Lunar Lander Propulsion System: 100% Design Review             AE 445 Spacecraft Detail Design            Department of Ae...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                Pag...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                  P...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION                                                   ...
Llv Propulsion System 100 Pct Design Report
Llv Propulsion System 100 Pct Design Report
Llv Propulsion System 100 Pct Design Report
Llv Propulsion System 100 Pct Design Report
Llv Propulsion System 100 Pct Design Report
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Llv Propulsion System 100 Pct Design Report
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Llv Propulsion System 100 Pct Design Report
Llv Propulsion System 100 Pct Design Report
Llv Propulsion System 100 Pct Design Report
Llv Propulsion System 100 Pct Design Report
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Llv Propulsion System 100 Pct Design Report

  1. 1. 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
  2. 2. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page i Team MembersExecutive StaffNicholas H. Perry, Project ManagerJohn C. Swatkowski, Research DirectorMitchell Graves, Budget ManagerShailesh Kumar, System Safety OfficerMission Assurance DivisionJohann Schrell, LeadEric McLaughlinBrendan McMahonConfiguration Management DivisionRobert J. Scheid, LeadJessica ChenTodd A. SnyderThermal Management DivisionJade Pomerleau, LeadBen KlammChi ZhangManufacturing, DivisionChukwuma Akosionu, LeadKristopher GreerJoe TaborSystems Integration & Testing DivisionEric J. Thompson, LeadMaggie CordovaAutumn GeeGary KelleyAndrew KreshekNicholas Rehak
  3. 3. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page ii Special ThanksTeam Cynthion would like to thank the following people who helped make this project areality:to Bogdan Udrea, Ph.D., for collaborating with Team Cynthion to make the lunar landerpropulsion 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 theproject as a mentor and instructor.
  4. 4. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page iii Table of Contents1. Introduction ................................................................................................................. 12. Main Rocket Thrusters Design Parameters................................................................. 2 2.1. Performance Metrics ............................................................................................ 2 2.2. Engine Dimensions ............................................................................................ 12 2.3. Injector Design ................................................................................................... 13 2.4. Engine Throttling ............................................................................................... 14 2.5. Ignition System .................................................................................................. 15 2.6. Backup Ignition System ..................................................................................... 15 2.7. ERPL Liquid Rocket Engine Creator ................................................................. 16 2.8. NASA Chemical Equilibrium Analysis ............................................................. 193. Propulsion System Configuration ............................................................................. 20 3.1. CATIA................................................................................................................ 20 3.2. Connections and Assembly ................................................................................ 27 3.3. NASTRAN Pressure Analysis ........................................................................... 29 3.4. Mass Budget ....................................................................................................... 395. Manufacturing Plan ................................................................................................... 68 5.1. Raw Materials and Hardware ............................................................................. 68 5.2. Fabrication Process ............................................................................................ 70 5.3. Assembly Process ............................................................................................... 82 5.4. Part Assembly .................................................................................................... 826. Systems Integration & Testing.................................................................................. 84 6.1. Feedline System ................................................................................................. 84 6.2. Electrical System for Solenoid Valves ............................................................... 85 6.3. Test Stand ........................................................................................................... 86 6.4. Data Acquisition (DAQ) .................................................................................... 87 6.5. Test Preparation Procedures ............................................................................... 93 6.6. Fuel Tank Air Evacuation .................................................................................. 95 6.7. Fuel Tank Fill (Propane) .................................................................................... 97 6.8. Fuel Tank Fill (Water)........................................................................................ 99 6.9. Assembly Procedure ......................................................................................... 100 6.10. Test Area Overview ...................................................................................... 103 6.11. Feed Line System Testing Overview............................................................ 114 6.12. Instrumentation ............................................................................................. 120 6.13. Igniter Testing Procedure ............................................................................. 122 6.14. Ignition Test Results ..................................................................................... 123 6.15. Engine Firing Test ........................................................................................ 1257. Project Economics .................................................................................................. 1288. System Safety Program Plan ................................................................................... 1319. Appendix ................................................................................................................. 158 9.1. NASA CEA Output .......................................................................................... 15810. References ............................................................................................................ 198
  5. 5. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page iv Table of Figures, Tables, and EquationsFigure 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 andManifold ............................................................................................................................ 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
  6. 6. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page vFigure 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
  7. 7. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page viFigure 88: Oxidizer System ............................................................................................ 117Figure 89: Pressurant System.......................................................................................... 118Figure 90: Fuel System ................................................................................................... 120Figure 91: Fault Tree Analysis ....................................................................................... 145Table 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 ........................................................................ 130Equation 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
  8. 8. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page viiEquation 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
  9. 9. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 11. 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 teams requirements. It is the intention of this report to provide a complete record of Team Cynthions design project since the fusion of Team ALLSTAR and Team Medea. A full overview of the has been included.
  10. 10. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 22. Main Rocket Thrusters Design Parameters2.1. 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. 3 2.5 2 Pc =10 Pc = 15 Pc = 20 1.5 Pc = 25 Pc = 30 Pc = 35 1 Pc = 38 0.5 0 0 1 2 3 4 5 6 Ae/At 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.
  11. 11. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 3 3 y = 0.0066x4 - 0.1435x3 + 1.228x2 - 5.0356x + 9.153 2.5 2 Tave 1.5 Tmin Tmax 1 0.5 0 0 1 2 3 4 5 6 Ae/At 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.
  12. 12. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 4 250 245 240 235 230 225 220 215 210 205 200 0 1 2 3 4 5 6 7 8 O/F Figure 3: Specific Impulse for Varied O/F Ratio 1650 1600 1550 1500 1450 1400 1350 0 1 2 3 4 5 6 7 8 O/F Figure 4: Characteristic Velocity for Varied O/F Ratio
  13. 13. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 5 3500 3000 2500 2000 1500 1000 500 0 0 1 2 3 4 5 6 7 8 O/F Figure 5: Chamber Temperature for Varied O/F Ratio 2.95 2.9 2.85 2.8 2.75 2.7 0 1 2 3 4 5 6 7 8 O/F Figure 6: Exit Mach Number for Varied O/F Ratio
  14. 14. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 6 120000 100000 80000 60000 40000 20000 0 0 1 2 3 4 5 6 7 8 O/F Figure 7: Nozzle Exit Pressure for Varied O/F Ratio 1.473 1.472 1.471 1.47 1.469 1.468 1.467 1.466 1.465 0 1 2 3 4 5 6 7 8 O/F Figure 8: Coefficient of Thrust for Varied O/F Ratio
  15. 15. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 7 1.55 1.5 1.45 1.4 1.35 1.3 1.25 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
  16. 16. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 8 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. 1.4 1.2 1 0.8 0.6 0.4 0.2 0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 Thrust (N) Figure 10: Exit Pressure Over Thrust Range
  17. 17. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 9 1.80000 1.60000 1.40000 1.20000 1.00000 0.80000 0.60000 0.40000 0.20000 0.00000 0 500 1000 1500 2000 2500 3000 3500 4000 4500 Thrust (N) Figure 11: Mass Flow Rate Over Thrust Range 245.2 245 244.8 244.6 244.4 244.2 244 243.8 0 500 1000 1500 2000 2500 3000 3500 4000 4500 Thrust (N) Figure 12: Specific Impulse Over Thrust Range
  18. 18. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 10 3270 3260 3250 3240 3230 3220 3210 3200 3190 3180 0 500 1000 1500 2000 2500 3000 3500 4000 4500 Thrust (N) Figure 13: Chamber Temperature Over Thrust Range
  19. 19. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 11 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 Chamber Pressure 29 bar Specific Impulse 244.74 sec Characteristic Velocity 1629.9 m/s Characteristic Length 0.89 m Mass Flow Rate 1.272 kg/s Exit Pressure 0.959 bar 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 Chamber Pressure 38 bar Specific Impulse 204.57 sec Characteristic Velocity 1367.7 m/s Characteristic Length 0.89 m Mass Flow Rate 1.994 kg/s Exit Pressure 1.30 bar Exit Mach 2.799 Ae/At 4.7 Ac/At 8
  20. 20. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 12 Table 5: Performance Metrics for Thrust Chamber at 60% Nominal at O/F=3 Operating Performance Metrics Thrust 3053 Chamber Pressure 29 bar Specific Impulse 204.23 sec Characteristic Velocity 1365.9 m/s Characteristic Length 0.89 m Mass Flow Rate 1.524 kg/s Exit Pressure 0.986 bar Exit Mach 2.719 Ae/At 4.7 Ac/At 82.2. 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.
  21. 21. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 132.3. 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 Propane Nitrous Oxide 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 Propane Nitrous Oxide 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
  22. 22. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 142.4. 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.
  23. 23. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 152.5. 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.
  24. 24. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 162.7. 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.
  25. 25. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 17 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
  26. 26. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 18Outputs: 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:
  27. 27. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 19 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 Check2.8. 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.
  28. 28. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 203. Propulsion System Configuration3.1. 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.
  29. 29. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 21 Figure 14: Assembled Engine
  30. 30. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 22 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. Figure 15: Supersonic The lip is designed to provide a 6.35mm flat area that can Nozzle 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 Figure 16: Combustion Chamber row of injectors to the top of the combustion chamber will be used for mounting other components.
  31. 31. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 23 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 Figure 17: Cooling Jacket measure of safety to the r mass 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 Figure 18: Fuel Inlet Manifold 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.
  32. 32. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 24 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 Figure 19: Nitrous Injection Plate 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.
  33. 33. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 25 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. Figure 20: Nitrous Injection Dome 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 Figure 22: Copper Connection Ring 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 Figure 21: Steel Connection Ring 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.
  34. 34. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 26 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.
  35. 35. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 273.2. 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.
  36. 36. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 28 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
  37. 37. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 293.3. 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)
  38. 38. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 30 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.
  39. 39. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 31 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.
  40. 40. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 32 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.
  41. 41. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 33 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.
  42. 42. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 34 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.
  43. 43. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 35 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.
  44. 44. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 36 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.
  45. 45. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 37 Figure 31: Section Cut of Jacket and Chamber: Von Mises Stress of Assembly
  46. 46. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 38 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.
  47. 47. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 393.4. 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 0.994Nozzle Tellurium Copper 1 0.625 0.625Cooling Jacket Steel 1 1.565 1.565Injector Plate Stainless Steel 1 1.813 1.813Connection Ring (Copper) Tellurium Copper 1 1.685 1.685Connection Ring (Steel) Steel 1 1.417 1.417Fuel Inlet Manifold Red Brass 1 0.487 0.487Injector Dome Stainless Steel 1 0.321 0.321O-Rings Permatex Sealant 2 0.003 0.006Bolts Steel 4 0.070 0.280Igniter (several materials) 3 0.004 0.012Washers Steel 8 0.008 0.064Nuts Steel 4 0.015 0.060Engine 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)..
  48. 48. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 404. Thermal Analysis4.1. 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. g .8 .4 g g g E Equation 1: Gas/Wall Heat Transfer Coefficient 1/ 5 2/3 l l l l Equation 2: Coolant/Wall Heat Transfer Coefficient
  49. 49. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 41 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
  50. 50. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 42 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: y rc1 sin 1 r rth cos 1 rc1 Equation 4: Section 1 y-position Equations Equation 5: Section 2 y-position Equation y y3 rc 3 sin r3 r 1 rc3 cos rc 3 Equation 6: Section 3 y-position Equations y rc 4 sin 1 r rth cos 1 rc 4 Equation 7: Section 4 y-position Equations 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
  51. 51. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 43 1 0 0 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 Therm Temperature Density Viscosity Cp Gamma Conduct (K) (kg/m^3) (millipoise) (kJ/(kg*K)) (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 w q Twg Twc tw 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
  52. 52. Lunar Lander Vehicle Propulsion System: 100% Design ReviewTeam CYNTHION Page 44 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. qA m Cp c Tc 2 Tc1 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: w w Tg 1 Tc t w hc t w hg Twg w w w w 1 1 t w hc t w hg t w hc t w hg Equation 15: Gas Wall Temperature w Twg Tc t w hc Twc w 1 t w hc Equation 16: Coolant Wall Temperature By combining Equation 13 and Equation 14, the coolant temperature at the end of the disc is found: hc Twc 1 Tc1 A Tc 2 Tc1 m Cp c 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)

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