AAE 439
Ch6 –1
6. LIQUID ROCKET PROPULSION (LRP)SYSTEMS
6. LIQUID ROCKET PROPULSION (LRP)SYSTEMS
AAE 439
Ch6 –2
6.1 INTRODUCTION6.1 INTRODUCTION
AAE 439
Ch6 –3
Hardware/SubsystemsHardware/Subsystems
Thrust Chamber or Thruster Injector, combustion chamber, nozzle.
Liquid propellants are metered, injected, atomized, mixed and burned to form hotgaseous reaction products, which are accelerated and ejected at high velocity.
Cooling: Propellant cooled: fuel is circulated through cooling jackets
Radiation-cooled: high-temperature material radiates away excess heat
Ablative-cooled: heat-absorbing materials
Tanks: Propellant storage,
Feed Mechanism: Force propellants from tanks to thrust chamber(s),Pressure-fed
Pump-fed
Power Source: Furnish energy for feed mechanism,
Plumbing/Piping: Transfer of liquids,
Structure: Transmit thrust force,
Control Devices: Initiate and regulate propellant flow and thus thrust.
AAE 439
Ch6 –4
Booster Engines Sea Level Operation
Core Engines Sea Level Operation in combination with Boosters
Sustainer Engines 2nd Stage Operation
Upper Stage Engines 2nd, 3rd or 4th Stage Engines
Boosters Very High Thrust
Core High Thrust + High Isp
Sustainers Medium Thrust + High Isp
Upper Stage Low Thrust + Very High Isp
Requirements
For an exhaustive list of rocket systems, see http://www.astronautix.com/spaceflt.htm
Liquid Rocket Engine (LRE)Liquid Rocket Engine (LRE)
Applications
AAE 439
Ch6 –5
Propellant Types/SystemsPropellant Types/Systems
BipropellantTwo separate liquid propellants: oxidizer and fuel,
Stored separately and mixed in combustion chamber.
MonopropellantSingle substance containing oxidizing agent and combustible matter.
Stable at atmospheric conditions, but decompose when heated or catalyzed.
Cold Gas PropellantStored at very high pressure.
Cryogenic PropellantLiquefied gas at low temperatures,
Issues: venting and vaporization losses.
Gelled PropellantLiquid with gelling additive behaving like jelly or thick paint.
AAE 439
Ch6 –6
PROPELLANT FEED SYSTEMPROPELLANT FEED SYSTEM
Functions:Raising pressure of propellants.
Feeding propellants to one or more thrust chambers.
Pump-Fed System Pressure-Fed System
AAE 439
Ch6 –7
APPLICATIONS OF LRPAPPLICATIONS OF LRP
XCold Gas
XMonopropellant(electrical heating)
XXXMonopropellant
XXXXPressure-FedBiprop System
XXPump-Fed BipropSystem
Attitude ControlOrbit
Maintenance &Maneuver
Orbit InsertionLaunch
Applications
System Type
AAE 439
Ch6 –8
Engine CyclesEngine Cycles
High Pressure > 30 MPa
Large Thrust Range up to 10 MN
Booster, Core(nothing’s flying yet)
Full Flow StagedCombustion
High Pressure 13 - 26 MPa
Large Thrust Range 80 kN - 8 MN
Booster, Core,Upper Stage
Staged Combustion
Low-Medium Pressure 3 - 7 MPa
Small Thrust Range 80 - 200 kN
Upper StageExpander
Medium Pressure < 15 MPa
Large Thrust Range 60 kN - 7 MN
Booster, Core,Upper Stage
Gas Generator
Low Pressure < 2 MPa
Low Thrust < 40 kN
Upper Stage,
Space Propulsion
Pressure Drive
SpecificsApplicationCycle
AAE 439
Ch6 –9
6.2 PRESSURE–FED SYSTEMS6.2 PRESSURE–FED SYSTEMS
AAE 439
Ch6 –10
Types of Rocket EnginesTypes of Rocket Engines
Booster engines and upperstages; extreme space engines
Upper stages, space engines, small thrusters(sat prop & RCS); ‘big dumb booster’Applications
Gain from higher pc allowinghigher ε, for same pa orenvelope
Cold gas and monprop < 200 s
Biprop up to 340 sPerformance
Extra weight of TPAHeavier tanks – ptank ≈ 1.5 pchamberWeight
Expander
Gas generator
Staged combustion
Blowdown
Regulated pressure
More reliableComplexity
Pump–FedPressure–Fed
AAE 439
Ch6 –11
Pressurization OptionsPressurization Options
AAE 439
Ch6 –12
Pressure vs. Pump-FedPressure vs. Pump-Fed
Some rules/issues to consider when deciding on a pressurization mechanism:Lack of pumps drives a higher tank pressure.
Large propellant-tank volume drives the use of pumps.
In general, big systems use pumps, small systems don’t.
With larger system size, decreasing tank mass (lower pressure ⇒ thinner wallthickness) justifies additional complexity and cost of pumps.
The larger the tank, the bigger the mass savings (inert mass fraction!).
New/advanced/improved tank material and manufacturing technologies must beevaluated.
AAE 439
Ch6 –13
6.2.1 MONOPROPELLANT SYSTEMS6.2.1 MONOPROPELLANT SYSTEMS
AAE 439
Ch6 –14
Monopropellant SystemsMonopropellant Systems
Since Earth storable bipropellants have so much better performance, why everuse monopropellant systems?
The answer is simplicity and cost!!
Biprop systems need twice as many tanks, components, etc.
Biprop systems can only operate over a tight range of mixture ratios and inletpressures: Mixture Ratio (r) is defined as oxidizer/fuel flowrate.
Tight r control implies tight thermal control of tanks, etc.
Small biprop thrusters are not available.
Biprop systems cost approximately an order of magnitude more than monopropsystems.
Dual mode propulsion (NTO/N2H4) allows high performance main enginesand monoprop attitude control thrusters with shared fuel storage and control.
NTO: nitrogen tetroxide
N2H4: hydrazine
An interesting fact about NTO/MMH thrusters is that the typical mixtureratio r for these thrusters results in roughly equal-volume propellant tanks.
MMH: monomethyl hydrazine
AAE 439
Ch6 –15
Pressure-Fed System TypesPressure-Fed System Types
Basic Blowdown Monoprop System
Monoprop Thrusters
N2H4
GN2
P
T
T T T T T T T T
LService valve
Temperature transducerT
Pressure transducerP
Legend
Filter
Inlet/Outlet filter
L Latch Valve
Flow Control Orifice
AAE 439
Ch6 –16
DefinitionDefinition
Description: A single liquid propellant feeds into a catalyst bed in a thrusterwhere it catalytically and/or thermally decomposes to create a hot gas that isexhausted through a nozzle.
Benefits:Low system complexity (no combustion and often a regulator is not even needed),
Relatively inexpensive,
Requires only one set of tanks, components, and plumbing,
Good storability.
Drawbacks/Issues:Moderate to poor performance,
Moderately dangerous from a handling standpoint,
Contamination of a spacecraft’s external surfaces from exhaust gases.
Typically chosen as the propulsion system when:Small thrust levels, minimum impulse bits, and/or Δv are required,Operational lifetime of the system is long,
Science dictates known and quantifiable contamination of external body.
AAE 439
Ch6 –17
ExamplesExamples
Typical Monopropellant Thruster
Typical Monopropellant Gas Generator
AAE 439
Ch6 –18
Important ConsiderationsImportant Considerations
Stable operation of monopropellant thrusters over a wide range of inletpressures typically allows operation in blow-down mode.
Catalyst bed and valve heaters are usually required on all monopropellantthrusters:For thermal catalyst beds, propellant decomposition will not take place unless cat
bed is pre-heated.The firing duty cycle will dictate how long and often the cat. bed heaters are
required. If cat. bed heaters are not wanted in the design with a spontaneous catalyst, a
warming pulse (“cold start”) could be used before the maneuver(s) are conductedto “gently” warm up the cat. beds. Cold starts rapidly degrade the catalyst bed and are not recommended unless operational
lifetime is short.
A diaphragm positive expulsion device (PED) is typically used for hydrazinesystems:Several elastomers which are compatible with hydrazine exist,HAN-based monoprop systems cannot use elastomeric diaphragms/bladders yet.
HAN: hydroxylammonium nitrate
Monopropellant systems will always have some hold-up/residual remaining inthe propellant tanks at the end of the mission that must be accounted for.
AAE 439
Ch6 –19
Other OptionsOther Options
Blow-down systems typically pay a mass penalty for their simplicity:Option for reducing propellant tank mass are:
Recharge systems
Regulated systems.
If off-the-shelf tanks are slightly too small, “ullage tanks” can be used toincrease effective tank volume.
Pumps can also be considered to allow storage of propellant at low pressureand with minimal beginning-of-life-ullage.
AAE 439
Ch6 –20
6.2.2 BIPROPELLANT SYSTEMS6.2.2 BIPROPELLANT SYSTEMS
AAE 439
Ch6 –21
DefinitionDefinition
Description: Liquid oxidizer and liquid fuel feed into a thrust chamber wherethey mix and react chemically; the combustion gases then accelerate and areexhausted through a converging-diverging nozzle.
Benefits:Relatively high achievable performance for chemical propulsion.
Drawbacks/Issues:High cost and complexity (“twice” as many components as a monopropellant
system, many more variables to be controlled/analyzed),
Bipropellants are often dangerous from a system safety/handling standpoint,
Reaction temperatures are often high enough that combustion temperatures aretoo high to be accommodated by the thruster materials, Sometimes need extremely tight mixture ratio control
Nasty exhaust products (contamination of spacecraft’s external surface or thesurface of an ephemeral body).
Typically chosen as the propulsion system when:High thrust levels, total impulse, and/or ΔVs are required to offset dry mass,Operational lifetime of the system is long,
The spacecraft temperature is tightly maintained.
AAE 439
Ch6 –22
Pressure-Fed System TypesPressure-Fed System Types
Basic Biprop Propulsion System
Service valve
Temperature transducerT
Pressure transducerP
Legend
Filter
Inlet/Outlet filter
L Latch Valve
Flow Control Orifice
AAE 439
Ch6 –23
Bipropellant SystemsBipropellant Systems
Unless thruster inlet conditions are carefully controlled, thrusters can burnthrough or otherwise fail.Not a problem with monoprop thrusters,
Many variables affect thruster inlet conditions: Schematic
Magnitude and location of pressure drops
Relative pressure drops in oxidizer and fuel feed systems
Propellant and propellant tank temperature and pressure,
Propellant vapor pressure and state of pressurant saturation.
Due to typically large propellant masses and need to carefully control inletconditions, most biprop systems include a pressurization system.Usually can’t afford the mass or volume penalty, or large excursion in inlet
pressures associated with blow-down operation for the entire mission.
Oxidizer is not compatible with most elastomers, so all-metal PropellantManagement Devices are typical in oxidizer tanks.
Separate propellants need to be kept apartBoth in liquid and vapor forms.
AAE 439
Ch6 –24
Important ConsiderationsImportant Considerations
Bipropellant systems will always have some hold-up/residual remaining in thepropellant tanks at the end of the mission that must be accounted for (moreso than monopropellant systems due to mixture ratio uncertainties),
Chugging (feed system frequency interaction with engines),
Vapor migration in the feed system,
Center of gravity shift during operation (must number, locate, and size tanksto account for this),
A repeatable minimum impulse bit is difficult to achieve due to mixture ratioand combustion uncertainties,
The design of the propellant management devices (PMDs) in the tanks aremuch more complicated than with monopropellant systems due to tighterpropellant management control and material incompatibilities (e.g., NTO isincompatible with many elastomeric materials),
Bipropellant engines typically have tight operating boxes (mixture ratio,pressure, temperatures), tighter the operating box, the more complex the overall propulsion system.
AAE 439
Ch6 –25
Important ConsiderationsImportant Considerations
Design Considerations:Plumbing in the pressurization system has to:
Provide isolation of the high pressure pressurant tank(s) from the relatively low pressurepropellant tanks.
Prevent migration and mixing of propellant vapors, if a pressurization system common toboth propellants is used.
Prevent mixing of propellants (except in thrusters, of course).
Maintain control of the flow of both propellants such that the thruster inletconditions stay within acceptable limits.
Control pressure drops in the system such that effects associated with pressurantcoming out of solution are within acceptable limits
Biprop systems often provide for the isolation of the pressurization system duringlong periods of system inactivity, or after enough of the propellant has beenexpelled that blow-down operations are possible. Dual mode systems often provide for the isolation of the oxidizer system after the biprop
main engine has been used, either leaving the hydrazine tank in blow-down, or keepingthe hydrazine tank at regulated pressure
AAE 439
Ch6 –26
ExampleExample
Viking Orbiter Bipropellant & Cold-Gas Propulsion System
Legend
Filter
Pressure transducerP
Temperature transducerT
Service valve
Pyrotechnic valve (normallyclosed)Pyrotechnic valve(normally open)
Latch valveLGas regulator
Flexible line
Field Joint
Orifice
Burst Disc / ReliefValve Assembly
T
P P
MMH
GHe
L
T P
L
TP
T
T
GHe
P
P
GN2 GN2
L L
TT
P P
Cold Gas Attitude ControlSubsystem (ACS)
Propulsion Subsystem
NTO
GHeT
T
T
T
Check valve
1323 N Main engine
134 mN Thrusters
GimbalActuator
AAE 439
Ch6 –27
6.3 CYCLE DEFINITION FOR BIPROP SYSTEMS6.3 CYCLE DEFINITION FOR BIPROP SYSTEMS
AAE 439
Ch6 –28
Engine CyclesEngine Cycles
High Pressure > 30 MPa
Large Thrust Range up to 10 MN
Booster, Core(nothing’s flying yet)
Full Flow StagedCombustion
High Pressure 13 - 26 MPa
Large Thrust Range 80 kN - 8 MN
Booster, Core,Upper Stage
Staged Combustion
Low-Medium Pressure 3 - 7 MPa
Small Thrust Range 80 - 200 kN
Upper StageExpander
Medium Pressure < 15 MPa
Large Thrust Range 60 kN - 7 MN
Booster, Core,Upper Stage
Gas Generator
Low Pressure < 2 MPa
Low Thrust < 40 kN
Upper Stage,
Space Propulsion
Pressure Drive
SpecificsApplicationCycle
AAE 439
Ch6 –29
Pump-Fed Cycle TypesPump-Fed Cycle Types
Open Cycles
Gas Generator Cycle Combustion Tap-Off Cycle Coolant Bleed Cycle
Open cycle
GG burns non-stoichiometric toeliminate turbine cooling
+ Fairly simple
+ Wide thrust operating range
− Turbine exhaust gives low Isp ≈effective loss in performance
− GG required
RS-27, MA-5, STME, Titan
Open cycle similar to GG, but useschamber pressure rather than GG todrive turbine
+ No GG required
− Difficult to throttle, start
− Narrow thrust operating range
Saturn V
Open cycle similar to combustiontap-off, but uses coolant bleed(vaporized) to run turbine
+ No GG required
− Limited to cryogenic fuels
− Pressure and thrust limited bythermal properties
AAE 439
Ch6 –30
Pump-Fed Cycle TypesPump-Fed Cycle Types
Closed Cycles
Staged Combustion CycleExpander Cycle
Closed cycle; most of coolant fed tolow pressure ratio turbines
+ Good performance, i.e. closed cycleefficiency
+ Simple design
+ Low weight
+ Self starting
− Limited to low pc < 1,100 psi
− Limited to cryogenic fuel
Closed cycle with preburnerreplacing GG
+ High performance
+ High pc and thrust capability
− Very complex, lesser reliability
− Advanced turbine/pump requiredfor high pc (boost pumps)
AAE 439
Ch6 –31
Pump-Fed Propulsion System OptionsPump-Fed Propulsion System Options
Thermodynamic cycle chosen is hopefully the optimal for the application.
AAE 439
Ch6 –32
Pump-Fed Propulsion SystemPump-Fed Propulsion System
AAE 439
Ch6 –33
6.4 COMPONENTS OF LRP SYSTEMS6.4 COMPONENTS OF LRP SYSTEMS
AAE 439
Ch6 –34
TANKAGETANKAGE
All liquid propellant systems require tanks for storage of propellants.Separate fuel and oxidizer tanks in bipropellant systems,
Single tanks for monopropellant systems,
Number of tanks dictated by system redundancy and control of vehicle’s center ofgravity.
Tank Shape:Launch Vehicle: cylindrical shape to reduce frontal area and thus minimize drag,
Space Vehicle: spherical shape to increase packaging efficiency.
Tank Arrangement:
Ox
Fuel
Ox Ox
Fuel
Ox
Fuel
Tandem Concentric Twin
AAE 439
Ch6 –35
TANKAGETANKAGE
Surface Area vs. Shape:
Volume Requirement is dictated by:Required mass of liquid propellant: Vprop
Changes in liquid density due to temperature the tank might encounter (Ullage): Vullage
Boil-Off of given propellants: VBO
Trapped propellant in tanks or/and feed lines: Vtrap.
V
total= V
prop+ V
ullage+ V
BO+ V
trap
Asphere
Acylinder
=1.5
2
3 !2 L D( )2
3
1+ 2 ! L D( )
where L/D=length-to-diameter of cylindrical tank
0
0.2
0.4
0.6
0.8
1
0 0.5 1 1.5 2
Su
rface
Are
a R
ati
o
L/D Ratio
AAE 439
Ch6 –36
TANKAGETANKAGE
Tank SizingPressure at the bottom of tank
Case Thickness
Buckling Analysis Static Equilibrium of Forces on Tank
Stress in Thin-Walled Tank
Critical Stress in Thin-Walled Tank
p
totalt( ) = p
ullaget( ) + !
propellantg
earth"a t( ) "h t( )# p
ambientt( )
h t( ) 2R
p
ambientt( )
F
axial
p
totalt( )
p
ullaget( )
t
case= p
total
R
!
VehicleAcceleration
PropellantDensity
Max. AllowedStress
F
z! = " R2
pullage
# pambient( ) # F
axial
! "F
A=
# R2
pullage
$ pambient( ) $ F
axial
2# R t
!critical
= "E 9 #t
case
R
$
%&'
()
1.6
+ 0.16 #t
case
L
$
%&'
()
1.3*
+
,,
-
.
//
L
Young’sModulus
!
critical> !
Non-Buckling Condition
AAE 439
Ch6 –37
TANKAGETANKAGE
Propellant Management throughout Mission ProfileSupplying engines with gas-free propellant,
Draining maximum amount of loaded propellant minimizing residuals,
Preventing propellant slosh: no forces or moments transmission to spacecraftstructure which might overwhelm attitude-control system,
Changing acceleration environment of the mission profile complicates propellantmanagement.
Propellant–Expulsion Devices:Positive–expulsion devices use physical barriers between propellant and pressurant
gas. Examples: bladders, pistons, diaphragms, bellows.
Passive–expulsion devices use surface tension on propellant to keep the fluid incontact with the propellant drain. Examples: vanes, porous sheets, screens.
Gas can accumulateat the outlet and keepengines from starting.
AAE 439
Ch6 –38
TANKAGETANKAGE
Positive–Expulsion Devices for Propellant ManagementComparison of Different Technologies Examples of Different Technologies
AAE 439
Ch6 –39
COMBUSTION CHAMBERCOMBUSTION CHAMBER
Propellant Choice is a strong influence on Sizing the Combustion Chamber.
Cross-sectional area is dictated by injector design.
Length of the combustion chamber is determined by the residence time toensure complete combustion of the propellants.Residence Time:
Characteristic Chamber Length:
tresidence
=l
chamber
vpropellant
=!
propellantA
chamberl
chamber
!m=
c *
RTpropellant
"A
chamberl
chamber
Athroat
Geometric Featuresof
Combustion Chamber
PropellantCharacteristics
L* =A
chamberl
chamber
Athroat
AAE 439
Ch6 –40
COMBUSTION CHAMBERCOMBUSTION CHAMBER
Combustion InstabilitiesLow Frequency: Chugging Mode
Characteristic Frequency range is between 10 and 200 Hz. This instability occursdue to an interaction between chamber and feed system.
Medium Frequency: Buzzing Mode
Characteristic Frequency range is (arbitrarily) placed between 20 and 1000 Hz.This instability is a result from either flow instabilities or resonance with chamberstructure.
High Frequency: Screaming Mode
Characteristic Frequency range is above 1000 Hz. This instability is due tointeractions of the combustion process with chamber acoustics.
AAE 439
Ch6 –41
INJECTORSINJECTORS
Injectors are responsible for metering, distributing, and atomizing propellantsfor efficient combustion within the combustion chamber.
Injector Classification:Doublet & Triplet Impinging Stream Patterns,
Self-Impinging Stream Pattern,
Shower Head Stream Pattern,
Spray Injection Pattern,
Splash Plate Pattern,
Coaxial Type, etc.
Injector SizingMass Flow Rate of an Injector Element (Bernoulli’s Equation)
Number of Elements
Injector Face Area
!m
element= c
DA
element2! g "p
N =!m
total
!melement
Achanber
=
Nox+ N
fuel
ND
N
D=
N
A= # of elements per area due to design
AAE 439
Ch6 –42
INJECTORSINJECTORS
E
xamples
Ch6–39 a
INJ
EC
TO
RS
INJ
EC
TO
RS
AA
E 439
AAE 439
Ch6 –43
6.5 EXAMPLES OF LRP CYCLES6.5 EXAMPLES OF LRP CYCLES
AAE 439
Ch6 –44
Pump-Fed Propulsion SystemPump-Fed Propulsion System
AAE 439
Ch6 –45
Gas Generator Engines - Production LineGas Generator Engines - Production Line
79.1
93.8
58.5
82.1
128.7
—
—
82.7
97.6
120.6
55.4
70.1
77.8
87.5
—
41.3
Thrust/Weight
Sea Level
Vacuum
2528103533361284458335711953341Dry Weight [lb]
1225849.2154510.4862.5Area Ratio
521.87367198278271647876521.8ChamberPressure
2.2452.272.251.861.915.301.704.77Mixture Ratio
254.8
302.1
220
308
265263.3
318
246.1
304
340
431
248.5
278.4
—
445.1
Specific Impulse
Sea Level [lbf]
Vacuum [lbf]
fixedfixedfixedfixedfixedfixedfixedThrottle
199,945
237,067
60,500
85,000
429,500—
106,200
447,300
552,500
198,000
250,500
152,000
171,000
—
14,100
Thrust
Sea Level [lbf]
Vacuum [lbf]
LOX/RP1LOX/RP1LOX/RP1N2O4/Aerozine5
N2O4/Aerozine5
LOX/LH2N2O4/UH25LOX/LH2Propellant
Delta II/III
1st Stage
Atlas II
1st Stage
Atlas II
1st Stage
Titan IV
2nd Stage
Titan IV
1st Stage
Ariane 5
1st Stage
Ariane 4
1st Stage
Ariane 4
3rd StageStage
RocketdyneRocketdyneRocketdyneAerojetAerojetSEPSEPSEPContractor
RS-27MA-5ASUSTAINER
MA-5ABOOSTER
LR 91LR 87HM 60VIKING VOHM 7BENGINE
AAE 439
Ch6 –46
Operating Schematic - RS-68Operating Schematic - RS-68
Main LOXInlet
Main FuelInletHelium
Spin Line GGOxidValve
GGFuel
Valve
LOX TankPressurization
Roll ControlNozzle
Main FuelValve
RegenCooledChamber
HeatExchanger
Fuel Turbopump
FuelTankRepress
Gas Generator(GG)
MainOxidValve
OxidTurbopump
Ablative LinedNozzle
AAE 439
Ch6 –47
Saturn V/J-2 EngineSaturn V/J-2 Engine
Saturn II
AAE 439
Ch6 –48
Tap-Off CycleTap-Off Cycle
AAE 439
Ch6 –49
Expander CycleExpander Cycle
Fuel cool-down andpressure relief valve
Turbine
Oxidizer flowcontrol valve
Main fuelS/O valve
Venturi
Main propellant valves
Fuel bypassand thrust
control valve
Regenerativecooling channels
AAE 439
Ch6 –50
RL-10 EvolutionRL-10 Evolution
42.3
46
—
56
—
55
—
50
—
50
—
50
—
50
Thrust/Weight
Sea Level
Vacuum
316370300300300300300Dry Weight [lb]
4.28846157404040655Area Ratio
578570475475475475475ChamberPressure
65.555555Mixture Ratio
334.8
365.1
—
449
—
444.4
—
442.4
—
433
—
429
—
424
Specific Impulse
Sea Level [lbf]
Vacuum [lbf]
13,352
14,560
—
20,800
—
16,500
—
15,000
—
15,000
—
15,000
—
15,000
Thrust
Sea Level [lbf]
Vacuum [lbf]
LOX/LH2LOX/LH2LOX/LH2LOX/LH2LOX/LH2LOX/LH2LOX/LH2Propellant
CentaurCentaurCentaurCentaurCentaurCentaurCentaurStage
RocketdyneP & WP & WP & WP & WP & WP & WContractor
RL-10-A5RL-10-A4RL-10-A3-3ARL-10-A-1-3RL-10-A-1-1RL-10-A-3RL-10-A-1ENGINE
AAE 439
Ch6 –51
RL-60/RL-10 ComparisonRL-60/RL-10 Comparison
AAE 439
Ch6 –52
RL-60 Combustion ChamberRL-60 Combustion Chamber
AAE 439
Ch6 –53
Staged Combustion Engine - Production LineStaged Combustion Engine - Production Line
AAE 439
Ch6 –54
Space Shuttle Main EngineSpace Shuttle Main Engine
AAE 439
Ch6 –55
Space Shuttle Main EngineSpace Shuttle Main Engine
Components and Subcomponents
AAE 439
Ch6 –56
Space Shuttle Main EngineSpace Shuttle Main Engine
SSME Powerhead
AAE 439
Ch6 –57
Ox-Rich Staged Combustion (ORSC)Ox-Rich Staged Combustion (ORSC)
AAE 439
Ch6 –58
RD-180 Engine SchematicRD-180 Engine Schematic
FuelInlet
HeatExchanger
LOX BoostPump
Fuel BoostPump
Fuel InletValve
FuelPump
LOXPump
Preburner
Turbine
ChamberMain Fuel
Valve
StartBottle
MixtureRatio Valve
ThrustControlValve
Main LOX Valve
HypergolicStart
Ampoule
Hot GasBellows
LOXInlet
AAE 439
Ch6 –59
RD-180 Engine CharacteristicsRD-180 Engine Characteristics
Characteristics demonstrate heritage to RD-170Two chamber derivative of the RD-170,
Identical chambers, scaled turbopumps,
Staged combustion cycle - LOX rich PB,
LOX/kerosene propellants,
2 thrust chambers (+/- 8o gimbal),
LOX & fuel boost pumps,
Single shaft high pressure turbopump 2 stage fuel pump,
single stage LOX pump,
single stage turbine,
Self contained hydraulic system (valves, TVC)powered with kerosene from fuel pump,
Hypergolic ignition.
AAE 439
Ch6 –60
RD-180 Technical DataRD-180 Technical Data
Full Power (100%) Minimum Power (47%)Chamber Pressure (psi) 3,722 1,755Flow Rate (lb/s) 2,756 1,152Total Sea Level Thrust (lb) 860,200 365,500Total Vacuum Thrust (lb) 933,400 438,700Sea Level Isp (sec) 311.9 278.7Vacuum Isp (sec) 338.4 334.6Mixture Ratio 2.72Nozzle Area Ratio 36.9Throat Area (in2) 67.5Engine Length (in) 141Circumscribed Exit Diameter (in) 124Single Nozzle Exit Diameter (in) 56.9Gimbal Angle +/- 8 degWeight, Dry (lb) 12,225Weight, Wet (lb) 13,260Thrust/Wt (Sea Level, with TVC ) 70.4Thrust/Wt (Sea Level, No TVC ) 74.5
AAE 439
Ch6 –61
Overall Interface SchematicOverall Interface Schematic
Frame
Turbine
Hot GasDucts
LOX BoostPump
Fuel BoostPump
LOX Pump
Fuel Pump
GimbalActuator
HeatExchanger
Preburner
Nozzle
GimbalUnit
CombustionChamber(2)
FuelInlet Valve
AAE 439
Ch6 –62
Full-Flow Staged CombustionFull-Flow Staged Combustion