Transcript Part 1
CVA Summer School 2010 - Roma July 8, 2011
LIQUID ROCKET PROPULSION
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Christophe ROTHMUND SNECMA – Space Engines Division [email protected]
CONTENTS
Part 1 : LIQUID ROCKET ENGINES Part 2 : NEAR TERM PROPULSION SYSTEMS Part 3 : ELECTRIC PROPULSION Part 1 : LIQUID ROCKET ENGINES Part 2 : ANATOMY OF A ROCKET ENGINE Part 5 : ADVANCED PROPULSION Part 4 : NUCLEAR PROPULSION
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Part 1
LIQUID ROCKET ENGINES
LIQUID ROCKET ENGINE TYPES
4 NUCLEAR THERMAL
FUEL NUCLEAR REACTOR
SOLID GRAIN
LIQUID PROPELLANTS
MONOPROPELLANTS :
Hydrazine, Hydrogen peroxyde
BIPROPELLANTS :
FUELS : Kerosine, ethanol Liquid hydrogen, UDMH, MMH, Hydrazine
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OXIDIZERS : Liquid oxygen, N2O4, Hydrogen peroxyde
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).
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CHEMICAL REACTIONS
BIPROPELLANTS
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) CH2 + 1.5O2
→ 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)
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Specific impulse of various propulsion technologies Engine type Jet engine Solid rocket Bipropellant rocket Ion thruster VASIMR Specific impulse 300 s 250 s 450 s 3000 s 30000 s
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PROPULSION BASICS
Pressure-fed cycle
Pros :
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Simplicity
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Low complexity
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Cons :
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Low performance
-Oldest and simplest cycle, -Rarely used nowadays on launch vehicles, Powered France’s Launch vehicle family Diamant.
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Gas-generator cycle
Pros :
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Higher presssures
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Lowerturbine temperatures
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Highest performance
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Cons :
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Moderate performance
-Most widely used cycle in the western world,
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Preburner cycle
Pros :
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Higher presssures
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Lowerturbine temperatures
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Highest performance
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Cons :
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High complexity
-Used on the Shuttle and the H-2A main engine
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Expander cycle
Pros :
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Thermally challenging
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Highest performance
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Cons :
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For cryogenic engines only
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Limited power therefore not suited for high thrusts
-Oldest cryogenic engine in service (RL-10) -Used in Japan and Europe
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Full flow staged combustion rocket cycle
Pros :
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Higher presssures
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Lowerturbine temperatures
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Highest performance
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Cons :
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Very high complexity
-Never flown, -Demonstration only (IPD) so far
Hybrid rocket cycle Pros :
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Higher performance than solids
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Lower complexity than liquids
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Cons :
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Lower performance than liquids
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Higher complexity than solids
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Nuclear thermal rocket cycle
Pros :
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Highest performance
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Cons :
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Very high complexity
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Radiations
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Never flown,
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Demonstration only so far
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NOZZLES
THRUST CHAMBER ASSEMBLY
Chamber Characteristics: • Combustion • High pressure • High temperature • Very low net fluid velocity GIMBAL INJECTOR COMBUSTION CHAMBER Exit Characteristics: • Flow expands to fill enlarged volume • Reduced pressure • Reduced temperature • Very high fluid velocity NOZZLE EXTENSION
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Thrust chamber cooling methods
Nozzle design challenges Structural factors
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Physical
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Weight, Center of Gravity Service Life
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Running time, Number of starts Duty Cycle/Operating Range
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Continuous operation at one power level
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Throttled operation Materials
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Compatibility
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Strength, Heat transfer capability Structural concerns
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Loads
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Testing
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Altitude Simulation or Sea Level
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Sideloads
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Flight
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Lift off, shut down, restart
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Sideloads
• •
Transportation Temperature
19 June 16, 1997
KEY REQUIREMENTS FOR SAFE NOZZLE OPERATION
Sea level :
- stable operation on ground
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high performance
Vacuum :
- high vacuum performance - low package volume
Bell nozzle
Advanced concepts with altitude adaptation
Dual-bell nozzle
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Extendible nozzle Plug nozzle (“Aerospike”)
• • •
Nozzle Types Conical Nozzle
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Simple cone shape - easy to fabricate
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Rarely used on modern rockets Bell Nozzle
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Bell shape reduces divergence loss over a similar length conical nozzle
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Allows shorter nozzles to be used Annular Nozzles (spike or pug)
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Altitude compensating nozzle
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Aerospike is a spike nozzle that uses a secondary gas bleed to “fill out” the truncated portion of the spike nozzle
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Nozzle extension cooling systems Laser welded jacket to liner Brazed tubes inside jacket Brazed jacket to liner Materials : stainless steel, copper, nickel, copper/nickel
AEROSPIKE NOZZLES
Linear aerospike Rocketdyne XRS-2200 Rocketdyne AMPS-1 1960’s Annular aerospike
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World’s first aerospike flight, September 20, 2003 (California State Univ. Long Beach)
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Part 2
ANATOMY OF A ROCKET ENGINE
1 – Gas generator cycle : F-1 and Vulcain engines 2 – Staged combustion cycle : SSME 3 – Expander cycle : Vinci 4 – Linear aerospike XRS2200 5 – Nuclear : NERVA/RIFT
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Rocketdyne F-1
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Engine flow diagram Rocketdyne F-1
Engine characteristics Rocketdyne F-1
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Thrust (sea level): Burn time: Specific impulse: Engine weight dry: Engine weight burnout: Height: Diameter: Exit to throat ratio: Propellants: Mixture ratio:
Apollo 4, 6, and 8
6.67 MN 150 s 260 s 8.353 t 9.115 t 5.79 m 3.76 m 16 to 1 LOX & RP-1 2.27:1 oxidizer to fuel
Apollo 9 on
6.77 MN 165s 263 s 8.391 t 9.153 t
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Engine components Rocketdyne F-1
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TESTS AND LAUNCHES
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VULCAIN 2
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VULCAIN 1 TO VULCAIN 2 EVOLUTION
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Engine characteristics Vacuum thrust Mixture ratio Vacuum impulse Dry weight specific Chamber pressure Expansion ratio Design life Vulcain 1 1140 kN 4,9 to 5,3 430 s 1680 kg 110 bar 45 6000 s 20 starts Vulcain 2 1350 kN 6,13 434 s 2040 kg 116 bar 60 5400 s 20 starts VULCAIN 2
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Vulcain 2 Flow diagram VULCAIN 2
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Thrust chamber VULCAIN 2
Hydrogen turbopump VULCAIN 2
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34200 rpm, 11,3 MW
Oxygen turbopump VULCAIN 2
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13300 rpm, 2,9 MW
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Space Shuttle Main Engine
Propellants Vacuum thrust at 109 % Mixture ratio Vacuum specific impulse Dry weight Chamber pressure Expansion ratio Throttle range Design life
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Engine characteristics SSME LOX – LH2 2279 kN 6 452 s 3527 kg 206,4 bar 69 67 % - 109 % 7,5 hours 55 starts SSME
2,07 bar SSME Flow schematic 6,9 bar 190,3 bar 511,6 bar SSME 29,1 bar 206,4 bar 296,5 bar 450 bar
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Hydrogen
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SSME Powerhead
Fuel Preburner Oxidizer Preburner
SSME
Main chamber
Oxygen
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F1 vs. SSME test comparison SSME Last Planned SSME Test : July 29, 2009 Duration : 520 sec, NASA Stennis
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VINCI
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FLOW DIAGRAM VINCI
Engine characteristics VINCI
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Propellants Vacuum thrust Mixture ratio Vacuum specific impulse Chamber pressure Expansion ratio Vinci LOX – LH2 180 kN 5,80 465 s 60 bar 240
VINCI Hydrogen turbopump
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Oxygen turbopump VINCI
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Combustion chamber VINCI 122 co-axial injection elements 228 cooling channels
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Nozzle extension VINCI
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Vinci test facility with altitude simulation
TYPICAL SMALL ENGINES
MONOPROPELLANT BIPROPELLANT
Typical applications : Attitude control Roll control Soft landings (Moon, Mars, …)
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SPACESHIPONE HYBRID ENGINE
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AMROC (USA)
LIQUID ROCKET ENGINE TEST FACILITIES
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Major elements of a test program
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PF52 Vulcain powerpack test facility
Vulcain 1 – hydrogen TP test
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Vinci test facility with altitude simulation
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PF50 and P5 Vulcain engine test stands LOX Tank Engine
NASA A-3 J-2X altitude simulation engine test stand
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Ø 7 m
simulated altitude : approximately 30 km
BEWARE OF HYDROGEN IMPURITIES… - 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, ...)
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HYBRID ENGINE TEST FACILITY
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Part 4
NUCLEAR PROPULSION
INTEREST OF NUCLEAR THERMAL PROPULSION Energy !
→1 kg of fissionable material (U235) contains 10 000 000 times more energy, as 1 kg of chemical fuel.
Consequences :
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Higher specific impulse - higher useful load fraction - No oxidizer required !
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Nuclear - thermal propulsion
Size of 1972 Vintage 350 kN NERVA Engine
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Isp = 888.3 s
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F = 333.6 kN
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Tc = 2700 K
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m = 5.7 t
American nuclear rockets of the 1960’s Engine and reactor in Engine Test Stand One (Nevada Test Site)
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KIWI A 1958-60 100 MW 1.125 kN KIWI B 1961-64 1000 MW 11.25 kN Phoebus 1 Phoebus 2 1965-66 1000 MW 11.25 kN 1967 50000 MW 56.25 kN
UPPER STAGE COMPARISON
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Departure mass Dry mass Thrust Duration Isp Payload S-IV B 121.2 t 12.2 t 91 kN 475 s 4180 m/s 39 t NERVA 53.694 t 12.429 t 266.8 kN 1250 s 7840 m/s 54.5 t For Saturn V
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NERVA ultimate goal : manned Mars flight
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