Propulsion Systems Design MARYLAND. Rocket engine basics Survey of the technologies Propellant feed systems Propulsion systems design

Transcription

1 Propulsion Systems Design Rocket engine basics Survey of the technologies Propellant feed systems Propulsion systems design 2008 David L. Akin - All rights reserved 1

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 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 3

4 Thermal Rocket Exhaust Velocity Exhaust velocity is V e = 2γ RT o γ 1 M [ 1 ( pe p o ) γ 1 γ ] where M average molecular weight of exhaust R universal gas constant = Joules mole o K γ ratio of specific heats 1.2 p e pressure at nozzle exit plane p o combustion chamber pressure 4

5 Ideal Thermal Rocket Exhaust Velocity Ideal exhaust velocity is 2γ RT o V e = 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! 5

6 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 t = γ +1 A e 2 1 γ 1 p e p 0 ( ) A e m 1 γ γ +1 γ 1 1 p e p 0 γ 1 γ I sp = c g 0 6

7 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!!! 7

8 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 V e = 1 p e V e,ideal p 0 γ 1 γ I sp,vacuum =455 sec --> I sp,sl =333 sec 8

9 Solid Rocket Motor From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

10 Solid Propellant Combusion From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

11 Solid Grain Configurations From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

12 Short-Grain Solid Configurations From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

13 Advanced Grain Configurations From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

14 Liquid Rocket Engine 14

15 Liquid Propellant Feed Systems 15

16 Space Shuttle OMS Engine From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

17 Turbopump Fed Liquid Rocket Engine From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

18 Sample Pump-fed Engine Cycles From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

19 Gas Generator Cycle Engine From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

20 SSME Engine Cycle From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

21 Liquid Rocket Engine Cutaway From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

22 H-1 Engine Injector Plate 22

23 Injector Concepts From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

24 Solid Rocket Nozzle (Heat-Sink) From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

25 Ablative Nozzle Schematic From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

26 Active Chamber Cooling Schematic From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

27 Boundary Layer Cooling Approaches From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

28 Hybrid Rocket Schematic From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

29 Hybrid Rocket Combustion From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

30 Thrust Vector Control Approaches From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

31 Apollo Reaction Control System Thrusters 31

32 Space Shuttle Primary RCS Engine From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

33 Monopropellant Engine Design From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

34 Cold-gas Propellant Performance From G. P. Sutton, Rocket Propulsion Elements (5th ed.) John Wiley and Sons,

35 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 l V l 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 35

36 Boost Module Propellant Tanks Low-Cost Return to the Moon 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 Space Systems Laboratory University of Maryland

37 Boost Module Main Propulsion Low-Cost Return to the Moon 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 (300 psi = MPa) M He =185.1 kg Space Systems Laboratory University of Maryland ( ) 1.67 ( ) m 3 ρ He = p M g,0 RT 0

38 Air-Breathing Propulsion Much of the total energy required is to accelerate propellants for the last phases of launch Typical (LOX/LH2) propellants are 86% oxygen by mass (I sp = 450 sec I sp,fuel = 3150 sec) Logical conclusion: take in oxygen from ambient air during early portion of launch trajectory 38

39 Hypersonic Research Engine 39

40 X-43A Hypersonic Flight Test 40

41 Scramjet-Airframe Integration NASA Hyper-X Program Demonstrates Scramjet Technologies NASA Facts FS LaRC 41

42 X-43A Film Clip 42

43 Performance of Airbreathing Engines 43

44 Constraints on Airbreathing Trajectories 44

45 Design Trends of Air-Breathing Propulsion C. Trefny, An Air-Breathing Launch Vehicle Concept for Single-Stage-to-Orbit AIAA

46 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 46

47 Speculative Designs Including Nuclear Aerojet General, Payload, Cost, and Reliability Analysis of Saturn C-5 and Nova with NERVA or Chemical Third Stages AGC-2279, June