1. 1 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Liquid Bi-Propellant Thruster Preliminary Design Review Program Manager : Adam Pender Lead Designer and Bringer of Problems: Jason Wennerberg Thermal and Performance Analyst: Arun Padmanabhan Manufacturing and Materials Analyst: Josh Revenaugh

2. 2 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Outline Introduction Design Process Thermal Analysis Performance Analysis Structural Analysis Manufacturability Fluid Interfaces Future Steps Conclusion

3. 3 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Schedule Milestones Preliminary Design Review –Now Critical Design Review –April 7 Hardware Completed –May 5

4. 4 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Develop thruster for Prospector-7 –Cal State University-Long Beach / Garvey Aerospace flight demonstration rocket. –Prototype first stage for Nano-Launch Vehicle –246 lbm propellant –550 lb GLOW –T/W = 4 –Payload TBD –Max Alt: 30,000-50,000 ft. (Burnout at 10,000 ft) Mission / Application ~22 ft 26 in Prospector-6

5. 5 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Requirements Thrust: 2200 lbf Chamber Pressure: 300 psi Burn Time: 20 seconds Nozzle Designed for Sea Level to 10,000 ft. Operation O/F = 2.27 (propellants deplete at same rate) Thruster Mass < 15 lbm Injector Pressure drop of 70 psi C* efficiency of 95% Interface –AN fittings (-8 if possible) –10+ inch plate with bolt pattern communicated to CSULB Development Static Tests Performed at Purdue

6. 6 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Pre-Launch Test Requirements Engine-only Static Test –Verify Performance Rocket Static Test –Verify Engine/Rocket Integration –2-4 Seconds if engine to be reused Static tests may be performed by CSULB at Mojave or at HPL

7. 7 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Propene Saturation Curve 100 F Saturation Margin: 80 psi Temperature Limit: 323 K (122 F)

8. 8 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Process Jason Wennerberg

9. 9 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Components of TCA Injector Chamber Nozzle

10. 10 3/9/2005 Thrust Chamber Assembly Preliminary Design Review The first step in the design was to pick propellants –LOX – propylene chosen for several reasons Customer has experience and access Allow for partial self pressurization of propellant tanks The mixture ratio is specified by CSULB based on the ratio that will give the best operability = 2.27. This allows for the propellant tanks to empty at the same rate A chamber pressure must be chosen –300 psi was chosen by the customer. Current tanks can handle 450 psi  300 psi chamber pressure after losses Cooling by passive means is possible (No dump or regenerative cooling required) Design Process

11. 11 3/9/2005 Thrust Chamber Assembly Preliminary Design Review With the information available we run the NASA thermochemistry code to obtain some useful data: –Chamber Temp (T c ) = 6341 R –C* = 6044 ft/s –Exit pressure (p e ) = 5.66 psi –Exit velocity (v e ) = 9627.8 ft/s –Cf vac = 1.593 –Specific heat ratio γ = 1.1398 –Molecular weight = 21.313 –Isp vac = 327.6 s Design Process

12. 12 3/9/2005 Thrust Chamber Assembly Preliminary Design Review With this data we can continue with the design of the engine. We would like to use the equation that relates mass flow rate to force and Isp so first we need C f at sea level, and then Isp at sea level and then finally mass flow rate through the engine. From NASA code Design parameters From NASA code Design parameter Design Process

13. 13 3/9/2005 Thrust Chamber Assembly Preliminary Design Review We know our O/F ratio so we can then split the mass flow into fuel and oxidizer: Where r is the mixture ratio The throat area is found with: We choose a contraction ratio of 2 to help with combustion stability Design Process

14. 14 3/9/2005 Thrust Chamber Assembly Preliminary Design Review We use the design parameter L* to find the size of the combustion chamber. We used an L* of 42.5 in because it has worked successfully in the past with RP-1. This is the volume needed Length of converging section with θc the converging half angle Volume that the converging section makes Use a cylinder to make the rest of the volume Design Process

15. 15 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector design pressure loss is 70 psi. We use.2*Pc = 60 psi for the drop across the orifices Area for injection is found with the pressure drop from the manifold to the chamber with: Cd is discharge coefficient =.80 Design Process We need to select hole sizes based on drill bits that can be purchased. By selecting the number of orifices that we want we can find the hole sizes that we need. Going back we can find the new mass flows and actual O/F.

16. 16 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Requirements – Chamber L* = 42.5 in Should withstand heat flux for burn time Should withstand any transient pressure Should not be overly complicated (Cheap to build) Cannot use regenerative cooling because of lack of pressure budget Use ablative liner and film cooling or O/F bias. Convergence ratio of 2 Need to be able to flange onto injector

17. 17 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Specs - Chamber Chamber Diameter = 3.69 in Length of chamber = 20.73 in Length of converging section ≈.64 in Diameter of throat = 2.61 in

18. 18 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Current Chamber Design Put drawing here

19. 19 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Nozzle Design Process Adam Pender

20. 20 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Requirements – Nozzle Expansion ratio = 8 75% bell to assist in weight reduction Manufacturing must be taken into consideration –Conical nozzle used to be cheaper to manufacture –CNC manufacturing has reduced cost of bell nozzle Uncooled NASA Dryden

21. 21 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Nozzle Contour Design Nozzle contour designed from Sutton guidelines. Nozzle code written which automatically draws nozzle curve with the following inputs: –Throat Radius –Contraction Ratio –Expansion Ratio –% Length (based on 30-deg. conical nozzle)

22. 22 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Nozzle Contour Guidelines

23. 23 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Nozzle Contour Code Output Nozzle Code Outputs –Figure Illustrating Nozzle –Text File to be used in Pro-E Example Nozzle Plot

24. 24 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Specs - Nozzle Length of nozzle –8.91 in (15° cone) –6.68 in (75% bell) 75% Bell –Lower Weight –Better Performance Bell Nozzle on Pump-Fed LRE

25. 25 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Current Nozzle Design

26. 26 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector Design Jason Wennerberg

27. 27 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Requirements - Injector By far the most complicated part of design ΔP = 70 psi Shouldn’t melt or scorch Provide combustion stability No inter-propellant seals Total flow rate = 8.45 lbm/s Ox flow rate = 5.87 lbm/s Fuel Flow rate = 2.58 lbm/s

28. 28 3/9/2005 Thrust Chamber Assembly Preliminary Design Review O-F-O Impinging Injector Injector provides for propellant mixing by impinging jets. Two oxidizer jets impinge on one fuel jet. OOF Fan

29. 29 3/9/2005 Thrust Chamber Assembly Preliminary Design Review O-F-O Injector Well known design process Better performance compared to pintle Allows for O/F biasing against wall and film cooling Propellants are well suited for this option –SG propylene =.5 –SG LOX = 1.14 –O/F = 2.27

30. 30 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector Sizing 18 – triplets 18 film cooling elements Oversize outboard oxidizer element to ensure jets stay away from the wall Impingement point length/ diameter of orifice should be ~ 5 Bore length/diameter of orifice should be > 3.5 to ensure Cd =.80 Manifolds – 10*area of orifices they feed

31. 31 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector Concept

32. 32 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector Concept

33. 33 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector Concept Put the picture here

34. 34 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector Performance Analysis With these sizes: Stream Lengths Bore Lengths

35. 35 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector Lengths

36. 36 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manifold Sizes

37. 37 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manifold Sizes

38. 38 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manifold Sizes A ox,in =.08165 in 2 A ox,out =.01437 in 2 A fuel =.01452 in 2 A film =.00226 in 2 Flow Area/ Injection Area Ox in = 2.296 Ox out = 7.738 Fuel = 9.043 Film = 33.186

39. 39 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector performance Velocities: –Ox = 88.35 ft/s –Fuel = 120.45 ft/s Momenta –Ox_out = 222 lb-in/s 2 –Ox_in = 210 lb-in/s 2 –Fuel = 224 lb-in/s 2 0.9911 : 1.0000 : 0.9375

40. 40 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector Fill Times Volumes Vox = 7.13508 in 3 Vf =.26875 in 3 Volumetric flows Qox = 142 in 3 /s Qf = 118.5 in 3 /s Fill times t ox =.05 sec t fuel =.002 sec

41. 41 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Combustion Stability Stable Unstable

42. 42 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Current Concept Summary Injector: O-F-O Injector Chamber: Ablative Lining Nozzle: 80% Bell

43. 43 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Heat Transfer Analysis Arun Padmanabhan

44. 44 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Steady State Heat Transfer Analysis on Injector Actual Injector Half Modeled Injector Half

45. 45 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Outline of Procedure Used 1.Get Chamber Properties from NASA code Density Sonic Velocity Viscosity Specific Heat Thermal Conductivity 2.Pick Mach number tangent to surface: 0.4 3.Steady state heat transfer iteration for gas-side wall temperature for oxidizer and fuel sections 4.Compute fuel pressure loss through injector

46. 46 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Fuel Section Analysis Copper Wall Fuel Flow Camber Gases Convection Conduction Convection

47. 47 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Gas-side Wall Temperature Iteration Steps 1.Guess Gas-Side Wall Temperature, Twg 2.Bartz Equation:

48. 48 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Fuel Section Iteration 3.Gas-Side Heat Flux: 4.Fuel-side Wall Temperature:

49. 49 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Fuel Section Iteration 5.Seider-Tate Forced Convection: Correlations as a function of temperature at 350psi for transport and physical properties of Propene from NIST Chemistry Web book Temperature at previous position used, T initial =405°R

50. 50 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Fuel Section Iteration 6.Fuel-side Heat Flux: 7.If fuel-side heat flux = gas-side heat flux, continue, else choose another guess for T wg 8.Fuel Temperature at current position:

51. 51 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Fuel Section Iteration 9.Pressure Iteration: 10.Move to next axial position and repeat

52. 52 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Oxidizer Section Analysis Same procedure as fuel-side except T wo is fixed at liquid oxygen temperature Copper Wall Stagnant O 2 Camber Gases Conduction Convection Fixed Temperature, Liquid O 2 = 162°R Copper Wall

53. 53 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Fuel-Side Results, M = 0.4 dTwg ≈ 78°R

54. 54 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Overall Results, M = 0.4

55. 55 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Fuel Convective Heat Transfer Coefficient

56. 56 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Fuel Section Heat Flux

57. 57 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Oxygen Section Heat Flux

58. 58 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Adiabatic Flame Temperature vs. O/F Ratio

59. 59 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Cstar vs. O/F Ratio

60. 60 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Ivac vs. O/F Ratio

61. 61 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Thrust Coefficient vs. O/F Ratio

62. 62 3/9/2005 Thrust Chamber Assembly Preliminary Design Review First Order Structural Calculations Adam Pender

63. 63 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Structural Calculations Uses Low and High Temperature Yield Strength Limits for Quasi-Transient Analysis. Size 40-5 Pipe for Chamber Wall –OD: 5.563 in. –ID: 5.047 in. –Pipe Thickness: 0.258 in. –Resulting Ablative Liner Thickness: 0.6785 in. Carbon Steel, Stainless Steel, and Aluminum Evaluated. –Working temperature of 500F for 224 Aluminum was much lower than Stainless Steel –Aluminum rejected early due to poor high temperature strength

64. 64 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Chamber Stress Calculations Barlow’s Formula Solved for Max Pressure for Hoop Stress S=PD/2t  P=2tS/D S=hoop stress, in psi P=internal pressure D=outside diameter of the pipe in inches t=normal wall thickness, in inches Resulted in high pressure limits >2000 psi Lead to more appropriate calculation

65. 65 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Chamber Stress Calculations: Part Two Von Mises Stresses calculated for many material, temperature, and pressure combinations. –Takes into account the cumulative stresses on part –Hoop Stress –Longitudinal Stress –Radial Stress Modeled chamber wall as wall of a pressure vessel (it is) Determined the yield chamber pressure.

66. 66 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Stainless Steel Yield Strength vs. Temperature

67. 67 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Max Chamber Pressure vs. Temperature

68. 68 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Bolt Pattern Bolt pattern applies to: –Injector-Chamber Flange –Chamber-Nozzle Flange (If Necessary) Fastener Specifics from McMaster-Carr –Steel Bolts Zinc Plated.25 in thread diameter 150000 psi tensile stress limit Head Width: 7/16 in. Head Height: 5/32 in. –Stainless Steel Hex Nuts 18-8 Stainless Steel Width: ½ in. Height: 15/64 in. –Stainless Steel Washers 316 Stainless Steel ID: 0.281 in. OD: 0.625 in. Thickness: 0.062

69. 69 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Max Pressure vs. Number of Bolts

70. 70 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Flange Thickness Specification Flange sized to remain intact at very high chamber pressures. Calculations based on shear at fastener edges Assumptions: –6 bolts –0.625 in. washers –Shear concentrated on 30% of the washer circumference –800 deg F flange temperature –Minimum Stainless Steel Strength at 800 deg F: 28000 psi (310S and 316) Max Pressure=L shear t flange S / A chamber

71. 71 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Max Pressure vs. Flange Thickness

72. 72 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing Josh Revenaugh

73. 73 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Material Selection Copper Injector (c101) High resistance to particle impact High Lox pressure rating (8000 psi)

74. 74 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Engine Assembly

75. 75 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Ablative Throat

76. 76 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Parts – Lox Dome

77. 77 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Lox Dome

78. 78 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Lox Dome

79. 79 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Lox Dome

80. 80 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Parts – Deflector

81. 81 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Deflector

82. 82 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Deflector

83. 83 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Deflector

84. 84 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Parts – Lox Plate

85. 85 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Lox Plate

86. 86 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Lox Plate

87. 87 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Lox Plate

88. 88 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Parts – Injector

89. 89 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Injector

90. 90 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Injector

91. 91 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Injector

92. 92 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Parts – Chamber

93. 93 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Chamber

94. 94 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Parts – Nozzle

95. 95 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing - Nozzle

96. 96 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Nozzle

97. 97 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Lox Dome AN16-AN8 Fitting (1” to ½”)

98. 98 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing – Injector 8AN-4NPT Fittings ½” to ¼”

99. 99 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Fluid Interface Rocket Interface AN16( 1”) fittings for Propene and Lox –V fuel =16.8 ft/s (Before Split) –V fuel =14.7 ft/s (After Split) –V LOX =20.2 ft/s Component Interface Six ½” to 1/4” (8AN- 4NPT) fittings in the injector for fuel (Reduced Further in Injector Fuel Manifold) One 1”- ½” (AN16-AN8) fitting in the Lox Dome for Lox

100. 100 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Engine Masses Partweight (lb) Lox Dome7.40 Deflector0.17 Lox Plate3.00 Injector7.00 Chamber22.80 Nozzle22.20 Total62.57

101. 101 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Brazing Silver is used as the braze alloy (BAG-8) Braze temperature 1500 deg F A few hours in the oven

102. 102 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Manufacturing Timeline PartProcess days →123456789 10101 1212 1313 1414 1515 1616 1717 Lox Dome shipping manufacturing Deflector shipping manufacturing brazing Lox Plate shipping manufacturing brazing Injector shipping manufacturing brazing Chamber shipping manufacturing Nozzle shipping manufacturing

103. 103 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Price List – Materials DescriptionMaterialprice Lox Dome 2" x7 1/2" ODc101370.30 Deflector 2" x 2" ODc10151.48 Lox Plate 1/4" x 7 1/2" ODc10190.34 Injector 3/4" x 7 1/2" ODc101169.60 Flange 1/2" x 7 1/2" ODcarbon steel38.69 Chamber 40-5 black pipe (21ft)carbon steel266.85 Nozzle Extension 4 1/4" x 8 1/4" ODcarbon steel109.42 Nozzle Throat 5" x 6" ODcarbon steel38.44 Total1135.12

104. 104 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Price List – Manufacturing DescriptionquantitypriceTotal Cost Drill Bit #68 (.0310")201.8436.80 Drill Bit #48 (.0760")103.8338.30 Drill Bit 5/64" (.0781")102.8528.50 Drill Bit #47 (.0785")102.7227.20 Total130.80

105. 105 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Price List – Parts DescriptionquantitypriceTotal Cost Bolts 1/4-20, 2-1/2" Grade 8 (25)19.28 Nuts 1/4-20, ASTM Standards (25)19.87 Washers 1/4" SAE Standards (50)15.01 Gasket Graphite 1/16" thick 24"x24"140.53 O-ring 339 Teflon, 3-1/4 ID121.89 Total86.58

106. 106 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Costs Descriptionprice Materials1135.12 Manufacturing130.80 Parts86.58 Brazing350.00 Total1702.50

107. 107 3/9/2005 Thrust Chamber Assembly Preliminary Design Review We’re Almost Done, I Promise. Adam Pender

108. 108 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Potential Failure Modes Injector Face Melts (Thermal) Chamber Overheats and Bursts (Von Mises) Flange Bolts Fail (Tensile Failure) Flange Fails (Shear) Nozzle Distorts and Buckles During Startup Ablative Liner Fails (Potential throat blockage or chamber failure)

109. 109 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Future Action Items 2 nd Order Analyses –Acoustic Analysis of Chamber –Thermal (Improve Thermal Model) –Structural (Improve Structural Calculations) –Performance (Effect of O/F Biasing) Pressure Loss Analysis Finish Drawings Develop Manufacturing Tolerances Choose Suppliers

110. 110 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Questions?

111. 111 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Appendix Numbers Concept Design Review Documentation Control Additional Design Guidelines

112. 112 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Numbers Summary O/F2.2729 Pc300psi F2200lbf ε8 Tc6340R c*6044ft/s Pe5.66psi ve9628ft/s Cf)vac1.593 γ1.1398 MW21.313lb/lbmole Ivac327.6s Iopt299.2s L*42.5in εc2 Pc/Pa20.41 Cf1.44207 η0.95 Isp257.35s c8280ft/s

113. 113 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Numbers mdot8.4519lb/s mo5.8695lb/s mf2.5824lb/s Dc3.692inChamber Ac10.706in^2 Dt2.6107inThroat At5.353in^2 De7.3841inExit Ae42.82in^2 Lc20.72inLength of chamber

114. 114 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Numbers ΔP60psiinjector pressure drop ρ lox0.0412lb/in^3density lox ρ fuel0.0218lb/in^3 Cd0.8 discharge coeff Dfilm0.031infilm orifices Dox,out0.0781inoutside orifices Dox,in0.076ininside orifices Dfuel0.0785infuel orifices Aox0.167887in^2 Afuel0.100703in^2 Vox1060in/s88.33333ft/s Vfuel1445in/s120.4167ft/s mom_ox_o222lb-in/s mom_ox_in210lb-in/s mom_f224lb-in/s

115. 115 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Numbers ac49259in/s4104.917ft/s f1t7821Hz f2t12974Hz f3t17845Hz f1r16276Hz f2r29800Hz Df/Vf5.40E-05s

116. 116 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Liquid Bi-Propellant Thruster Concept Design Review Adam Pender Jason Wennerberg Arun Padmanabhan Josh Revenaugh

117. 117 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Outline Schedule and Milestones Mission System Requirements Requirements Compliance Matrix Design Process Concepts Considered Chosen Concept Details Application to SLV Project Conclusion

118. 118 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Schedule Milestones Preliminary Design Review –Now Critical Design Review –April 7 Hardware Completed –May 5

119. 119 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Requirements Compliance Interface: Document Fittings and Flange Bolt Pattern Nozzle: Design with Area Ratio for Max Efficiency from 0-10,000 ft. Thrust, Burn Time, C* Efficiency, O/F : Static Tests Thruster Mass: Scale

120. 120 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Process

121. 121 3/9/2005 Thrust Chamber Assembly Preliminary Design Review The first step in the design was to pick propellants –LOX – propene chosen for several reasons Performance is increased over LOX - RP-1 Customer has experience and access Allow for partial self pressurization of propellant tanks The mixture ratio is specified by CSULB based on the ratio that will give the best operability = 2.27 A chamber pressure must be chosen –300 psi was chosen by the customer. Current tanks can handle 450 psi  300 psi chamber pressure after losses Cooling by passive means is possible (No dump or regenerative cooling required) Design Process

122. 122 3/9/2005 Thrust Chamber Assembly Preliminary Design Review With the information available we run the NASA thermochemistry code to obtain some useful data: –Chamber Temp (T c ) = 6374 R –C* = 6073 ft/s –Exit pressure (p e ) = 5.7 psi –Exit velocity (v e ) = 9673.5 ft/s –Cf vac = 1.593 –Specific heat ratio γ = 1.1396 –Molecular weight = 21.232 –Isp vac = 329.3 s Design Process

123. 123 3/9/2005 Thrust Chamber Assembly Preliminary Design Review With this data we can continue with the design of the engine. We would like to use the equation that relates mass flow rate to force and Isp so first we need C f at sea level, and then Isp at sea level and then finally mass flow rate through the engine. From NASA code Design parameters From NASA code Design parameter Design Process

124. 124 3/9/2005 Thrust Chamber Assembly Preliminary Design Review We know our O/F ratio so we can then split the mass flow into fuel and oxidizer: Where r is the mixture ratio The throat area is found with: We want the diameter of the chamber to be at least twice the diameter of the throat to help with combustion stability Design Process

125. 125 3/9/2005 Thrust Chamber Assembly Preliminary Design Review We use the design parameter L* to find the size of the combustion chamber. We used an L* of 42.5 in because it has worked successfully in the past with RP-1. This is the volume needed Length of converging section with θc the converging half angle Volume that the converging section makes Use a cylinder to make the rest of the volume Design Process

126. 126 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector design pressure loss is 70 psi. Area for injection is given by: Cd is discharge coefficient =.65 for pintle or.72 for straight elements Design Process

127. 127 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Requirements – Chamber L* = 42.5 in Should withstand heat flux for burn time Should not be overly complicated (Cheap to build) Cannot use regenerative cooling because of lack of pressure budget Use ablative liner or thermal barrier coating and O/F biasing for cooling Convergence ratio of 3.5-4

128. 128 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Specs - Chamber Chamber Diameter = 5 in Length of chamber = 9.32 in Length of converging section = 3.63 in Diameter of throat = 2.61 in

129. 129 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Requirements - Injector By far the most complicated part of design ΔP = 70 psi Shouldn’t melt or scorch Provide combustion stability No inter-propellant seals Total flow rate = 8.51 lbm/s Ox flow rate = 6.21 lbm/s Fuel Flow rate = 2.3 lbm/s

130. 130 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Pintle Option Easy to manufacture Simpler design Increases combustion stability –Fuel orifices =.0292 in –Number of fuel elements = 152 –Diameter of pintle = 1.56 in –Diameter of oxidizer orifice = 1.64 in

131. 131 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Pintle Injector: Manufacturing Pros Few parts to manufacture Cons Little room for error in manufacturing (Tilted pintle would effect LOX injection)

132. 132 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Flat Face Injector More traditional option/better known Better performance O/F biasing against wall Harder to manufacture (hole sizes are small and numerous) –Fuel orifice size =.0319 in –Number of elements = 128 –Ox orifice size =.0320 in –Number of elements = 256

133. 133 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Flat Face Injector 1 Pros Can be manufactured with minimal brazing (Hand brazing or sweating only) Cons Many parts to manufacture Extra structural supports needed?

134. 134 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Flat Face Injector 2 Pros Few parts to manufacture Solid (no extra structural supports) Cons Outsourcing for brazing (time & $)

135. 135 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Injector Selection Criteria Wants: –Large knowledge base –Low-cost design –Easily manufactured –Meet performance requirements –Stable combustion

136. 136 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Materials Compatibility LOX is highly corrosive, need resistive materials High Resistance –Copper, Nickel, Nickel alloys, and copper- nickel alloys Medium Resistance –Stainless steels, aluminum alloys Low Resistance –Carbon steel, iron

137. 137 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Materials Monel (Copper-Nickel Alloys) –High flame resistance (up to 10,000 psi) –High strength to weight ratio Inconel (Nickel Alloys) –High flame resistance (up to 10,000 psi) –High strength

138. 138 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Materials Copper –High flame resistance (up to 8000 psi) –Cheap Stainless Steel –Medium flame resistance (304 up to 1000 psi, 316 up to 500 psi) –Good structural properties –Cheap

139. 139 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Preliminary Material Choices Copper or Stainless Steel for Lox components –Budget Constraints –Manufacturability Nickel plating for injector face heat management if necessary

140. 140 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Lox Seals All seals should be “Hard Seals” –Metal-on-metal contact Pure Teflon ® –One of a few materials that can be used for sealing components together if a hard seal is not possible –Must be many layers for a robust seal to allow for thermal contraction

141. 141 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Current Concept Summary Injector: Flat Face Injector Chamber: Ablative Lining Nozzle: 80% Bell (Conical Shown) 2” 9.44”3.61” 8.95” 20° 15° Injector Combustion Chamber 2.62” 5”  = 8

142. 142 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Applicability to SLV Engine could be used on a Purdue-designed rocket. Engine could be modified or scaled to suit the specific needs of the project. Design process could be followed for a new engine.

143. 143 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Near Future Steps Injector Iterations Heat Transfer / Thermal Analysis Material Selection Structural Analysis Interface Design Nozzle Design Failure Mode Analysis

144. 144 3/9/2005 Thrust Chamber Assembly Preliminary Design Review ¿ Questions?

145. 145 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Documentation Control –Anyone can post a file they deem useful, using the file format. –The Designers have sole access to editing any drawing files. File Format: –component_mmdd_ver.ext File Deletion –Old files are to be moved to the "OLD" directory for at least one week before they are deleted. –Only the Designers or the Project Manager can delete any file, which shall occur at regular intervals to clear out any unnecessary files.

146. 146 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Prospector 7 is a modified version of Prospector 5 Thrust: 2,200 lbf; GLOW: 550 lbm 6 fully loaded tanks of each propellant: 246 lbm Burnout altitude: 10,000 ft; Coast: 30,000 – 40,000ft Engine Design Guidelines: @ O/F: 2.27 Propylene and LOX will deplete at about the same rate Mass: 15 lbm Burntime: 20 seconds Injector: flat head designs due to higher performance potential and easier control of temperatures Tank Pressure: 450 psi; Injector Pressure Loss: 70 psi Chamber Design: Ablative chambers, with some including graphite throat inserts Silica tape from Cotronics and an epoxy from a NASA TPS L*: 1 meter C* eff : 0.95 (if possible) Specified Design Guidelines

147. 147 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Objectives Design a liquid propellant thrust chamber assembly to be launched as a flight demonstration engine for CalState Long Beach. Design with the constraint that hardware must be designed and built using Purdue facilities and personnel Cost?? Adam – feel free to add stuff here

148. 148 3/9/2005 Thrust Chamber Assembly Preliminary Design Review Design Objectives Level 2 Thrust ~ 2200 lbf (Sea Level) Chamber Pressure ~ 300 psi Propellants: LOX-propylene Mixture ratio (O/F) = 2.27 Burn time ~ 15s Engine mass ~ 15 lbm