Flip part around, bore other end

Chuck part and turn outside profile

Cooling Jacket (Nozzle Portion) Fabrication

A mandrel will have to be fabricated to machine the nozzle end of the cooling jacket

Turn the aluminum on the lathe, cutting out the outer profile of lower mandrel piece

Cut hole in the center of the lower mandrel

Turn aluminum on the lathe, cutting out the outer profile of upper mandrel piece

Cut hole in the center of the upper mandrel

Turn out aluminum rod

Thread both ends of aluminum rod

Slide rod through both halves of the mandrel and bolt both ends

Bore steel (intake end), till half way

Make flange on the inlet side to enable chucking

Dismount piece and flip around to bore the exit side of the nozzle

Mount entire piece in mandrel

Turn the outer profile of the part

Cut out the flange used to chuck the part

Estimated fabrication time: 2 weeks

Status: Nozzle portion completed



Figure 67: Cooling Jacket (nozzle portion)

Injection Plate Fabrication

The injector plate will be fabricated from stainless steel. Machining for the injector plate will be done on both the lathe and the mill. The fabrication process for the injector plate is as follows:

Lathe

Turn the inner profile of the part

Face both edges on the exit side of injector plate

Chuck top side of part and bore the exit side of the plate

Bore center hole from the exit side of the plate

Flip part around and machine opposite side to desired width

Mill

Make fixture to hold part in the mill

Cot out outer profile

Drill desired hole patterns

Estimated fabrication time: 1/3 week

Status: No work commenced

Connection Ring (Steel) Fabrication

This connection ring will be made of steel, and a required fixture will be made of aluminum. All the machining for this part will be performed on the mill. The fabrication process for the connection ring is as follows.



Mill

Drill required hole patterns

Machine fixture

Cut circular profile out

Cut grooves for O-rings



Connection (Copper) Plate Fabrication

This connection ring will be made of copper, and a required fixture will be made of aluminum. All the machining for this part will be performed on the mill. The fabrication process for the connection ring is as follows.

Mill

Drill required hole patterns

Machine fixture

Cut circular profile out

Cut grooves for O-rings



Injector Dome Fabrication

The nitrous injection dome will be made out of stainless steel. Machining for this part will be done on both the lathe and mill.

Lathe

Turn top side profile

Mill

Flip part around, cut inside profile

Cut notches on the mill

Cut central hole

Status: Work has commenced, inner profile completed



Figure 68: Injection Dome (while being machined)

Fuel Manifold

The propane intake manifold will be made out of brass. Machining for this part will be done on the lathe and mill.

Lathe

Turn outside profile of the manifold

Bore inside profile

Turn notch holder inside of the doughnut

Cut doughnut at the top with cutter bit

Face edges of ring

Cut ring off with cutter bit

Mill

Drill out desired hole patterns

Status: Work commenced, lathe work finished and mill work pending



Figure 69: Manifold Ring

Figure 70: Propane Manifold Assembly



Test Stand

A test stand was manufactured using stock steel from the ERAU Manufacturing Lab. The frame of the test stand will be constructed by welding together steel cut to the specifications of the design. The front support beams are bolted to the frame to allow ease in the replacement or exchange of support beams for multiple tests.

Figure 71: Test Stand

Connection Block Four connection blocks were made out of stainless steel and will be mounted at the end of the cantilever beams at the top of the test stand.

Figure 72: Connector Blocks



5.3. Assembly Process

The final rocket engine is a complex piece of equipment with many locations that need adjoining. It is important to make sure these joints are securely sealed as to prevent failure. The techniques of brazing, welding, and bolting will be used to hold pieces together.

Parts List

Table 17: Fabricated Parts List

Part Identification No. Part Description Raw Material Quantity

14 Fuel Inlet Manifold Red Brass 1

12 Main Engine Nozzle Tellurium Copper 1

11 Combustion Chamber Tellurium Copper 1

13 Engine Cooling Jacket Steel 1

9 Connection Ring Steel 1

8 Bottom O-Ring Gasket Sealant Paste 1

7 Connection Ring Tellurium Copper 1

6 Top O-Ring Gasket Sealant Paste 1

5 Nitrous Injection Plate Stainless Steel 1

1 Injection Dome Stainless Steel 1

Table 18: Hardware and COTS List

Part Identification No. Part Number Part Description Quantity

10 94805A224 McMaster-Carr Nut 4

4 AXE10008 Rossi R8 low Plug Igniter

3

3 91083A033 McMaster-Carr Washer 8

2 92865A356 McMaster-Carr Bolt 4

5.4. Part Assembly

The following steps are necessary to assemble the engine:

1. Braze the copper nozzle to the copper combustion chamber. 2. Weld the steel ring to the top half of the steel cooling jacket. 3. Cut the bottom half of the steel cooling jacket (the section covering the nozzle) in

half lengthwise (top to bottom). 4. Take the two bottom pieces of the cooling jacket and place them around the

copper nozzle. Weld them back together along the lines that were just cut.



5. Slide the top half of the cooling jacket over the copper combustion chamber, with the steel ring on top. Weld the top and bottom halves of the cooling jacket together.

6. Braze the copper ring to the top of the copper combustion chamber. 7. Braze the fuel intake manifold to the bottom of the copper nozzle. 8. Weld the fuel intake manifold to the bottom of the steel cooling jacket. 9. Place the stainless steel injector plate on top of the copper plate. 10. Bolt the stainless steel injector plate to the copper plate and the steel plate. 11. Weld the injector dome into the grove in the stainless steel injector plate.

For all of the brazing steps, mandrels must be manufactured to h



6.2. Electrical System for Solenoid Valves

The solenoid valves will be wired parallel to each other, guaranteeing that each valve will be able to receive the required amount of voltage (120 VAC) to operate and open. Indicator lights will be provided by Michael Potash, an associate professor at Embry-Riddle Aeronautical University, and will be wired parallel to each solenoid valve. This will guarantee that the solenoids are indeed open and letting propellant through the feedline system when the solenoids are powered. Figure 73 represents the assembled system. Although there is only one solenoid and indicator light in the diagram, additional pairs can be added to the system in parallel. Toggle switches will also be connected to the system before each pair, guaranteeing manual control over the system for each solenoid. This is due to safety concerns and dangers inherent with the rocket engine. A master toggle switch will be connected to the power source for both activation and switch off of the power flowing through the whole system. Should anything out of the ordinary arise during the rocket engine test, the switch and power can be easily cut off, thereby shutting the solenoid valves.

Wiring for the system will compose of extension cords running from the master control unit to each of the solenoids. The solenoid ends will be replaced with Edison connectors to cut down on components that can go wrong. The criteria for

Figure 73: Electrical System for Solenoid Valve and Indicator Light



the wiring system must be at minimum 18 gauge and insulated for 120VAC. Extension cords are insured to be rated for this.

6.3. Test Stand

Strain Gain Overview

One of the main purposes of the test stand is to measure the thrust produced by the engine. This is achieved by installing strain gages to the four beams that secure the engine to the test stand. During a test fire, force will be applied to the four beams on the test stand. As a result, one side of the beam will be in tension, and the other side will be in compression. When strain gages are attached to the beam, the tension or compression will cause the resistance of the strain gage to vary. This change in resistance will result in different output voltages, which will ultimately correspond to the thrust of the engine. The strain gage output voltages are measured and recorded by the DAQ system.

Configuration

A total of twenty Omega SGT-1/350-TY11 strain gages were purchased by Task Force Orion, and sixteen of the strain gages were installed on the test stand. On each beam, four strain gages were attached: two on top of the beam and two on the bottom of the beam. These four strain gages are wired together to form a Wheatstone bridge. The Wheatstone bridge was determined to be the best configuration for the strain gages because of its increased precision and accuracy measurements. During test fire, two of the strain gages will be in compression, and two will be in tension; the orientation of the engine fire (up or down) will determine which strain gages will be in tension or compression. The strain gages will be excited and their output voltage measured via the wires of the previous test stand. These wires are excellent for this purpose because they are insulated with a metal coating that shields the signal from outside noise. The wires will connect each Wheatstone bridge configuration to an amplifier; totaling four separate amplifiers for each of the four Wheatstone bridges. The amplifiers will convert the minute voltage produced by the Wheatstone bridge into a voltage appropriate for the DAQ system. A computer can then convert these voltage values into thrust values.

Calibration

In order for the software to accurately convert the voltage read by the strain gages into thrust, the system must be calibrated. This will be achieved by taking known

output voltage. The calibration will be in weight increments between zero and three hundred pounds. relationship between these two values. The relationship should be linear. In the case that it is not, the plot will be mapped to regions of linearity. Once this is achieved, the software can accurately measure the thrust produced by the engine, up to three hundred pounds.

Mounting Blocks



To properly attach the engine to the test stand, the engine will be bolted to bearing-based connectors that will slide in and out to accommodate the engine in the event that there is an error in sizing the pieces. These bolts have been scaled to meet the needs of the more powerful engine. The implementation consists of a simple arrangement of aluminum mounting blocks that will accept quarter-inch bore ball bearing rod ends, utilizing a sliding threaded standoff in order to accept multiple testing configurations. Rod ends will screw seamlessly into the aforementioned standoffs, which in turn could slide freely within the mounting blocks to allow for extension of each support. An overview of the RBC Bearings Product Catalogue determined that the rod end model M8CR fit the requirements of the engine design. The mentioned part is a commercial extra capacity rod end manufactured of industrial steel, and is more than capable of tolerating the full thrust of the Cynthion test engine. The increased overall size of the M8CR rod ends presented a problem at first, as threaded

However, Unicorp Inc carries both the required thread size as well as standoff manufactured from stainless steel, making their product more than tolerant for this application.

The mounting blocks will be manufactured in-house, as no supplier could be found. ERAU machine shop coordinator Richard Hedge created larger and stronger mounting blocks from solid steel, resulting in an assembly that will tolerate the expected stresses without need for concern. The CATIA drawing for this mounting back is shown in Figure 74.

Figure 74: Steel Mounting Block

6.4. Data Acquisition (DAQ)

DAQ



There are three types of data that will be recorded: thrust, pressure, and flow rate. Thrust will be measured using the strain gauges on the test stand mentioned earlier. Pressure will be measured with pressure transducers, mentioned in the feed line system section. Flow rate will be monitored using load cells. All of these methods transduce their respective physical phenomenon into a voltage. Data acquisition hardware is then used to measure, digitize, and transmit this voltage to a computer where it can be interpreted. The hardware to be used is a Measurement Computing DAQ.

Ideally, the DAQ would be programmed to record data at a specified sampling rate. While the Measurement Computing software can record data at a specified rate, this software will not be used. Instead, a LabVIEW program will read the DAQ. The reason for this is that LabVIEW provides the functionality of a full fledge language, which is required for testing purposes. Unfortunately, there is no means to set a sampling rate for the DAQ using LabVIEW. To remedy this, the software records data as fast as it can; the speed at which the software records information depends

Using this method, as each sample is recorded, it is also given a time stamp. Test results can then easily be interpreted with these time stamps. As long as no additional programs are activated on the testing computer during firing, then this should provide a sampling rate of at least 1000Hz, which is more than adequate.

Load Cells

To record the flow of propellant, the weight of the tanks will be recorded during firing. Load cells are used to record weight. For the purposes of the engine tests, at least two load cells are required: one for the fuel tank, and one for the oxidizer. One load cell has been provided by an Embry-Riddle student organization. This will be used for the fuel tank, as only one tank is required for the fuel and the load cell can easily support this. The oxidizer load cell will have to consist of both a load cell and a stand, so that all three tanks can be measured simultaneously. This stand, minus the load cells, can be seen in Figure 75.



Figure 75: Oxidizer Stand with Two Tanks

As a cost saving measure, load cells were created by adapting an inexpensive digital bathroom scale. Upon dismantling the digital scale, it was discovered that is consisted of 4 feet, each of which made up one half of a Wheatstone bridge. A Wheatstone bridge is a basic electrical configuration frequently used in load cells, and other transducers. The circuit board of the scale could not easily be adapted, as it was hard-wired for home use and not scientific use. Because of this constraint, a new electrical board was designed, as shown in Figure 76.

Figure 76: Half of the Symmetrical Load Cell Electrical Circuit



The signal is then sent to an instrumentation amplifier. The amplification is controlled using a potentiometer (shown as Rg), and the output is sent to the DAQ. The bridge is powered by a 9V battery. The voltage is actually dropped down by a resistor. The amplifier is powered by two 9V batteries. Because this load cell setup

required for the tank stand. The voltages of these two load cells could then be added together to make a single output, but it is easier to plug both outputs into the DAQ and sum them programmatically.

Software

The software has been written for the GUI and test fire control. Some transducer constants are still required to be inserted into the software. The front panel GUI can be seen in Figure 77. It is designed to be easy to use and compensate for user error. The software is easy to use because of simplified controls and outputs are few, and easily understood. It compensates for error as all recorded data is automatically saved in a unique, easy to identify file. This way, in the excitement of firing, information is not lost.



Figure 77: Testing Software Front Panel GUI

Figure 78 represents a sample of the software code. The code primarily consists of one large hile loop, which collects new data with each iteration. When the command to record data is entered, the iteration rate (and thus sampling rate) is sped up, and all data is saved to an array. At the end of firing, all data in the array is automatically saved into a unique text file. The text file name includes the firing time, making it easy to identify. As can be seen in Figure 78, each type of data has

These lanes keep the code clean and easy to follow.



Figure 78: LabVIEW Code.



6.5. Test Preparation Procedures

Feedline System / Section Pressure Test

Note that this test concerns feed lines and joints only. It does not apply to filters, valves, or any other components of the full feed line system.

Statement of Purpose: The primary objective of the feed line system/section pressure test is to determine whether an individual feed line section or a set of feed lines connected to form a coherent system carry the tolerance to withstand the calculated pressures and flow rates within the feed line system without failure. Failure for such a test is to be defined as follows: leaking, bending, buckling, or the appearance of other inelastic deformations. If any such failure occurs during the following testing procedure, the section or system is to be considered unfit for use. If serious failure, detailed in SSPP Sub-System Hazards Risk Index Matrix occurs despite expectations, such as a line suffering catastrophic failure (example: pressure burst), the test shall immediately be aborted, and safety personnel shall take action according to the proper procedure in the related SSPP specification.

Inventory Checklist: (1) Pressurized water reservoir. Must have a secure release valve as well as water

at the calculated pressure for the feed line section to be tested. Example: sections located before the primary pressure regulator in the FEED SYSTEM DIAGRAM must be capable of tolerating pressure to 4000 psi, while all subsequent lines (located after the primary pressure regulator) must be capable of tolerating 1500 psi without failure.

(1) Sealant cap. (N) Feed line sections that comprise system under test. Exact number of feed line

sections will depend upon system configuration. (N) Fittings measured to properly connect sections of system under test to one

another as well as sealant cap and release valve of pressurized water reservoir.

Test Procedure: 1. Obtain pressurized water reservoir.

a. 99. 2. Obtain feed line(s). 3. Obtain proper fittings for sealant cap and pressurized water reservoir. 4. If multiple feed line sections are to be tested, connect feed line fittings to

produce desired configuration of the system under test: a. Visually inspect lines for defects. b. Line interior of seals with Teflon tape. c. Torque until established specifications have been met. Do not cease

until a tight seal is created. 5. Cap feed line system at one end of system:

a. Line interior of seal with Teflon tape.



b. Torque until established specifications have been met. Do not cease until a tight seal is created.

6. Connect pressurized fitting at other end of system to pressurized water reservoir:

a. Line interior of seal with Teflon tape. b. Torque until established specifications have been met. Do not cease

until a tight seal is created. 7. Pressurize system:

a. Open exit valve of pressurized water reservoir to allow entry of pressurized water into system under test.

8. Check for failure: a. For feed line section test only, leaks at the joints of the system are not

relevant and may be ignored in the observation. If the entire system is under test (feed lines as well as joints), the aforementioned does not apply.

9. Document failure, if any pertinent case exists. 10. Depressurize system:

a. Close exit valve of pressurized water reservoir to stop entry of pressurized water into system under test.

b. Disconnect system end cap to depressurize system. Take caution: water under pressure can carry respectable force.

11. Conclude test: a. Disconnect pressurized water fitting from reservoir. b. Disconnect system configuration if necessary. If no pressing reason

exists to deconstruct present configuration, allow system to remain as such in case further testing with present configuration is required.



6.6. Fuel Tank Air Evacuation

Statement of Purpose: The primary objective of this procedure is to evacuate a fuel tank, which is slated to be filled with propane, of 99.9% of the air initially present within the pressure vessel. In order to accomplish this, the fuel tank shall be repeatedly pressurized and vented with nitrogen, as described in detail in the following steps. Once purged of air, the tank is then kept at a significant enough positive pressure (approximately three times ambient, or 3 atm.) to prevent any amount of air from re-entering the tank and posing an ignition hazard, even in the event that a slight leak should develop. Positive pressure in case of leakage shall allow safety personnel ample time to take action according to SSPP

Fuel Tank Specifications: Type: DOT-3AL1800 EN0001300

M4002 06C07 N020 TC-3ALM124 CATALINA Manufacturer: Nitrogen Oxide Systems, 1801 Russellville Rd. Bowling Green, KY 42101 Full/Empty Weight: 45 lb 2 oz / 25 lb 2 oz Rating: 3000psi Current Location: LB 184/187 - test stand lab Contents (planned): Propane, compressed Contents (present): Air, atmospheric pressure Fitting 1: CGA unknown male; ma Fitting 2:

Nitrogen (Pressurant) Tank Specifications: Type: DOT-E10869-45006

HO-264057 Manufacturer: Holox Atlanta Full/Empty Weight: Unknown Rating: Unknown Current Location: LB 189 - test stand lab Contents (planned): Nitrogen, compressed Contents (present): Nitrogen, compressed, UN 1066 Fitting: CGA 580 (Compressed Gas Association nitrogen fitting) female;

Fuel Tank Air Evacuation Procedure:

1. Obtain fuel tank from designated storage area: a. Visually inspect for defects.

2. Obtain nitrogen tank from designated storage area: a. Visually inspect for defects.

3. Obtain pressure transducer with propane fitting.



a. Inspect if reasonably possible. b. Connect pressure transducer to fuel tank propane fitting. c. Line interior of seals with Teflon tape. d. Torque until established specifications have been met. Do not cease

until a tight seal is created. 4. Open fuel tank propane valve.

a. Check pressure transducer. 5. Connect transfer line to nitrogen fittings of fuel and nitrogen tanks:

a. Visually inspect transfer line for defects. b. Line interior of seals with Teflon tape. c. Torque until established specifications have been met. Do not cease

until a tight seal is created. d. Set nitrogen regulator to just exceed 100 atmospheres (1500 psi.) of

absolute pressure. 6. Open fuel tank nitrogen valve. 7. Open nitrogen tank nitrogen valve. 8. Pressurize fuel tank from nitrogen tank:

a. Pressurize until pressure transducer reads approximately 100 atmospheres (1500 psi.) of absolute pressure.

9. Close fuel tank nitrogen valve. 10. Partially vent fuel tank:

a. Open fuel tank bleed valve. b. Allow fuel tank to vent nitrogen/air mixture until pressure transducer

reads approximately 3 atmospheres (44.1 psi) of absolute pressure within fuel tank.

c. Caution: nitrogen gas will escape into the air. Perform venting only in a well-ventilated area.

d. Warning: avoid inhaling nitrogen. e. Close fuel tank bleed valve.

11. Repeat steps 6-10 four more times. This should almost completely evacuate fuel tank of air. If unsure, repeat until comfortable that 99.9% of the initial oxygen has been evacuated.

12. Close nitrogen tank nitrogen valve. 13. Disconnect transfer line from fuel tank:

a. Caution: transfer line is under pressure. Disconnect very slowly at first, until nitrogen gas begins to vent from line, then cease disconnecting until pressure has equalized i.e. nitrogen gas stops or very nearly stops escaping. In addition, ensure transfer line is secured.

b. Caution: nitrogen gas will still be present within transfer line. Disconnect only in a well-ventilated area.

c. Warning: avoid inhaling nitrogen. 14. Disconnect transfer line from nitrogen tank. 15. Return nitrogen tank to designated storage area.



6.7. Fuel Tank Fill (Propane)

Statement of Purpose: The primary objective of this procedure is to fill the fuel tank with liquid propane from a standard propane supply tank. In order to accomplish this, the tanks shall be connected to one another via transfer line, followed by the transfer of propane from supply tank to fuel tank, as detailed in the procedure below. If at any point a leak is discovered, whether by the presence of liquid propane or another indicator, the transfer shall immediately be aborted and safety personnel are to take action according to SSPP

Fuel Tank Specifications: Type: DOT-3AL1800 EN0001300

M4002 06C07 N020 TC-3ALM124 CATALINA Manufacturer: Nitrogen Oxide Systems, 1801 Russellville Rd. Bowling Green, KY 42101 Full/Empty weight: 45 lb 2 oz / 25 lb 2 oz Rating: 3000psi Current Location: LB 189 - test stand lab Contents (planned): Propane, compressed Contents (present): Air, atmospheric pressure Fitting 1: CGA unknown Fitting 2:

Supply Tank Specifications: Can vary, but must be standard propane supply tank compliant with safety standards outlined in SSPP

Fuel Tank Filling Procedure - Propane: 1. Obtain fuel tank from designated storage area:

a. Visually inspect for defects. b. Verify that fuel tank has been evacuated of air. If not, perform Fuel

Tank Air Evacuation. 2. Obtain propane supply tank from designated storage area:

a. Visually inspect for defects. 3. Prepare transfer line:

a. Visually inspect transfer line for defects. b. Line interior of seals with Teflon tape.

4. Connect transfer line to propane fitting of supply tank: a. Torque until established specifications have been met. Do not cease

until a tight seal is created. 5. Open supply tank propane valve. 6. Connect transfer line to propane fitting of fuel tank:



a. Wait a few moments before connecting to allow propane under pressure to fill transfer line and evacuate air remaining within.

7. Open fuel tank propane valve. 8. Fill fuel tank from supply tank:

a. Invert supply tank to allow liquid propane to drain into fuel tank by means of gravity feed.

b. Open fuel tank bleed valve. c. Continue to fill until bleed valve sounds, indicating that fuel tank is

nearing capacity. Bleed valve will sound with a loud hissing noise that is difficult to miss. Nonetheless, attention should be on bleed valve at all times during this portion of the procedure.

d. Upon bleed valve sounding, close fuel tank bleed valve. 9. Close fuel tank propane valve. 10. Drain transfer line:

a. Right supply tank. b. Raise fuel tank to allow liquid propane in transfer line to drain back

into supply tank 11. Close supply tank valve. 12. Disconnect transfer line from fuel tank:

a. Caution: transfer line is under pressure. Disconnect very slowly at first, until propane begins to vent from line, then cease disconnecting until pressure has equalized i.e. propane venting stops or very nearly stops escaping. In addition, ensure transfer line is secured.

b. Caution: gaseous and liquid propane may still be present within transfer line. Disconnect only in a well-ventilated area and inspect surroundings for potential ignition sources. If any sources of ignition are present, do not disconnect transfer line at this location.

c. Warning: avoid any form of contact with any form of propane. 13. Disconnect transfer line from supply tank. 14. Return supply tank to designated storage area.



6.8. Fuel Tank Fill (Water)

Statement of Purpose: The primary objective of this procedure is to fill the fuel tank with liquid water from a pressurized water feed, which may be a reservoir, line pump, or any other

ressurized water

water feed via a transfer line, followed by the filling of liquid water into the fuel tank, as detailed in the procedure below. If at any point a leak is discovered, the transfer shall immediately be aborted and safety personnel are to take action

Fuel Tank Specifications: Type: DOT-3AL1800 EN0001300 M4002 06C07 N020 TC-3ALM124 CATALINA Manufacturer: Nitrogen Oxide Systems, 1801 Russellville Rd. Bowling Green, KY 42101 Full/Empty Weight: 45 lb 2 oz / 25 lb 2 oz Rating: 3000psi Current Location: LB 184/187 - test stand lab Contents (planned): Water Contents (present): Air, atmospheric pressure Fitting 1: Fitting 2:

Fuel Tank Filling Procedure - Water: 1. Obtain fuel tank from designated storage area:

a. Visually inspect for defects. 2. Obtain pressurized water feed:

a. Pump with line, hose from wall, etc. from now on referred to as

3. Open fuel tank bleed and propane valves: a. Allow pressure inside and outside fuel tank to equalize.

4. Connect pressurized water feed line to propane fitting of fuel tank. 5. Fill fuel tank with water:

b. Start pressurized water flow. c. Fill until fuel tank bleed valve begins to discharge water. a. Close fuel tank bleed valve.

6. Close fuel tank propane valve. 7. Stop pressurized water flow. 8. Disconnect pressurized water feed line from fuel tank. 9. Drain pressurized water feed line.



6.9. Assembly Procedure

At this point, the torque values for connecting each component and the exact fitting sizes of each component are not known. These values will be researched more in depth and included before this procedure is worked. Each connection must include Teflon tape around the threads to prevent against leaks.

To begin the assembly of the feed line system; make sure the following instrumentation is present:

Torque wrench

Crows foot adapters

Thread sealant

Sockets

Wrenches

Bolts

C-clamps

Sand Bags

Hammer

Drill and required bits

Pressurant System:

1. Verify all valves to nitrogen tank are closed. 2. Place nitrogen tank in proper compartment. 3. Secure nitrogen tank sandbags at the base. 4. Connect regulator to nitrogen tank outlet:

a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

5. Connect solenoid valve to outlet of regulator: a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

6. Connect pressure transducer to inlet of fuel tank: a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

7. a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

Fuel System:

1. Verify all valves to fuel tank are closed. 2. Place fuel tank in proper compartment ( 3. Figure 79). 4. Place on top of load cell and secure with sandbags at base. 5. Connect solenoid valve to outlet of fuel tank:

a. Line interior of seals with Teflon tape.



b. Torque until established specifications have been met. 6. Connect a tee fitting on the outlet of solenoid valve:


7. Connect pressure transducer to top of tee fitting: a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.


9. Run hose from fuel tank location to test stand location securing every two feet with tie straps.

10. Connect a check valve to the other end of the ten to thirteen foot hose with the flow arrow pointing downstream (away from the fuel tank).


11. Connect a solenoid valve to the check valve: a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

12. Connect a tee fitting on the outlet of solenoid valve: a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

13. Connect pressure transducer to top of tee fitting: a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

14. adapter:


15. et to an (4 each) inlet on the cooling jacket:


Oxidizer System:

1. Verify all valves for each oxidizer tank are closed. 2. Place oxidizer tanks (three each) in proper compartment ( 3. Figure 79). 4. Place on load cells and secure oxidizer tanks (three each) with sandbags. 5.


6. Connect other end of (three each) hose to (three each) inlet holes in manifold; a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

7. Connect a tee fitting on the outlet of manifold;




8. Connect pressure transducer to top of tee fitting; a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.


10. Run hose from oxidizer tank location to test stand location securing every two feet with tie straps.

11. Connect check valve to the other end of the ten to thirteen foot hose with the flow arrow pointing downstream (away from the oxidizer tanks).


12. Connect a tee fitting on the outlet of the check valve; a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

13. Connect pressure transducer to top of tee fitting; a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

14. Connect solenoid valve to outlet of tee fitting; a. Line interior of seals with Teflon tape. b. Torque until established specifications have been met.

15. read into the forward dome.



Figure 79: Top View of Housing Compartments

6.10. Test Area Overview

Validation Test Area Overview

Qualification test for the Lunar Lander vehicle will take place inside a steel container measuring thirty feet long by eight feet wide by four feet deep. The container has three areas that are isolated from each other. These sections contain sand bags used to insulate and shield the container (sand bags are on the inside of the container in all sections). In the case of a catastrophic failure, the sand bags will absorb kinetic energy from debris. This prevents debris from ricocheting off container walls and provides additional protection to the surrounding environment. Another benefit of this design is the immediate containment of any chemical or metal fires that may occur (refer to fire containment section on Page 107). The sand and dust will become a suppressant, absorbing fuel and oxidizer to deprive any fire of necessary fuel/oxidizer for further combustion.

Each area is sectioned off using a steel blast plate approximately 0.25 inches thick which is held in place by C-clamps, sand bags that insulate the wall, and a second blast plate that is positioned perpendicular to the first. This will isolate the engine from both the fuel and oxidizer tanks as well as isolating the tanks from each other. The steel structure of the testing container will reinforce the sandbags and act as a secondary shield.

The area also has a tent above it to protect the fuel and oxidizer tanks from direct sun light which protects the tanks from significant pressure changes due to solar heating. In addition, the tanks are fastened to the testing container utilizing a setup similar to what one would see for a fire extinguisher attached to a wall in a building. The fuel and oxidizer feed system is insulated in a series of cinder blocks covered with sand bags until it intersects the primary blast plate that separates the tanks from the test engine (Figure 80).

STEEL BLAST



Figure 80: Housing Illustration, Top View

CINDER BLOCKS WHICH ARE COVERED WITH SAND BAGS

FUEL AND OXIDIZER LINES



The test container is sealed around its base so fluid that can potentially exit the test system does not have a transit path to exit the test container; there is a small hole to prevent rain water from building up but this will be sealed prior to the test. In addition, there is a thirty-that can open but new containment procedures will be developed if use of this variation is appropriate). As a secondary containment measure to prevent fluids from contaminating the environment, the testing container is placed on a polypropylene hold excess liquid if necessary.



Fire Containment

Automatic water sprinkler fire suppression systems are not viable for the test area, due to its distance from buildings and the fact that it would be ineffective against metal fires. Current plans employ sand as the primary fire control tool. Sand sprays in cases of sand bag rupture which occurs if either debris or fire compromises the structural integrity of the sand bag (refer to Figure 82). Specifically, the sand bags are made of a plastic material which melts under high temperatures produced by fire. When the sand bags at the base melt, the sand will shift because of force applied by the other sand bags above and the remaining sand bags will collapse as shown by the arrows in Figure 82. These sand bags will cover the area that contains the fire. To ensure stability of the sand bags in the case of no fire out breaks, the sand bags will be held in place with thin plastic cords (not illustrated in Figure 82). As a secondary fire control measure, fire department and emergency medical personnel will be on site with additional equipment and for observation. The test site will be located in an area with an abundance of sand, serving as a large fire buffer zone.

Figure 82: Testing Container Illustration, Front View

Site selection also depends on the ability for a heavy vehicle to drive on the sand to deliver the test stand.

AREA OF FIRE

BASE SAND

BAGS

BASE SAND

BAGS

UNSTABLE

RESULT



Ventilation

Test site and test container will be in an outdoor location that meets proper ventilation standards set by OSHA. Elevated terrain on three sides of the container prevents exhaust of toxic gases from entering populated areas and controls its direction. The direction with no elevated terrain must not contain populated areas, since it is very likely any gases will flow in that direction due to air currents.

Decontamination of Personnel

Personnel conducting the test will be in a containment area that is shielded from the test site and is located a safe distance from test container. Should contamination occur, adequate water and medical supplies as well as fire and medical personnel will be on site to facilitate decontamination of personnel.

Material Compatibility

All materials utilized in test area will be reviewed for proper storage conditions. The test container has separate storage areas for the fuel and oxidizer due to their potential reactive nature. Dissipation of static charge is accomplished by grounding testing system to testing container before flow starts.

Spill Containment

The test container is the primary containment entity. The test container is located above apolypropylene tarp which provides secondary containment.

Public Access

Access to primary main engine test area is easily regulated due to its isolated location.

Personnel Safety

Personnel shall be located at an undetermined distance suitable for all safety risks. To mitigate risk, a safety wall or shield will be in place.



1. Entrance to test site number one which is accessible from Clyde Morris Boulevard. This entrance is least likely to have a heavy vehicle become immobilized due to loosely packed sand. In the case of an incident of any kind, this area is easily accessible from a major road way that will minimize the time between an accident and the arrival of emergency personnel. 2. The actual location of the test container is located at this point. There is an abundance of sand for filling bags which will insulate the inner portion of the testing container. In addition, the surrounding terrain is elevated with respect to this location, providing more protection for surrounding areas. In case of a fire, which would most likely be either a fuel or metal fire, the abundance of sand provides a wide fire buffer not available at other locations. Polypropylene will be used to cover the ground and surrounding area using a technique that will prevent and accidental spills from the testing container from contaminating the environment. 3. This area has been cleared of trees and other vegetation.



1. Entrance accessible from building 1 (Information Technology) parking lot. Testing location is more difficult to reach and heavy vehicles have an elevated risk of becoming immobilized due to loosely packed sand. 2. Entrance accessible from building 2 (ROTC) parking lot. Testing location is more difficult to reach and heavy vehicles have an elevated risk of becoming immobilized due to loosely packed sand. 3. Location of testing container. There is an abundance of sand for filling sand bags that will insulate the inner portion of the testing container. In addition, the surrounding terrain is elevated with respect to this location providing more protection for surrounding areas. In case of a fire, which would most likely be either a fuel fire or metal fire, the abundance of sand provides a safe guard that would not be possible at other locations. Polypropylene will be used to cover the ground and surrounding area using a technique that will prevent and accidental spills from the testing container from contaminating the environment.



Figure 86: Alternate test site easily accessible from Richard Petty Blvd.

6.11. Feed Line System Testing Overview



Figure 87



Figure 88

Figure 88

Figure 88: Oxidizer System

Figure 89Figure 89



Figure 89: Pressurant System



Figure 90



Figure 90: Fuel System

6.12. Instrumentation

The instrumentation below will be used to measure the different values needed for an accurate and thorough analysis:

1. Thermometers 2. Pressure transmitters 3. Power Generator 4. Load Cells 5. Data acquisition (DAQ)



Procedures 1. Establish the testing facility. 2. Secure the testing area, assuring that only required personnel are present. 3. Assemble the feed line system by following the Assembly Procedures Assembly

Procedure. 4. Connect and zero data acquisition software, step motors, and the emergency stop

system. a. Attach the power cord for the DAQ to both the DAQ and an electrical

outlet. b. Connect the DAQ USB cord to the DAQ and the computer. c. There are four color coded wires which connect from the trolley to the

DAQ. These transmit the pressure transmitters and the flow meter data. Attach these cords.

d. Attach the COM wires to the step motors and the computer. These are labeled.

e. Plug in the Power Supply box and set it to 12 V DC. f. Attach the Power Supply ends to the COM cable. g. The LabVIEW software should read inputs from the various sensors.

Verify that these readings are reasonable and that the test can proceed. 5. Sub-Test 1: Control Verification:

1.a.Check the emergency cutoff system (the solenoid valves.) b.a. Toggle then reset the pressurant switch. b.b. Toggle then reset the fuel switch. b.c. Toggle then reset the oxidizer switch.

6. Sub-Test 2: Leak Test for Pressurant System:

Note: The following procedures shall be fulfilled while referencing the safety checklist.

a. Fill Nitrogen tank by following filling procedures (Page Error!

Bookmark not defined..) b. Set regulator to approximately 1000 psi. c. Connect Nitrogen to corresponding solenoid valve. Refer to Assembly

Procedures Page 100. d. Connect Nitrogen system with appointed pressure gage. Refer to

Assembly Procedures Page 100. e. Connect Nitrogen system, following the pressure gage to the empty fuel

tank. Refer to Assembly procedures Page 100. f. Open ball valve. g. Read value, and compare to regulator setting. h. Reiterate steps a-g until connections and lines are leak free.

7. Sub-Test 3: Flow and Pressure a. Assemble remaining parts of the feed line system by following the

Assembly Procedures, to comply with the configuration specified. b. Fill fuel tank by following filling procedures Page 97.



c. Open all manual tank valves. d. All personnel must go to feedback station, at least 100 yards from test

facility. e. Open solenoid for pressurant system using toggle switch. f. Take pressure reading from fuel pressure transmitter. g. Take pressure reading from oxidizer pressure transmitter. h. Compare readings with predicted values, and make adjustments.

8. Clean-up fluid contained by solenoid valves by following safety 9. Command LabVIEW to stop logging measurements. 10. Purge system with nitrogen gas. 11. Clean area equipment stowed and data saved.

Results/Analysis

A successful test will produce the information needed to ensure that the system is functional and efficient. A functional system will ignite when given the signal, with predicted amounts of fluid flow through the feed lines. As for measurable data, we are interested in analyzing the pressure, temperature, and flow rates of the system. Through graphical analysis, a comparison will be made between measurements obtained from the test and the theoretical values desired. We expect the operation of the feed line system to match theoretical calculation within 15% difference

The plots to be analyzed are as follows:

Mass flow rate vs. time

Fuel tank 3 Oxidizer tanks

Pressure vs. Time

Fuel line before cooling jacket manifold Oxidizer line before pressure chamber

6.13. Igniter Testing Procedure

The equipment needed for testing are as follows i. (1) Variable Power source ii. (3) Glow plug iii. (3) Locking cap w/ cord iv. (4) Set of Banana Clips v. Socket Set

Installing Glow Plug:

1. Using a Rossi R8 glow plug, with the threaded side facing down, place the copper washer over the threaded end and slide the washer upward until it can no longer move upward anymore.

2. Holding the copper washer in place lightly with your finger, place the threaded end of the Rossi R8 glow plug into the top of the threaded hole offset



from the center of the oxidizer dome and turn clockwise until glow plug is snug using your hand.

3. Using a 5/16 socket and socket wrench, fit the socket over the top of the glow plug and proceed to tighten until 20N of force is used.

4. At this point the glow plug should be set into its proper slot with the copper washer, provided with the glow plug, creating a seal between the face plate and the glow plug.

5. Glow plug installation complete.

Connecting Power Source to Glow Plug(s): 1. Assuring the variable power source is unplugged from the outlet and the

variable power source is off; connect the glow plug locking head cord(s) up to the DC plugs on the power source while ensuring that the three locking head caps are connected up in parallel.

2. Next, place the glow plug locking cap over the glow plug 3. Press down with the locking cap and turn counter clockwise and then release.

This should effectively lock the igniter cap on top of the glow plug ensuring a complete electrical connection as well as prevent the cap from falling off the end of the glow plug.

4. The power source should now be connected to the glow plug via the glow plug locking head cord.

Turning on the Power Source / Glow Plug:

1. With the voltage and current on the variable power source turned down to zero, plug in the variable power source to the outlet and turn on.

2. Turn up the Voltage and Current slowly, while trying not to produce too high of a voltage that could blow out the filament, until you reach about 1.5V and 3A for a single glow plug (or 1.5V and 9A for three glow plugs in parallel). This should make the glow plug filament(s) produce an orange glow.

3. While not changing the current, turn up the voltage slowly until just under 2.1V or until the glow plug filament glows a bright white. This will be the hottest temperature until the filament will break and will severely reduce the life of the glow plug filament. Using this voltage of 2.1V, will be the voltage used across the filament to ensure the greatest possibility for combustion, which is also the hottest temperature the glow plug will reach before being destroyed.

4. The glow plug should now be glowing white and at its hottest temperature.

6.14. Ignition Test Results

The testing for the glow plug ignition system took place on April 14, 2008 outside the rocket lab utilizing a propane grill and three glow plugs connected up in parallel. There were two types of configurations that were tested, one being three glow plugs spaced directly next to each other placed directly on top of the propane



outlet where the gas would come out to create a flame. The other configuration that was tested was using the three glow plugs evenly spaced around the burner to possibility create a more uniform heat and ignition. The ignition tests were done in the following sequence; propane would be turned on first to prevent any combustion while there was a user close by. Then the glow plugs would be heated up while waiting for combustion. If no combustion occurred after ten seconds, the glow plugs would be turned off and then the propane would be turned off. Although all ignition tests were unsuccessful, the fact that it did not replicate the actual firing does not mean the ignition will not occur. There were many factors in this test that are not similar to the actual firing. For instance, the pressure of the propane that was exiting the burner was much less than the propane that will be injected into the combustion chamber. Also, the pressure build-up that will occur in the combustion chamber was not present. The propane used in the test vented directly into atmospheric pressure.



6.15. Engine Firing Test

The purpose for the engine firing test is to ensure that the engine successfully ignites and produces thrust. The engine firing test will also evaluate the overall performance of the engine at different oxidizer to fuel ratios. The data collected from the test will be compared to theoretical models and used to determine if the design is effective as a rocket engine and if the design can be improved.

The set up of the engine and the feed line system are the same as that of the feed line system test, shown in Error! Reference source not found.. The testing facility

Test Area Overview 103.

Engine Firing Test Instrumentation

The instrumentation below will be used to measure and record different values which will be used for an accurate and thorough analysis: 1. Test Stand 2. Tank Scales 3. Pressure gauges/transducers 4. Thermometer (temperature sensor) 5. Flashlight 6. Hand mirror 7. Digital Camera

The following values will be measured during this test:

Thrust Using strain gauges attached to the test stand, the total engine thrust can be calculated and recorded. The deformation of the steel test stand can be used to find the total force exerted from the engine.

Temperature Using a thermometer located on the outside of the engine wall, the temperature of the engine wall will be recorded. From this value the temperature of the fuel flowing through the cooling system can be calculated.

Pressure in the feed line Using pressure transducer in the same configuration as the feed line system (Error! Reference source not found.), the data acquisition system will measure the pressures for the fuel and oxidizer in the feed line system.

Mass flow rate The are two options for calculating the mass flow rate. The first is to place the tanks on a digital scale system used in the previous test. This will record the tank weight as the engine is fired.

Option two is to weigh the propane tank before and after the engine is fired. Each firing will be strictly timed. From these two values the mass flow rate of the propane will be calculated.



Test Fire procedure

1. Secure the test

2. tank:

i. Ensure propane and oxidizer tank solenoid valves are completely closed by deactivating the Emergency Kill button.

ii. Detach the propane tank from the feed line system. iii. Fill propane tank. (See Fuel Tank Filling Procedure steps 2 through 11.) iv. Reattach propane tank to the feed line system in same configuration as the

feed line test. v. Ensure that the solenoid valve is closed before opening tank propane

valve. 3. Feed line system from previous test should be in testing configuration, already

assembled (see assembly procedure, Page 100.) 4. Connect and zero data acquisition software. 5. Run software to assure that all sensors are being read. 6. If required, record weight of fuel and oxidizer tanks. 7. Ensure control and solenoid valves are closed. 8. Manually set fuel and oxidizer pressure regulators to the specific pressures

corresponding to an OF ratio of 3. 9. Evacuate the testing chamber. The area should again be verified that it is clear of

be looked at. 10. Set LabVIEW program to record all data. 11. Activate the solenoid valves using the Emergency Kill buttons to allow the

required flow through the lines, always looking for possible leaks. 12. Ignite engine. 13. Allow engine to run for 3 seconds. 14. Stop flow of propellant by closing the control valve (pressing the Emergency Kill

Buttons.) This stops the combustion in the engine. 15. If required, record the weights of fuel and oxidizer tanks. 16. Save the data collected by the LabVIEW Program. 17. Let engine cool. (15 minutes) 18. Thoroughly check feed lines, engine, and testing supports (stand and harness) for

damage. Record observations in a detailed report with photographs/drawings.

rst aid should be present in the vicinity for emergency action. Use the Observation Testing Document.

i. When checking lines, look for structural damage and component damage. Visually inspect each component of the feed line system.

ii. Visually inspecting the engine is most important.



i. First check the outside of the engine wall for burn spots and any deformation. Ensure that the temperature sensor at the top of the engine is still in place.

ii. Check the inside of the engine chamber and nozzle for deformation and burn spots. A flash light and hand mirror will most likely be needed to make a thorough check.

iii. Visually inspect the engine test stand and harness. i. Check where the engine is mounted to the test stand to make sure it

is secure enough to with stand another test. ii. Ensure all strain gages are still attached to the stand. iii. Check for any signs of inconsistent loading on the stand. We want

to ensure that test stand is structurally capable of test conditions. iv. Visually inspect the testing container and surrounding testing facility for

damage. 19. If no damage is detected in step 15, repeat steps 3 through 16 for O/F ratios 4, 5,

6, 7, and 8. 20. If damage is reported, stop testing, ensure that accurate and thorough observations

are documented. 21. Clean up testing site and engine system.

Test Firing Results/Analysis

Most of the analysis of this test is observational. This test is to ensure the functionality of the engine system as a whole (feed line system, cooling system, and engine). By thoroughly checking every component in the system, the effects of the O/F ratio at each firing will be concluded. A maximum O/F ratio for the engine should be found through this test. It is important to find a safe O/F ratio for the engine, where it can be properly and safely cooled and fired. The measured results acquired from this test will be compared to theoretical values. Final changes regarding the best operational O/F ratio will be made based on the test results.

The plots to be analyzed are as follows:

Thrust vs. O/F ratio

Temperature of engine wall vs. O/F ratio Additional Results:

Pressure readings from feed lines at each O/F ratio

Mass flow rate measurements for the fuel and oxidizer in the feed line system



7. Project Economics

An initial budget of $6000 was acquired for the Lunar Lander Challenge project. One half, or $3000, was allocated toward the puchase, manufacturing, and testing of the propulsion system. The purchase of fuel, various miscellaneous parts such as bolts and screws, raw materials for test engine construction, and the feed system are to be purchased with these funds. By current estimates, building the test engine will cost $1,878.63; this total cost includes all raw materials, some purchased items for the test fire, and the addition of nuts, bolts, washers, O-ring paste, and brazing paste. However, this estimate does not include the cost of the feed system. The overall estimate and the actual cost are close, but the actual cost is slightly higher by less than $20 due to some items, such as the strain gages, being above estimated cost. Other items, such as the nuts and bolts, were well below estimated cost. Note that the difference between the two budgets can also be attributed to the estimated cost of propane. This will significantly increase the actual cost, but it has been compensated for. Table 20: Total Cost EstimateTable 20 is a representation of the total cost estimate for the fabrication and testing of a single engine. Many of the materials and items have already been purchased in the following table.



Table 20: Total Cost Estimate Part Name Part Description Cost Item Quantity Total Price

Rocket Test Stand Located in LB 189 $0.00 1 $0.00

Propellant Propane; 3 gallons; $1.87/gal $5.67 1 $5.67

Oxidizer Nitrous oxide; 5 gal; $36.59/gal $182.95 1 $182.95

Pressurant Gas Nitrogen; 3 gal; $2/gallon $6.00 1 $6.00

Main Rocket Thruster

Machinable Copper; obtained from Acme Industrial Surplus, located in Sanford, FL $568.11 1 $568.11

Main Rocket Ignition System

Several Components $113.50 1 $113.50

Connector Plate 10"x10"x0.25" Stainless Steel Plate Part 6620K21 McMaster-Carr $102.50 1 $102.50

Intermeidate Plate 12"x12"x0.25" Copper Plate Part 8995K11 McMaster-Carr $160.86 1 $160.86

Injector Dome 4"x2" Stainless Steel Barstock $67.22 1 $67.22

Connector Plate 12"x12"x0.25" Steel Plate Speedy Metals $18.34

Strain Gages Testing effects on components $70.00 1 $70.00

Propane Tank Holds propane $396.00 1 $396.00

Testing Dumpster Testbed $35.00 1 $35.00

Combustion Chamber Cooling Jacket

4 1/4"OD x 3 3/4"ID x 6" Steel Tube / SpeedyMetals

$15.72 1 $15.72

Intake Manifold - Propane

4"x2" Round Barstock - Brass $45.63 1 $45.63

Brazing paste $18.00 1 $18.00

Bolts 1/2"x2" Steel Bolts McMaster PN 90201A426 Pkg: 5 $9.30 1 $9.30

Nuts 1/2"x3/4"x5/16" Steel Nut Mcmaster PN 91847A525 Pkg: 25 $10.44 1 $10.44

Washer 1/2"ID x 1 1/4"OD x 1/4" Felt Washer McMaster PN 95571A845 Pkg: 25 $11.25 1 $11.25

O-Ring Goop $15.12 4 $60.48

Total Cost of Acquired Materials: $1,878.63

All materials for manufacturing have been purchased, leaving the testing division with $1,104.38 to use. The initial proposal for the feed system was $997. This proposal severely limited future purchasing options of fuel and miscellaneous emergency costs. The current cost of $300 reflects the cables that have been purchased. Other components for the feed system have not been purchased but are estimated to cost a total of approximately $450. This new proposal will provide a budget cushion. Table 21 represents the current budget and reflects which items have been purchased.



Table 21: Current Budget

Items Purchased

Part Name Cost

Main Rocket Thruster $568.11

Ignition System $113.48

Injector Manifold $45.36

Steel Shower Head $66.60

Steel Nozzle Cooling Jacket $15.58

Steel Combustion Chamber Cooling Jacket $15.72

Connector Plates $285.57

Propellant Tank $396.29

Head Lock $12.75

Brazing Paste $44.20

Strain Gages $141.80

Sand Bags $125.00

Bolts, Nuts, Washer $4.68

O-Ring Goop $60.48

Feed System $300.00

Total Budget $3,000.00

Total Used $2,195.62

Total Budget Available $804.38

The overall cost of the engine has decreased significantly due to calculation of single components rather than three. The estimated total of the engine test fire configuration is under $1,000 which brings the total cost estimate down also. A margin of error has been added to both the feed system and the rocket thruster system. Coincidentally, the budget breaks even. Table 22 represents an estimation of the cost of a single flightweight engine as well as a complete propulsion system configuration.

Table 22: Propulsion System Total Cost System Name Cost

Feed System $1,000.00

Rocket Thruster System $2,000.00

Total cost estimate (1 engine) $3,000.00

Full Total Cost $9,000.00



8. System Safety Program Plan

In order to analyze the hazards system the following guideline will be used in the table which follows for system analysis of specific system parts:

HAZARD RISK INDEX MATRIX

Probability of Occurrence

Hazard Categories

I Catastrophic

2 Critical

3 Marginal 4

Negligible

A Frequent 1A 2A 3A 4A

B Probable 1B 2B 3B 4B

C Occasional 1C 2C 3C 4C

D Remote 1D 2D 3D 4D

E Improbable 1E 2E 3E 4E

Hazard Risk Index Severity Probability Suggested Criteria

1 1A, 1B, 1C, 2A, 2B, 3A Unacceptable

2 1D, 2C, 2D, 3B, 3C Undesirable (Management Decision Required)

3 1E, 2E, 3D, 3E, 4A, 4B Acceptable with Review by Management

4 4C, 4D, 4E Acceptable without Review



Description Category

Environmental,

Safety, and Health

Result Criteria

Catastrophic 1

Could result in death, permanent total disability or irreversible severe environmental damage that violates law or

regulation.

Critical 2

Could result in permanent partial disability, injuries or occupational illness that may result in

hospitalization of more than one person or

reversible environmental damage causing a violation of law or

regulation.

Marginal 3

Could result in injury or occupational illness resulting lost work

days(s), loss exceeding $10K, or environmental damage without violation

of law or regulation where restoration activities can be accomplished

Negligible 4

Could result in injury or illness not resulting in a lost work day, loss less than $10K, or minimal environmental damage not violating law or

regulation.



Description Level Specific

Individual Item

Fleet or

Inventory

Frequent

A

Likely to occur often in the life of an item, with a probability of occurrence greater than 0.1 in that life.

Continuously experienced

Probable B

Will occur several times in the life of an

item, with a probability of occurrence less

than 0.1 but greater than 0.01in that life.

Will occur frequently

Occasional C

Likely to occur some time in the life of an

item, with a probability of occurrence less

than 0.01 but greater than 0.001 in that life

Will occur several times.

Remote D

Unlikely but possible to occur in the life of

an item, with a probability of occurrence less

than 0.001 but greater than 10^-6 in that life.

Unlikely, but can reasonably be

expected to occur.

Improbable E

So unlikely, it can be assumed occurrence

may not be experienced, with a

probability of occurrence less

than 10^6 in that life.

Unlikely to occur, but possible.



Assessing Hazards

Sub-System:

ITEM COMPONENT FUNCTION HAZARD HAZARD HAZARD

NO. DESCRIPTION EFFECTS CATEGORY

& PROB.

DESIGN TO

ELIMINATE

HAZARDS

DESIGN TO

CONTROL

HAZARDS

PROVIDE

WARNING

DEVICES

HAZARD

ACCEPTED

GAS LEAK DUE

TO EXCESS

PRESSURE

BUILD UP

1D6 RELIEF VALVE

REDUCE

PRESSURE IN

TANKS IN CASE

IT EXCEEDS

EQUIPMENT

FAILURE

(MANUFACTURE)

/ UNABLE TO

MAINTAIN

DISCHARGE

RATE

DESIGN TO

CONTROL

HAZARDS

5

PRESSURIZED

NITROGEN

GAS

THE GAS USED

IN ATTITUDE

CONTROL

SYSTEM

VERY COLD IN

LIQUID STATE

3D

MATERIAL

USED FOR

COVERING THE

ATS THRUSTER

FAILURE DUE TO

STRESS OR

STRAIN

CAN CAUSE

FROSTBITE IF

RELEASED

2C

MONITOR

PRESSURE IN

THE TANK

FAILURE

(MANUFACTURE)

/ IMPROPER

INSTALLATION

DIFFICULT TO

DIAGNOSE A

PRESSURE

PROBLEM

4E

3SOLENOID

VALVES

CONTROL THE

FLOW BY

CUTOFF/ON

FAILURE IN THE

EQUIPMENT OR

FAILURE DUE TO

IMPURITY

UNCONTROLLED

THRUST3B

4PRESSURE

TRANSDUCER

2

AL 6061

THRUSTER

CASING

1PRESSURE

REGULATOR

MAINTAIN A

CONSTANT

PRESSURE IN

GAS TANK

FAILURE IN THE

EQUIPMENT

(MANUFACTURE

OR IN FLIGHT)

GAS TANK

EXPLOSION OR

LEAK

1C

CHANGE THE

ORIENTATION

OF GAS

THRUSTERS

COLD GAS-THRUSTER SYSTEM _________________PREPARED BY___________

CONTROLLER POWER____

RECOM.

CONTROL

ANALYSIS TYPE:: SUB-SYSTEM HAZARD ANALYSIS

DESIGN TO

CONTROL

HAZARDS





& PROB.


COOLING SYSTEM _________________PREPARED BY___________


RECOM.

1

COOLANT

(PROPANE)

FLOW

COOL THE

THRUST

MECHNISM

UNEVEN FLOW

TRANSPORT

COOLANTCAN LEAK

CONTROL

VARIED THRUST 2C

PROVIDE

SAFETY

DEVICES

OVERHEATING

CAN CAUSE

ENGINE SHUT

DOWN

2D2 FEEDLINES

3CHECK

VALVES

MAINTAIN

CONSTANT

TEMP.

FAILURE IN ITS

INTERNAL

EQUIPMENT

THRUSTERS

OVERHET1D

PREVENT

COOLANT TO

FLOW IN

WRONG

MANUFACTURE

FAILURE

INCREASE IN

TEMP. CAN

CAUSE

PROPANE TO

BOIL

2D4 TRANSDUCER

PROVIDE

SAFETY

DEVICES

PROVIDE

WARNING

DEVICES

PROVIDE

WARNING

DEVICES



& PROB.

4E4 BLEED VALVE

DESIGN TO

ELIMINATE

HAZARD

PROVIDE

WARNING

DEVICES

DESIGN TO

CONTROL

HAZARD

DRAIN TANKS

FAILURE IN THE

EQUIPMENT

(MANUFACTURE)

/ FUEL IMPURITY

THE SYSTEM

CANNOT BE

DRAINED AND

FILLED

FUEL TANK

EXPLOSION OR

LEAK

1C

CAUSED

WATER AND

IRON

COMPONENTS

STRUCTURAL

STABILITY IS

LOST / FUEL

STORAGE

FUEL LEAK /

GAIN IN

STRESSES

2

CLEAN FUEL

OR OXIDIZER

OFF DUST AND

RUST

FILTERS CAN

OBSTRUCT

FLOW

THRUST

VARIATION

FAILURE IN THE

EQUIPMENT

(MANUFACTURE

OR IN FLIGHT)

FUEL /

OXIDIZER

FILTER

PROVIDE

WARNING

DEVICES

4D

3ON/OFF

VALVES

1PRESSURE

REGULATOR

MAINTAIN A

CONSTANT

PRESSURE IN

TANKS


TANKS _________________PREPARED BY___________


RECOM.

CONTROL

4D





& PROB.

DESIGN TO

ELIMINATE

HAZARD

DESIGN TO

CONTROL

HAZARD

DESIGN TO

CONTROL

HAZARD

PROVIDE

WARNING

DEVICES

1C

4

SHOWER

HEAD

INJECTOR

EVENLY MIX

THE FUEL

WITH OXIDIZER

FAILURE IN THE

EQUIPMENT

(MANUFACTURE)

FEW UNBURNT

FUEL LEFT4D

3OXIDIZER

INTAKE TUBE

INJECT

OXIDIZER INTO

THE ENGINE

HIGH PRESSURE

COULD CAUSE

STRESS OR

STRAIN FAILURE

OXIDIZER AND

OR FUEL LEAK /

CAN CAUSE

FIRE

CASING

COPPER

MATERIAL

USED FOR

COVERING THE

ENGINE

FAILURE DUE TO

STRESS OR

STRAIN

(THERMAL OR

OXIDIZER LEAK

1C

2FUEL INTAKE

MANIFOLD

DISTRIBUTE

COOLANT

EVENLY FROM

NOZZLE UP

IMPURITY CAN

OBSTRUCT

FLOW /

MAUFACTURE

VARIABLE

COOLING FLOW

RATE / ENGINE

OVERHEATING

1D

1


RECOM.

CONTROL


ENGINE _________________PREPARED BY___________



System Hazards Analysis:



& PROB.

ANALYSIS TYPE:: SYSTEM HAZARD ANALYSIS

SYSTEM ELECTRICAL COMPONENT _________________PREPARED BY___________


RECOM.

CONTROL

1POWER

CABLE

TRANSFERS

POWER

BETWEEN

SOURCE AND

PUMP

PUMP

CONTROLLER

POWER

CABLE FAILS -

LACK OF

LOSS OF

PUMPING

CAPABILITY -

WATER

FLOODS

1D

PROVIDE

WARNING

DEVICES

2 BATTERY

SUPPLY

POWER TO

CONTROLLERS

LEAKAGE OF

BATTERY

LOSS OF

MANUVERING

CAPABILITY

2E

PROVIDE

SPECIAL

TRAINING

3 IGNITOR BURN FUEL

FAILS TO

WORK DUE

TO SHORT

CIRCUIT

CONFUSION

WHETHER

THE FUEL HAS

IGNITED OR

3D

PROVIDE

SPECIAL

TRAINING

FAILURE OF

INSTRUMENTS

ON BOARD

4D

PROVIDE

WARNING

DEVICES

4 WIRECURRENT

FLOW

BURN OUT OF

WIRE OR

SHORT

CIRCUIT



& PROB.


FUEL SYSTEM _________________PREPARED BY___________


RECOM.

CONTROL

1 TANKSCONTAIN FUEL

AND OXIDIZER

LEAKAGE OF

TANKS

EXPLOSION

OR BURNING

OF THE FUEL

1C

DESIGN

TO

ELIMINATE

HAZARD

2TESTING

MOUNT

TEST THE

THRUST FROM

THE MOTOR

FAILURE

CAUSED BY

FAULTY

MATERIALS

ENGINE

MOTOR

EXPLOSION

OR FUEL

LEKAGE

1D

DESIGN

TO

ELIMINATE

HAZARD

3 FUEL PUMP

DELIVER FUEL

TO

COMBUSTION

CHAMBER

FAULTY PUMP

CAUSES FUEL

TO LEAK OR

NOT DELIVER

FUEL BURNING 2D

DESIGN

TO

CONTROL

HAZARD





& PROB.


STRUCTURAL COMPONENT _________________PREPARED BY___________


RECOM.

CONTROL

1NOZZLE AND

FUEL TANKSTHRUST

YIELD

FAILURE MIS-

CALCULATED

STRESSES

FUEL LEAK OR

OXIDIZER

LEAK

1B

PROVIDE

SAFETY

DEVICES

2COOLING

SYSTEM

COOL THE

THRUST

MECHNISM

THERMAL

FAILURE AND

OR FAULTY

MATERIAL

OVERHEATING

CAN CAUSE

ENGINE SHUT

DOWN

2D

PROVIDE

SAFETY

DEVICES

3 STRUCTURE

KEEP THE

ENGINES IN

PLACE

FATIGUE

FAILURE -

DUE TO

TESTING

THE

STRUCTURE

FAILS

CAUSING THE

ENGINE TO

FAIL

1D

DESIGN

TO

ELIMINATE

HAZARD

ITEM COMPONENT HAZARD HAZARD HAZARD


& PROB.


ENVIRONMENTAL EFFECTS _________________PREPARED BY___________


RECOM.

1STRONG

WINDS

STABILITY OF

THE

PROPULSION

SYSTEM

POOR

LANDING OF

THE LANDER

OVER HEAT

THE ENGINES /

BATTERY

COOLING

SYSTEM

FAILS TO

COOL

CONTROL

4CPROVIDE SPECIAL

PROCEDURES

3CDESIGN TO

ELIMINATE HAZARD

3 RAIN

CAN CAUSE

SHORT

CIRCUIT

LOSS OF

POWER AND

CONTROL OF

SYSTEM

3BDESIGN TO

CONTROL HAZARD

2 HIGH TEMP.

4DDESIGN TO

CONTROL HAZARD4 RUST

STRUCTURAL

STABILITY IS

LOST / FUEL

STORAGE

FUEL LEAK /

GAIN IN

STRESSES



Failure Mode and Effects Analysis:

This is a risk assessment technique which uses a systematic technique in identifying potential failures in a system or a process. It is a reliability analysis where the focus is on single events or components which will effect the entire system. The worksheet below shows how a component failure can cause a system to fail.



Resolve Hazards

Hazard Reduction procedure. Risk management involves the recognition of risk, risk assessment, developing strategies to manage it, and mitigation of risk using managerial resources. In ideal risk management, a prioritization process is followed whereby the risks with the greatest loss and the greatest probability of occurring are handled first, and risks with lower probability of occurrence and lower loss are handled in descending order. In our case we are going to use a step wise risk reduction procedure by defining each system as below:

Design to Eliminate Hazards

If Not Eliminated

Design to Control Hazards

If Not Controlled

Provide Safety Devices

If Not Provided

Provide Warning Devices

If Not Provided

Provide Special Procedures or Training

If Not Provided

Hazard Acceptance or System Disposal

If any of the above process is possible from top first then the hazard is reduced, if the first process is not possible then the next step is taken in to consideration until there is no choice but to accept the last step that is Hazard Acceptance or System Disposal process.



a) Design to Eliminate Hazards: This method or strategy of reducing/eliminating hazards deals with acquisition of new equipment or facility. It can also be related with a change in design of current equipment or a change in facility without degrading the performance of the system or facility.

b) Design to Control Hazards: In cases where hazards cannot be eliminated

completely and are inherent, the hazards should be controlled through design. In this safety process it is important to include safety features that are fail-safe or have capabilities to handle contingencies through redundancies of critical elements in the system. If the control design increases the likelihood of hazard occurrence due to its complexity, that particular process should be avoided.

c) Provide Safety Devices: Hazards that cannot be eliminated or controlled through

design should make use of appropriate safety devices in order to reduce its hazard risk. Safety devices such as pressure transducers and relief valves are a critical part of operations if the system or subsystem, and should be incorporated in the system to reduce hazards.

d) Provide Warning Devices: Where it is impossible to preclude the existence or

occurrence of the hazard, visual or audible warning devices should be used for timely detection of conditions that precede an occurrence of hazard. These alarms should be designed and tested before installation in order to avoid false alarms.

e) Provide Special Procedures and Training: Where a hazard cannot be eliminated or

reduced using the above defined procedures then special malfunction or emergency procedures should be developed and formally implemented. These specific operations should be standardized and used in test, operational and maintenance activities. (E.g. Personal Protective Equipment Usage)

f) Hazard Acceptance or System Disposal: In situations where hazard cannot be

reduced by any means, then a decision must be documented in order to either accept the hazard or dispose the system causing the hazard.

Many times in order to reduce the hazard a combination of above procedures can be used. The above hazard controls are implemented in the system and sub-system hazards matrix shown above.



Fault Tree Analysis: A fault tree analysis is a technique used to state an undesired state of a system or its component; and the system is then analyzed in view of its environment and operation to figure out all possible means in which that undesired event could occur. A fault tree analysis is not a method of figuring out all the possible ways in which a system could fail but it includes only those failures that will cause a top level event to fail. In the FTA diagram shown below uses gates, these gates show the relationship of events needed for the occurrence of a "higher" event which is the "output" of the gate; the "lower" events are the "inputs" to the gate. The two gates used are AND and OR gates.

AND gate: An AND gate shown below, a top level system will fail if ALL inputs fail.

OR gate: In an OR gate shown below, a top level system will fail if ANY of its inputs fail.



Operational Hazards Analysis

Hazard Location Comments

Fuel and Oxidizer Storage Tanks

Centralized segregated storage areas for chemicals and gases should be provided in the rear of the

facility. A system safety model should be followed, which specifies appropriate Exposure effects and First Aid procedures when handling chemicals.

Nitrogen Storage Tanks

Although nontoxic and inert, nitrogen can act as a simple asphyxiate by displacing needed oxygen in the air. Users/handlers should follow strict safety guidelines for storage and handling, and consult

nitrogen Material Safety Data Sheets

O/F Ratio Test Pressure Tanks The test procedures should be closely followed. The engine cool time will vary according to the O/F ratio

(this should be in mind before the next step).

Static Charge Test Atmosphere

Caution should be taken when connecting the ammeter to the feed line system. The atmosphere should be checked to see if there is presence of

propane or nitrous oxide in the atmosphere before the test is performed.

Fill Tanks Tanks

A predefined sequence or procedure should be carried out in order to fill the tanks. Caution must be

taken when connecting the fuel line or the fuel adaptor as freeze burns are a hazard during the

process and therefore proper apparel must be worn while connecting or disconnecting the line.



Safety Issues with Propane

1. Properties of Propane:

Physical Data:

Boiling Point: -45/14.7 psia° F Vapor Pressure: 188/psia@100°F Specific Gravity: .504/60°F Solubility (H20): <0.1% Evaporation Rate: Gas at normal ambient conditions Freezing Point: -305°F Molecular Weight: 44

Appearance Odor

Colorless Gas Unpleasant odor caused by odorant.

Fire Fighting and Explosion Data:

Flash Point: -156°F Auto ignition: 742°F Lower Explosive Limit (%): 2.3 Upper Explosive Limit (%): 9.5 Extinguishing Media

Water spray, dry chemical, C02, or Halon

Special Fire Fighting Instructions:

Evacuate the area. Stay upwind of vapors. Stop flow of gas. Use water to keep fire exposed containers and piping cool. Use water spray to disperse un-ignited gas. If ignition has occurred and no water is available, tank or piping metal may fail from overheating. Approach containers from sides, not from ends.



2. Storage and Handling:

Store and use cylinders and tanks in well-ventilated areas, away from heat and sources of ignition.

All Liquid Propane Gas (LPG) tanks should be stored in a cage designed for its protection in an area where the temperature of gas cylinders should not exceed 50ºC.

Cylinders stored in an area outside a building must be a minimum distance of 20 feet from flammable gases or combustible material.

No electrical cord is to hang on or around any gas cylinder.

No smoking near storage or use.

Cylinders must be secured at all time to prevent tipping, falling, or rolling

Only personnel trained in the proper transportation and safe use of gas cylinders shall handle cylinders. Students are not to transport gas cylinders in any vehicle or by foot.

Contents of any compressed gas cylinder must be identified and labeled properly.

Gas cylinder shall only be used in an area where adequate ventilation is provided.

All cylinders shall be kept far enough away or shielded while in the work area in order to prevent contact with sparks, flame or radiant heat

All cylinders need a have a valve protection cap unless they are being used in a manifold.

When personnel have finished using a compressed gas cylinder for the day, the cylinder valve shall be closed and the pressure in the regulator and associated equipment released properly.

No person should direct a high pressure gas at another person.

One must use hands to open and close valves; in case there is a difficulty in opening a valve, contact supplier or vendor.

Empty cylinders must be labeled or marked as "empty". Do not refill an empty cylinder.

If the gas cylinder is accidentally dropped or mishandled, contact the vendor immediately to perform an inspection of the cylinder. The cylinder shall be secured to a wall and tagged inoperative until an inspection can be completed.

Only personnel with sufficient physical stature and strength should handle the gas tanks to avoid any potential hazard resulting from the size and weight of the cylinders.

No more than one cylinder shall be handled at a time except on carts designed to transport more than one cylinder.

Before purchasing any Hazardous Material requisitions, the EH&S Office must be added as an on-line approver on requisitions for compressed gases. Only approved Department of Transportation certified trucks with properly certified drivers are to handle cylinders that are full or empty.



3. Exposure Effects and First Aid Matrix

Route of

ExposureEffects First Aid

Inhalation

Exposure may produce rapid

breathing, headache, dizziness, visual

disturbances, muscular weakness,

tremors, narcosis, unconsciousness,

and death, depending on concentration

and duration of exposure.

Immediately move personnel to area

of fresh air. For respiratory distress,

give air, oxygen, or administer CPR

(cardiopulmonary resuscitation), if

necessary. Obtain medical attention if

breathing difficulties continue.

Skin

This material is not expected to be

absorbed through the skin. Non-

irritating; but liquid form of this

material and pressurized gas can cause

freeze burns.

Frozen tissues should be flooded or

soaked with warm water. DO NOT

USE HOT WATER. Cryogenic

burns, which result in blistering or

deeper tissue freezing, should be

promptly seen by a physician.

Eyes

Direct contact with

liquefied/pressurized gas or frost

particles may produce severe and

possibly permanent eye damage from

freeze burns.

Vapors are not expected to present an

eye irritation hazard. If contacted by

liquid, immediately flush the eye(s)

gently with warm water for at least

15 minutes. Seek medical attention if

pain or redness persists.

IngestionLiquid form of this material and the

pressurized gas can cause freeze burns.

Induce vomiting with warm water

(one quart) only if patient is

conscious. Immediately obtain

medical attention.

Health Conditions Aggravated by Exposure

Personnel with pre-existing chronic respiratory diseases should avoid exposure to this material.



4. Spill, Leak and Disposal Procedures

Steps taken during spill, leak or release

Eliminate all potential sources of ignition. Evacuate all non-essential personnel to an area upwind (at least ½ mile in all directions if tanks are involved in fire). Stop source of release with non-sparking tools before putting out any fire. Ventilate enclosed areas to prevent formation of flammable or oxygen-deficient atmospheres. Liquid spills will vaporize forming cold, dense vapor clouds that do not readily disperse. Avoid vapor cloud even with proper respiratory equipment. Contact Health and Safety department as soon as the problem occurs.

Waste Disposal Methods

Defective, empty, or partially used portable containers should be returned to the supplier and appropriate tags. Do not handle disposal of propane tanks without contacting proper personnel.

5. Special Protective Measures Matrix

Special Protective Measures

Category Comments

Ventilation

Local exhaust and general room ventilation may both be essential in work areas to prevent accumulation of explosive

mixtures. If mechanical ventilation is used, electrical equipment must meet National Electrical Code requirements.

Eye Protection

Use chemical-type goggles and face shields when handling liquefied gases. Safety glasses and/or face shields are

recommended when handling high-pressure cylinders and piping systems and whenever vapors are discharged.

Skin Protection

Prevent potential skin contact with cold liquid/vapors. Use insulated, impervious plastic or neoprene-coated canvas

gloves and protective gear (apron, face shield, etc.) to protect hands and other skin areas.

Respiratory Problems

For excessive gas concentrations, use only NIOSH/MSHA-approved, self-contained breathing apparatus.

Work/Hygienic Practices

Emergency eye wash fountains and safety showers for first aid treatment of potential freeze burns should be available in the vicinity of any significant exposure from compressed gas

release. Personnel should not enter areas where the atmosphere is below 19.5% Vol. oxygen without special procedures/equipment. Respirator use should comply with

OSHA 29 CFR 1910.134 or equivalent.



Safety Issues with Nitrous Oxide

1. Properties of Nitrous Oxide:

Physical Data:

Chemical Formula N2O Molecular Weight 44.013 Boiling Point -88.56ºC (-127.4ºF) Melting Point -93.06ºC (-131.5ºF) Vapor Pressure 5,238 kPa @ 21.1ºC (759.7 psia @ 70ºF) Specific Gravity (air = 1) 1.53 Gas Density (21.1ºC (70ºF) @1atm) 1.836 kg/m3 (0.1146 lb/ft3) Liquid Density (saturation pressure at 0ºC) 0.913 kg/l (57.0 lb/ft3) Specific Volume (21.1ºC (70ºF) @ 1atm) 0.5447 m3/kg (8.738 ft3/lb) Critical Temperature 36.4ºC (97.6ºF) Critical Pressure 7,254 kPa (1,052.2 psia)

Appearance Odor

Nitrous oxide is a colorless, non-corrosive, non-toxic, liquefied compressed gas with a faintly sweet odor and taste.

Special Fire Fighting Instructions:

Nitrous oxide is not flammable but is an oxidizer, which means it can support and enhance combustion. Violent decomposition can also occur in case of fire. Cylinders exposed to fire may have their pressure relief devices activate, if present, and the cylinders themselves may fail, especially aluminum cylinders. From a safe distance, cool the cylinders with a water spray. Use an extinguishing medium appropriate for the surrounding fire.

2. Storage and Handling:

Similar instructions as that used for Propane tanks are to be used except for the few listed below specifically for Nitrous Oxide cylinders.

All nitrous oxide cylinders and manifolds shall be at least 20 feet away from or separated by a one-hour rated fire resistant partition from all flammable gases and materials (such as oil, grease, and all petroleum products in general) in the area of use.

All manifold enclosures for nitrous oxide in excess of 200 cubic feet of manifold capacity shall be vented to the outside and the cylinder or manifold shall be protected with check valves or alarms.



3. Exposure Effects and First Aid Matrix

Similar instructions are to be followed as propane documentation for Inhalation, eye and skin contact.

4. Spill, Leak and Disposal Procedures

Eliminate all potential sources fuel and ignition. Evacuate all non-essential personnel to an area upwind. Disposal Return unused product to the supplier. Disposal of nitrous oxide must be done in an environmentally acceptable manner in compliance with all applicable national and local codes. Contact Health and Safety department as soon as the problem occurs.

5. Special Protective Measures Matrix

Special Protective Measures

Category Comments

Ventilation

Local exhaust and general room ventilation may both be essential in work areas to prevent accumulation of explosive

mixtures. If mechanical ventilation is used, electrical equipment must meet National Electrical Code requirements.

Eye and Skin Protection

Safety glasses with side shields, leather gloves, and safety shoes. If prevent exposure to liquid phase when handling the refrigerated liquid, add a long-sleeved shirt and face shield.

Respiratory Problems

For excessive gas concentrations, use only NIOSH/MSHA-approved, self-contained breathing apparatus.



Fault Tree Analysis for Feedline system test: A fault tree analysis as mentioned in the pages above will be used for analysis of the Feedline system test closely.

AND gate: An AND gate shown below, a top level system will fail if ALL inputs fail.

OR gate: In an OR gate shown below, a top level system will fail if ANY of its inputs fail.



SAFETY CHECKLIST

Location: Date:

Inspected by:

YES NO N/A COMMENTS

Ventilation

Removal of fumes and dust

Temperature and humidity control

Noise levels

Vibration

Lighting - General purpose

- For a particular task

- Absence of glare

- Work Area

- Exterior

Ergonomics - layout of work area

Work Environment

YES NO N/A COMMENTS

Drinking Water

Washing Facilities

Clean Area

Hygiene

YES NO N/A COMMENTS

Training provided to each employee assigned to a job

and/or machine

Employees receive training in the usage of the required

personal protective equipment provided

Training



YES NO N/A COMMENTS

Procedures

Safety Showers

Eye Wash units

Breathing apparatus and/or respirators

First Aid kit

First Aid Officers - name

Antidotes where applicable

Torches

Loud Hailer

Emergency lighting

Emergency Equipment

YES NO N/A COMMENTS

Fire separation - walls and floors

- Tanks

- Doors and Windows

Isolation of risk areas

Fire detection system

Alarm and Emergency Evacuation system

Are Alarms audible in all areas

Fire extinguisher - type

- location

- serviced

- operator training

Fire hose reels

Fire hydrants

Access for Fire Brigade

Fire Protection

YES NO N/A COMMENTS

Procedures Established

Display of floor plans and escape routes

Communication system

Exits - unobstructed & open from inside

Evacuation

YES NO N/A COMMENTS

Coats/Overalls

Eye protection

Gloves

Footwear

Respirators

Personnel Protection



YES NO N/A COMMENTS

OS&H Policy displayed

OS&H committee - employee representative

Responsibilities defined

Departmental Safety Officer (DSO)

Safety manual

Qualifications and training of operators

OS&H review of research projects

Accident and incident reporting

Hazard reporting and follow up

Provision for visitors

Management

YES NO N/A COMMENTS

OSHA standards comliant

Stored in a flammable liquids cabinet

Quantities are not excessive

Suitable containers

Spark proof electrics

Static energy control

Class B fire extinguisher near by

Correct signage

Flammable Liquids

YES NO N/A COMMENTS

Number of cylinders inside rooms

Cylinders secured

Segregation of incompatible gases

Transported on appropriate trolley

Serviced regularly

Tamper proof

Correct signage

Check lines carrying gas

Compressed Gas



YES NO N/A COMMENTS

Material Safety Data Sheets (MSDS)

Register of chemicals

Containers - suitable type

- condition

Correct labeling

Only sufficient stock

Storage - suitable

- condition

Segregation into classes

Waste collection and disposal Spill procedures

Chemicals

.



9. Appendix

9.1. NASA CEA Output



10. References

Hill, Philip. Petersen, Carl. Mechanics and Thermodynamics of Propulsion. 2nd Ed. Ashley-Wesley. 1992.

Holman, J.P. Heat Transfer. 9th Ed. McGraw-Hill. 2002. Humble, Ronald W. Henry, Gary N. Larson, Wiley J. Space Propulsion Analysis and Design. McGraw-Hill. 1995. Poling, Bruce E. The Properties of Gases and Liquids. 5th Ed. McGraw-Hill. 2001. Air Products and Chemicals, Inc.,, "Nitrous Oxide (N20)." 01/15/08 <http://www.airproducts.com/nr/rdonlyres/8c46596e-2f7d-4895-b12a-e54cd63e1996/0/safetygram20.pdf>. Environmental Health and Safety, "Compressed Gas Cylinder Policy." (2007): "MATERIAL SAFETY DATA SHEET: PROPANE." 11/6/02 1-5. 01/14/08 <http://www.georgepropane.com/PDF/MATERIAL%20SAFETY%20DATA%20SHEET.pdf>. "NASA Technical Standard." Facility System Safety Guidebook. January 1998. NASA. 03 Sep 2007 <http://www.hq.nasa.gov/office/codeq/87197c-6.pdf>. "Other Hazard Analysis Methodologies." January 1998. NASA. 04 Sep 2007 <http://www.hq.nasa.gov/office/codeq/871977-c.pdf>. "System and Safety Office." 03/18/2008 <http://www.safety.uwa.edu.au/__data/page/10280/workplace_safety_checklist.pdf>.

llv propulsion system 100 pct design report

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