6. liquid rocket propulsion (lrp) systems · aae 439 ch6 –1 6. liquid rocket propulsion (lrp)...

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

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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.

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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

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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.

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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

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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

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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



Gas Generator

Low Pressure < 2 MPa

Low Thrust < 40 kN

Upper Stage,

Space Propulsion

Pressure Drive

SpecificsApplicationCycle

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Ch6 –9

6.2 PRESSURE–FED SYSTEMS6.2 PRESSURE–FED SYSTEMS

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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

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Ch6 –11

Pressurization OptionsPressurization Options

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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.

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Ch6 –13

6.2.1 MONOPROPELLANT SYSTEMS6.2.1 MONOPROPELLANT SYSTEMS

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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

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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

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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.

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Ch6 –17

ExamplesExamples

Typical Monopropellant Thruster

Typical Monopropellant Gas Generator

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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.

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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.

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Ch6 –20

6.2.2 BIPROPELLANT SYSTEMS6.2.2 BIPROPELLANT SYSTEMS

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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.

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Ch6 –22

Pressure-Fed System TypesPressure-Fed System Types

Basic Biprop Propulsion System

Service valve



Legend

Filter

Inlet/Outlet filter

L Latch Valve

Flow Control Orifice

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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.

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Ch6 –24


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.

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Ch6 –25


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

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Ch6 –26

ExampleExample

Viking Orbiter Bipropellant & Cold-Gas Propulsion System

Legend

Filter



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

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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



Staged Combustion

Low-Medium Pressure 3 - 7 MPa

Small Thrust Range 80 - 200 kN

Upper StageExpander

Medium Pressure < 15 MPa



Gas Generator

Low Pressure < 2 MPa

Low Thrust < 40 kN

Upper Stage,

Space Propulsion

Pressure Drive

SpecificsApplicationCycle

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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

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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)

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Ch6 –31

Pump-Fed Propulsion System OptionsPump-Fed Propulsion System Options

Thermodynamic cycle chosen is hopefully the optimal for the application.

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Ch6 –32

Pump-Fed Propulsion SystemPump-Fed Propulsion System

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Ch6 –33

6.4 COMPONENTS OF LRP SYSTEMS6.4 COMPONENTS OF LRP SYSTEMS

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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

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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

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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

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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.

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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.

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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

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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


Components and Subcomponents

AAE 439

Ch6 –56


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

6. liquid rocket propulsion (lrp) systems · aae 439 ch6 –1 6. liquid rocket propulsion (lrp)...

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