Propulsion Systems Design
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1 Propulsion Systems Design Rocket engine basics Survey of the technologies Propellant feed systems Propulsion systems design David L. Akin - All rights reserved
2 Propulsion Taxonomy Mass Expulsion Non-Mass Expulsion Thermal Non-Thermal Chemical Non-Chemical Ion Solar Sail Monopropellants Bipropellants Nuclear Electrical MPD Beamed Laser Sail Microwave Sail Solar Cold Gas MagnetoPlasma Solids Hybrids Liquids Air-Breathing ED Tether Pressure-Fed Pump-Fed 2
3 Thermal Rocket Exhaust Velocity Exhaust velocity is where V e = 2γ γ 1 RT 0 M 1 p e p 0 γ 1 γ M average molecular weight of exhaust R universal gas const.= Joules mole K γ ratio of specific heats 1.2 3
4 Ideal Thermal Rocket Exhaust Velocity Ideal exhaust velocity is 2γ RT V e = 0 γ 1 M This corresponds to an ideally expanded nozzle All thermal energy converted to kinetic energy of exhaust Only a function of temperature and molecular weight! 4
5 Thermal Rocket Performance Thrust is T = m V e + ( p e p amb )A e Effective exhaust velocity T = m c c = V e + p e p amb Expansion ratio ( ) A e m I sp = c g 0 A t = γ +1 A e 2 1 γ 1 p e p 0 1 γ γ +1 γ 1 1 p e p 0 γ 1 γ 5
6 A Word About Specific Impulse Defined as thrust/propellant used English units: lbs thrust/(lbs prop/sec)=sec Metric units: N thrust/(kg prop/sec)=m/sec Two ways to regard discrepancy - lbs is not mass in English units - should be slugs Isp = thrust/weight flow rate of propellant If the real intent of specific impulse is I sp = Ṫ m and T = ṁv e then I sp = V e!!! 6
7 Nozzle Design Pressure ratio p 0 /p e =100 (1470 psi-->14.7 psi) A e /A t =11.9 Pressure ratio p 0 /p e =1000 (1470 psi-->1.47 psi) A e /A t =71.6 Difference between sea level and ideal vacuum V e γ 1 V e = 1 p γ e V e,ideal p 0 I sp,vacuum =455 sec --> I sp,sl =333 sec 7
8 Solid Rocket Motor From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
9 Solid Propellant Combustion Characteristics From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
10 Solid Grain Configurations From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
11 Short-Grain Solid Configurations From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
12 Advanced Grain Configurations From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
13 Liquid Rocket Engine 13
14 Liquid Propellant Feed Systems 14
15 Space Shuttle OMS Engine From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
16 Turbopump Fed Liquid Rocket Engine From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
17 Sample Pump-fed Engine Cycles From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
18 Gas Generator Cycle Engine From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
19 SSME Schematic 19
20 SSME Powerhead Configuration 20
21 SSME Engine Cycle From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
22 Liquid Rocket Engine Cutaway From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
23 H-1 Engine Injector Plate 23
24 Injector Concepts From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
25 TR-201 Engine (LM Descent/Delta) 25
26 Solid Rocket Nozzle (Heat-Sink) From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
27 Ablative Nozzle Schematic From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
28 Active Chamber Cooling Schematic From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
29 Boundary Layer Cooling Approaches From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
30 Hybrid Rocket Schematic From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
31 Hybrid Rocket Combustion From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
32 Thrust Vector Control Approaches From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
33 Gemini Entry Reaction Control System 33
34 Apollo Reaction Control System 34
35 Apollo CSM RCS Assembly 35
36 Lunar Module Reaction Control System 36
37 LM RCS Quad 37
38 Viking Aeroshell RCS Thruster 38
39 Space Shuttle Primary RCS Engine From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
40 Monopropellant Engine Design From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
41 Cold Gas Thruster Exhaust Velocity Assume nitrogen gas thrusters V e = 2γ T 0 1 γ 1 M pe p o γ 1 γ M = 28 p 0 = 300 psi V e = T 0 = 300 K = (1.4) (300) 28 1 p e =2psi γ = = 689 m sec 41
42 Cold-gas Propellant Performance From G. P. Sutton, Elements (5th ed.) John Wiley and Sons,
43 Total Impulse Total impulse I t is the total thrust-time product for the propulsion system, with units <N-sec> I t = Tt = ṁv e t t = ρv ṁ To assess cold-gas systems, we can examine total impulse per unit volume of propellant storage 43 I t = ρv v e I t V = ρv e
44 Performance of Cold-Gas Systems 44
45 Self-Pressurizing Propellants (CO 2 ) 45
46 Self-Pressurizing Propellants (N 2 O) Density 1300 kg/m 3 Density 625 kg/m 3 46
47 N 2 O Performance Augmentation Nominal cold-gas exhaust velocity ~600 m/sec N 2 O dissociates in the presence of a heated catalyst 2N 2 O 2N 2 + O 2 engine temperature ~1300 C exhaust velocity ~1800 m/sec NOFB (Nitrous Oxide Fuel Blend) - store premixed N 2 O/hydrocarbon mixture exhaust velocity >3000 m/sec 47
48 Pressurization System Analysis P g0, V g Adiabatic Expansion of Pressurizing Gas P gf, V g p g,0 V g γ = p g, f V g γ + p V γ Known quantities: P L, V L Initial P L, V L Final P g,0 =Initial gas pressure P g,f =Final gas pressure P L =Operating pressure of propellant tank(s) V L =Volume of propellant tank(s) Solve for gas volume V g 48
49 Boost Module Propellant Tanks Gross mass 23,000 kg Inert mass 2300 kg Propellant mass 20,700 kg Mixture ratio N 2 O 4 /A50 = 1.8 (by mass) N 2 O 4 tank Mass = 13,310 kg Density = 1450 kg/m 3 Volume = m 3 --> r sphere =1.299 m Aerozine 50 tank Mass = 7390 kg Density = 900 kg/m 3 Volume = m 3 --> r sphere =1.252 m 49
50 Boost Module Main Propulsion Total propellant volume V L = m 3 Assume engine pressure p 0 = 250 psi Tank pressure p L = 1.25*p 0 = 312 psi Final GHe pressure p g,f = 75 psi + p L = 388 psi Initial GHe pressure p g,0 = 4500 psi Conversion factor 1 psi = 6892 Pa Ratio of specific heats for He = 1.67 ( 4500 psi)v 1.67 g = ( 388 psi)v 1.67 g psi V g = m 3 Ideal gas: T=300 K --> ρ=49.7 kg/m 3 (4500 psi = MPa) M He =185.1 kg 50 ( ) 1.67 ( ) m 3 ρ He = p M g,0 RT 0
51 Nuclear Thermal Rockets Heat propellants by passing through nuclear reactor Isp limited by temperature limits on reactor elements (~900 sec for H2 propellant) Mass impacts of reactor, shielding High thrust system 51
52 VASIMR Engine Concept 52
53 Ion Propulsion Uses electrostatic forces to accelerate ions Injects electrons to keep beam neutral High Isp (~3000 sec) at low thrust (~10 N) Substantial mass penalty for electrical power generation 53
54 Solar Sails Sunlight reflecting off sail produces momentum transfer T = 2m V = 2m c E = mc 2 m = E m = E 1 = P c 2 t c 2 c 2 At 1 AU, P=1394 W/m 2 c=3x10 8 m/sec T=9x10-6 N/m 2 54
55 Propulsion Taxonomy Mass Expulsion Non-Mass Expulsion Thermal Non-Thermal Chemical Non-Chemical Ion Solar Sail Monopropellants Bipropellants Nuclear Electrical MPD Beamed Laser Sail Microwave Sail Solar Cold Gas MagnetoPlasma Solids Hybrids Liquids Air-Breathing ED Tether Pressure-Fed Pump-Fed 55
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