Presentation is loading. Please wait.

Presentation is loading. Please wait.

1 LIQUID ROCKET PROPULSION CVA Summer School 2010 - Roma Christophe ROTHMUND SNECMA – Space Engines Division July 8, 2011.

Similar presentations


Presentation on theme: "1 LIQUID ROCKET PROPULSION CVA Summer School 2010 - Roma Christophe ROTHMUND SNECMA – Space Engines Division July 8, 2011."— Presentation transcript:

1 1 LIQUID ROCKET PROPULSION CVA Summer School Roma Christophe ROTHMUND SNECMA – Space Engines Division July 8, 2011

2 2 Part 1 : LIQUID ROCKET ENGINES Part 2 : NEAR TERM PROPULSION SYSTEMS Part 3 : ELECTRIC PROPULSION Part 4 : ANATOMY OF A ROCKET ENGINE Part 5 : ADVANCED PROPULSION Part 1 : LIQUID ROCKET ENGINES Part 2 : ANATOMY OF A ROCKET ENGINE Part 3 : TEST FACILITIES Part 4 : NUCLEAR PROPULSION CONTENTS

3 3 Part 1 LIQUID ROCKET ENGINES

4 4 LIQUID ROCKET ENGINE TYPES NUCLEAR THERMAL FUEL NUCLEAR REACTOR SOLID GRAIN

5 5 LIQUID PROPELLANTS MONOPROPELLANTS : BIPROPELLANTS : Hydrazine, Hydrogen peroxyde FUELS : OXIDIZERS : Kerosine, ethanol Liquid hydrogen, UDMH, MMH, Hydrazine Liquid oxygen, N2O4, Hydrogen peroxyde

6 6 HYBRID PROPELLANTS liquid propellant : oxidizer - solid propellant : fuel (solid oxidizers are problematic and lower performing than liquid oxidizers) oxidizers : gaseous or liquid oxygen nitrous oxide. fuels : polymers (e.g.polyethylene), cross-linked rubber (e.g.HTPB), liquefying fuels (e.g. paraffin). Solid fuels (HTPB or paraffin) allow for the incorporation of high- energy fuel additives (e.g.aluminium).

7 7 Dimethylhydrazine (UDMH) : H2N – N (CH3)2 (l)+ 2 N2O4 (l) - > 3 N2 (g) + 4 H2O (g) + 2 CO2 (g) Hydrazine hydrate (N2H4,H2O) : 2 (N2H4,H2O) (l) + N2O4 (l) - > 3 N2 (g) + 6 H2O (g) Monomethylhydrazine (MMH) : 4 H2N – NHCH3 (l) + 5 N2O4 (l) - > 9 N2 (g) + 12 H2O (g) + 4 CO2 (g) Kerosene (CH2 is the approximate formula ) with hydrogen peroxide : CH2 + 3H2O2 → CO2 + 4H2O Kerosene and liquid oxygen (LOX) CH O2 → CO2 + H2O Hydrogen and oxygen (liquids) : 2 H2 (g)+ O2 (g) - > 2 H2O (g) MONOPROPELLANTS Hydrogen peroxyde (H2O2) H2O2 (l) - > H2O (l) + 1/2 O2 (g) Hydrazine (N2H4) N2H4 (l) - > N2 (g) + 2 H2 (g) CHEMICAL REACTIONS BIPROPELLANTS

8 8 Engine typeSpecific impulse Jet engine300 s Solid rocket250 s Bipropellant rocket 450 s Ion thruster3000 s VASIMR30000 s Specific impulse of various propulsion technologies

9 9 PROPULSION BASICS Pressure-fed cycle Pros : -Simplicity -Low complexity -Cons : -Low performance -Oldest and simplest cycle, -Rarely used nowadays on launch vehicles, -Powered France’s Launch vehicle family Diamant.

10 10 Pros : -Higher presssures -Lowerturbine temperatures -Highest performance -Cons : -Moderate performance -Most widely used cycle in the western world, Gas-generator cycle

11 11 Pros : -Higher presssures -Lowerturbine temperatures -Highest performance -Cons : -High complexity -Used on the Shuttle and the H-2A main engine Preburner cycle

12 12 Pros : -Thermally challenging -Highest performance -Cons : -For cryogenic engines only -Limited power therefore not suited for high thrusts -Oldest cryogenic engine in service (RL-10) -Used in Japan and Europe Expander cycle

13 13 Pros : -Higher presssures -Lowerturbine temperatures -Highest performance -Cons : -Very high complexity -Never flown, -Demonstration only (IPD) so far Full flow staged combustion rocket cycle

14 14 Hybrid rocket cycle Pros : - Higher performance than solids - Lower complexity than liquids Cons : - Lower performance than liquids - Higher complexity than solids

15 15 Nuclear thermal rocket cycle Pros : -Highest performance -Cons : -Very high complexity -Radiations -Never flown, -Demonstration only so far

16 16 NOZZLES

17 17 Chamber Characteristics: Combustion High pressure High temperature Very low net fluid velocity Exit Characteristics: Flow expands to fill enlarged volume Reduced pressure Reduced temperature Very high fluid velocity THRUST CHAMBER ASSEMBLY NOZZLE EXTENSION COMBUSTION CHAMBER INJECTOR GIMBAL

18 18 Thrust chamber cooling methods

19 19 June 16, 1997 Physical –Weight, Center of Gravity Service Life –Running time, Number of starts Duty Cycle/Operating Range –Continuous operation at one power level –Throttled operation Materials –Compatibility –Strength, –Heat transfer capability Structural factors Loads Testing –Altitude Simulation or Sea Level –Sideloads Flight –Lift off, shut down, restart –Sideloads Transportation Temperature Structural concerns Nozzle design challenges

20 20 Sea level : - stable operation on ground -high performance Vacuum : - high vacuum performance - low package volume KEY REQUIREMENTS FOR SAFE NOZZLE OPERATION Bell nozzle Dual-bell nozzle Plug nozzle (“Aerospike”) Extendible nozzle Advanced concepts with altitude adaptation

21 21 Nozzle Types Conical Nozzle –Simple cone shape - easy to fabricate –Rarely used on modern rockets Bell Nozzle –Bell shape reduces divergence loss over a similar length conical nozzle –Allows shorter nozzles to be used Annular Nozzles (spike or pug) –Altitude compensating nozzle –Aerospike is a spike nozzle that uses a secondary gas bleed to “fill out” the truncated portion of the spike nozzle

22 22 Brazed jacket to liner Laser welded jacket to liner Brazed tubes inside jacket Nozzle extension cooling systems Materials : stainless steel, copper, nickel, copper/nickel

23 23 AEROSPIKE NOZZLES Annular aerospike Linear aerospike Rocketdyne AMPS ’s Rocketdyne XRS-2200 World’s first aerospike flight, September 20, 2003 (California State Univ. Long Beach)

24 24 1 – Gas generator cycle : F-1 and Vulcain engines 2 – Staged combustion cycle : SSME 3 – Expander cycle : Vinci 4 – Linear aerospike XRS – Nuclear : NERVA/RIFT Part 2 ANATOMY OF A ROCKET ENGINE

25 25 Rocketdyne F-1

26 26 Engine flow diagram Rocketdyne F-1

27 27 Apollo 4, 6, and 8 Apollo 9 on Thrust (sea level):6.67 MN6.77 MN Burn time:150 s165s Specific impulse:260 s263 s Engine weight dry:8.353 t8.391 t Engine weight burnout: t9.153 t Height:5.79 m Diameter:3.76 m Exit to throat ratio:16 to 1 Propellants:LOX & RP-1 Mixture ratio:2.27:1 oxidizer to fuel Engine characteristics Rocketdyne F-1

28 28 Engine components Rocketdyne F-1

29 29 TESTS AND LAUNCHES 100 m 1000 m

30 30 VULCAIN 2

31 31 VULCAIN 1 TO VULCAIN 2 EVOLUTION

32 32 Vulcain 1Vulcain 2 Vacuum thrust1140 kN1350 kN Mixture ratio4,9 to 5,36,13 Vacuum specific impulse 430 s434 s Dry weight1680 kg2040 kg Chamber pressure110 bar116 bar Expansion ratio4560 Design life6000 s 20 starts 5400 s 20 starts Engine characteristics VULCAIN 2

33 33 Vulcain 2 Flow diagram VULCAIN 2

34 34 Thrust chamber VULCAIN 2

35 rpm, 11,3 MW Hydrogen turbopump VULCAIN 2

36 36 Oxygen turbopump rpm, 2,9 MW VULCAIN 2

37 37 Space Shuttle Main Engine

38 38 SSME PropellantsLOX – LH2 Vacuum thrust at 109 %2279 kN Mixture ratio6 Vacuum specific impulse452 s Dry weight3527 kg Chamber pressure206,4 bar Expansion ratio69 Throttle range67 % % Design life7,5 hours 55 starts Engine characteristics SSME

39 39 SSME Flow schematic 511,6 bar 296,5 bar 29,1 bar 6,9 bar 2,07 bar 190,3 bar 450 bar 206,4 bar SSME

40 40 SSME Powerhead Hydrogen Oxygen Main chamber Oxidizer Preburner Fuel Preburner SSME

41 41 F1 vs. SSME test comparison Last Planned SSME Test : July 29, 2009 Duration : 520 sec, NASA Stennis SSME

42 42 VINCI

43 43 FLOW DIAGRAM VINCI

44 44 Vinci PropellantsLOX – LH2 Vacuum thrust180 kN Mixture ratio5,80 Vacuum specific impulse465 s Chamber pressure60 bar Expansion ratio240 Engine characteristics VINCI

45 45 Hydrogen turbopump VINCI

46 46 Oxygen turbopump VINCI

47 47 Combustion chamber 228 cooling channels 122 co-axial injection elements VINCI

48 48 Nozzle extension VINCI

49 49 Vinci test facility with altitude simulation

50 50 TYPICAL SMALL ENGINES MONOPROPELLANT BIPROPELLANT Typical applications :Attitude control Roll control Soft landings (Moon, Mars, …)

51 51 SPACESHIPONE HYBRID ENGINE

52 52 AMROC (USA)

53 53 LIQUID ROCKET ENGINE TEST FACILITIES

54 54 Major elements of a test program

55 55 PF52 Vulcain powerpack test facility Vulcain 1 – hydrogen TP test

56 56 Vinci test facility with altitude simulation

57 57 PF50 and P5 Vulcain engine test stands Engine LOX Tank

58 58 NASA A-3 J-2X altitude simulation engine test stand simulated altitude : approximately 30 km 59 m Ø 7 m

59 59 - Avoid risky configurations (e.g. low points,...) - Define and integrate the cleansing constraints - Easily accessible equipment after commissioning, - Rapid reaction instrumentation (e.g. fast pressure sensors,...) BEWARE OF HYDROGEN IMPURITIES…

60 60 HYBRID ENGINE TEST FACILITY

61 61 Part 4 NUCLEAR PROPULSION

62 62 INTEREST OF NUCLEAR THERMAL PROPULSION Energy ! →1 kg of fissionable material (U235) contains times more energy, as 1 kg of chemical fuel. Consequences : -Higher specific impulse - higher useful load fraction - No oxidizer required !

63 63 Isp = s F = kN Tc = 2700 K m = 5.7 t Nuclear - thermal propulsion Size of 1972 Vintage 350 kN NERVA Engine

64 64 Photo courtesy of National Nuclear Security Administration / Nevada Site Office Engine and reactor in Engine Test Stand One (Nevada Test Site) KIWI AKIWI BPhoebus 1Phoebus MW1000 MW MW kN11.25 kN kN American nuclear rockets of the 1960’s

65 65 S-IV BNERVA Departure mass121.2 t t Dry mass12.2 t t Thrust91 kN266.8 kN Duration475 s1250 s Isp4180 m/s7840 m/s Payload39 t54.5 t UPPER STAGE COMPARISON For Saturn V

66 66 NERVA ultimate goal : manned Mars flight

67 67 ANY QUESTIONS ?


Download ppt "1 LIQUID ROCKET PROPULSION CVA Summer School 2010 - Roma Christophe ROTHMUND SNECMA – Space Engines Division July 8, 2011."

Similar presentations


Ads by Google