Liquid BiPropellant Thruster Preliminary Design Review Program Manager

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 1

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

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

Mission / Application • 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) ~22 ft Prospector-6 26 in 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 4

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 5

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 6

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

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

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

Design Process • 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) 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 10

Design Process • With the information available we run the NASA thermochemistry code to obtain some useful data: – – – – 3/9/2005 Chamber Temp (Tc) = 6341 R C* = 6044 ft/s Exit pressure (pe) = 5. 66 psi Exit velocity (ve) = 9627. 8 ft/s Cfvac = 1. 593 Specific heat ratio γ = 1. 1398 Molecular weight = 21. 313 Ispvac = 327. 6 s Thrust Chamber Assembly Preliminary Design Review 11

Design Process 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 Cf 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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 12

Design Process 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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 13

Design Process 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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 14

Design Process • • • 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 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. 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 15

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 16

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 17

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

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

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 3/9/2005 NASA Dryden Thrust Chamber Assembly Preliminary Design Review 20

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) 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 21

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

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

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 24

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

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

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 27

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

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 29

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 30

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

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

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

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

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

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

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

Manifold Sizes • • • Aox, in =. 08165 in 2 Aox, out =. 01437 in 2 Afuel =. 01452 in 2 Afilm =. 00226 in 2 Flow Area/ Injection Area Oxin = 2. 296 Oxout = 7. 738 Fuel = 9. 043 Film = 33. 186 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 38

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 39

Injector Fill Times Volumes Fill times Vox = 7. 13508 in 3 tox=. 05 sec Vf =. 26875 in 3 tfuel=. 002 sec Volumetric flows Qox = 142 in 3/s Qf = 118. 5 in 3/s 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 40

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

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

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

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

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 45

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

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

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

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

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

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

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

Fuel-Side Results, M = 0. 4 d. Twg ≈ 78°R 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 53

Overall Results, M = 0. 4 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 54

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

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

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

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

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

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

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

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

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 500 F for 224 Aluminum was much lower than Stainless Steel – Aluminum rejected early due to poor high temperature strength 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 63

Chamber Stress Calculations • Barlow’s Formula Solved for Max Pressure for Hoop Stress • S=PD/2 t P=2 t. S/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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 64

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. 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 65

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

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

Bolt Pattern • Bolt pattern applies to: – Injector-Chamber Flange – Chamber-Nozzle Flange (If Necessary) • Fastener Specifics from Mc. Master-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 • • 3/9/2005 316 Stainless Steel ID: 0. 281 in. OD: 0. 625 in. Thickness: 0. 062 Thrust Chamber Assembly Preliminary Design Review 68

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

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 (310 S and 316) • Max Pressure=Lsheartflange. S / Achamber 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 70

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

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

Material Selection • Copper Injector (c 101) • High resistance to particle impact • High Lox pressure rating (8000 psi) 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 73

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Manufacturing – Lox Dome AN 16 -AN 8 Fitting (1” to ½”) 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 97

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

Fluid Interface Rocket Interface • AN 16( 1”) fittings for Propene and Lox – Vfuel=16. 8 ft/s (Before Split) – Vfuel=14. 7 ft/s (After Split) – VLOX=20. 2 ft/s 3/9/2005 Component Interface • Six ½” to 1/4” (8 AN 4 NPT) fittings in the injector fuel (Reduced Further in Injector Fuel Manifold) • One 1”- ½” (AN 16 -AN 8) fitting in the Lox Dome for Lox Thrust Chamber Assembly Preliminary Design Review 99

Engine Masses Part 3/9/2005 weight (lb) Lox Dome 7. 40 Deflector 0. 17 Lox Plate 3. 00 Injector 7. 00 Chamber 22. 80 Nozzle 22. 20 Total 62. 57 Thrust Chamber Assembly Preliminary Design Review 100

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

Manufacturing Timeline Part Process days → 1 2 3 4 5 6 7 8 9 1 0 1 1 1 2 1 3 1 4 1 5 1 6 1 7 Lox Dome shipping manufacturing Deflector shipping manufacturing brazing Lox Plate shipping manufacturing brazing Injector shipping manufacturing brazing Chamber shipping manufacturing Nozzle shipping manufacturing 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 102

Price List – Materials Description Material price Lox Dome 2" x 7 1/2" OD c 101 370. 30 Deflector 2" x 2" OD c 101 51. 48 Lox Plate 1/4" x 7 1/2" OD c 101 90. 34 Injector 3/4" x 7 1/2" OD c 101 169. 60 Flange 1/2" x 7 1/2" OD carbon steel 38. 69 Chamber 40 -5 black pipe (21 ft) carbon steel 266. 85 Nozzle Extension 4 1/4" x 8 1/4" OD carbon steel 109. 42 Nozzle Throat 5" x 6" OD carbon steel 38. 44 Total 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 1135. 12 103

Price List – Manufacturing Description quantity price Total Cost Drill Bit #68 (. 0310") 20 1. 84 36. 80 Drill Bit #48 (. 0760") 10 3. 83 38. 30 Drill Bit 5/64" (. 0781") 10 2. 85 28. 50 Drill Bit #47 (. 0785") 10 2. 72 27. 20 3/9/2005 Total Thrust Chamber Assembly Preliminary Design Review 130. 80 104

Price List – Parts Description quantity price Total Cost Bolts 1/4 -20, 2 -1/2" Grade 8 (25) 1 9. 28 Nuts 1/4 -20, ASTM Standards (25) 1 9. 87 Washers 1/4" SAE Standards (50) 1 5. 01 Gasket Graphite 1/16" thick 24"x 24" 1 40. 53 O-ring 339 Teflon, 3 -1/4 ID 1 21. 89 3/9/2005 Total Thrust Chamber Assembly Preliminary Design Review 86. 58 105

Costs Description Materials Manufacturing Parts 3/9/2005 1135. 12 130. 80 86. 58 Brazing Total price 350. 00 1702. 50 Thrust Chamber Assembly Preliminary Design Review 106

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

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) 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 108

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 109

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

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

Numbers Summary O/F 2. 2729 Pc 300 psi F 2200 lbf ε 8 Tc 6340 R c* 6044 ft/s Pe 5. 66 psi ve 9628 ft/s Cf)vac 1. 593 γ 1. 1398 MW 21. 313 lb/lbmole Ivac 327. 6 s Iopt 299. 2 s L* 42. 5 in εc 2 20. 41 1. 44207 0. 95 257. 35 s 8280 ft/s Pc/Pa Cf η Isp c 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 112

Numbers mdot 8. 4519 lb/s mo 5. 8695 lb/s mf 2. 5824 lb/s Dc 3. 692 in Ac 10. 706 in^2 Dt 2. 6107 in At 5. 353 in^2 De 7. 3841 in Ae 42. 82 in^2 Lc 20. 72 in 3/9/2005 Chamber Throat Exit Length of chamber Thrust Chamber Assembly Preliminary Design Review 113

Numbers ΔP 60 psi ρ lox 0. 0412 lb/in^3 ρ fuel 0. 0218 lb/in^3 0. 8 discharge coeff 0. 031 in film orifices 0. 0781 in outside orifices Dox, in 0. 076 in inside orifices Dfuel 0. 0785 in fuel orifices Aox 0. 167887 in^2 Afuel 0. 100703 Cd Dfilm Dox, out injector pressure drop density lox in^2 Vox 1060 in/s 88. 33333 ft/s Vfuel 1445 in/s 120. 4167 ft/s mom_ox_o 222 lb-in/s mom_ox_in 210 lb-in/s mom_f 224 lb-in/s 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 114

Numbers ac 49259 in/s f 1 t 7821 Hz f 2 t 12974 f 3 t 4104. 9 1 7 ft/s Hz 17845 Hz f 1 r 16276 Hz f 2 r 29800 Hz Df/Vf 5. 40 E-05 s 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 115

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

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

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

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 119

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

Design Process • 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) 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 121

Design Process • With the information available we run the NASA thermochemistry code to obtain some useful data: – – – – 3/9/2005 Chamber Temp (Tc) = 6374 R C* = 6073 ft/s Exit pressure (pe) = 5. 7 psi Exit velocity (ve) = 9673. 5 ft/s Cfvac = 1. 593 Specific heat ratio γ = 1. 1396 Molecular weight = 21. 232 Ispvac = 329. 3 s Thrust Chamber Assembly Preliminary Design Review 122

Design Process 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 Cf 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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 123

Design Process 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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 124

Design Process 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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 125

Design Process • • 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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 126

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 127

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 128

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 129

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 130

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

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 132

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

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

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

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

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 137

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 138

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

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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 140

Current Concept Summary • Injector: Flat Face Injector • Chamber: Ablative Lining • Nozzle: 80% Bell (Conical Shown) Injector Combustion Chamber 20° 5” 2” 3/9/2005 15° 2. 62” 9. 44” 3. 61” Thrust Chamber Assembly Preliminary Design Review = 8 8. 95” 141

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. 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 142

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

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

Documentation Control • 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. 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 145

Specified Design Guidelines 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, 000 ft • 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) 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 146

Design Objectives • Design a liquid propellant thrust chamber assembly to be launched as a flight demonstration engine for Cal. State 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 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 147

Design Objectives Level 2 • • • Thrust ~ 2200 lbf (Sea Level) Chamber Pressure ~ 300 psi Propellants: LOX-propylene Mixture ratio (O/F) = 2. 27 Burn time ~ 15 s Engine mass ~ 15 lbm 3/9/2005 Thrust Chamber Assembly Preliminary Design Review 148