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Liquid Propulsion

Stanford Aeronautics & Astronautics

Lecture 9, 4/30/2019

A A 1 0 3 A I R A N D S P A C E P R O P U L S I O N

AA 103Air and Space Propulsion

Contents

Liquid Propulsion

– Intro

– Liquid Rocket Propulsion

• Monoprops and Biprops

• Cycles

• Propellants

• Thrust Chambers

• Cooling

– Mission Design

• Delta-V Budget

• Mission Analysis


Liquid Propulsion Pros & Cons

• Advantages

– Highest achievable specific impulse for chemical

propulsion

– Controllability (throttle, pulse)

– Amenable to ground testing ad pre-launch checkout

– Low cost / high reliability

– Enables component redundancy

– Most have relatively clean exhaust

• Disadvantages

– Lower propellant density than solids

– Higher part count than other chemical systems

– More complicated in some cases

– Concerns about leakage / hazards

– Concerns about combustion instability

– Needs mixture ratio control

– Needs special design for zero-g start

3


Monopropellants

• A monopropellant is a single fluid capable of undergoing an exothermic decomposition reaction to yield a gaseous product

• Only a single tank and propellant feed system is required

• Injection is simplified – no mixing required

• Field operations and on-orbit operations (for satellite systems) are simpler

Pressurized

Monopropellant

Fluid Tank Main

Valve Catalyst

Bed

NozzleInjector

Nitrous Oxide Catalyst Bed


5

Decomposition of Monopropellants

• Catalytic decomposition

• Hydrogen peroxide (H2O2):

– Oxidizer

– ISP = 150 s

– 2H2O2 (l) + n H2O (l) → (n+2) H2O (g) + O2 (g) + heat

• Hydrazine (N2H4):

– 3N2H4 → 4 NH3 + N2 + heat

– Followed by ammonia decomposition (endothermic) into N2 and H2

– Most popular: good handling characteristics, relative stability under normal

storage conditions, clean decomposition, ISP = 225-230 s

– For high ISP, attempt to get 30-40% ammonia dissociation

– For gas generator (cooler), up to 80% ammonia dissociation

– Large system: NASA’s Tracking and Data Relay Satellite (TDRS)

• Advanced “monopropellants”: more complex; subject of R&D


6

Bipropellant – Pressure-fed

Pressure-regulated gas keeps tanks at desired pressure.

Tank pressure forces propellant into Thrust Chamber

Source: R.H. Humble, Space Propulsion

Analysis and Design


7

Pressure-fed System

Pressure drops continuously from the tank to the nozzle


Analysis and Design


8

Bipropellant – Turbopump-fed

System contains pump and pump drives.

This figure illustrates a gas generator driving the turbine

Merlin 1D

Vacuum

Merlin 1D

Engine

Source: R.H. Humble, Space Propulsion Analysis and Design


9

Turbopump Engine Cycles

Higher fuel pressure drop between pump outlet and

chamber

Can use “overboard” flow for roll control

These are 3 cycles that most often drive propellant pumps


Analysis and Design


10

Turbopump-fed System

Pressure drops continuously to the pump inlet. The pump

drastically drives up the pressure, which then continuously

drops again. This schematic is for a gas generator cycle


Analysis and Design


11

Bipropellant Combustion

• Energetic release of heat as the fuel vapors react with oxidizer vapors to

form combustion products

• As the gas temperatures rise, thermal energy causes dissociation of the

gases

• As the gases expand and cool in the nozzle, re-association and energy

release occurs, but at finite rate - so it is incomplete.

• Jets formed by injection orifices

• Droplets formed by collision of jets,

by ligament breakup of free jets, or

by drag-induced shattering

• Droplets vaporize from heat of

combustion gases

• Further mixing and reaction of

vapors due to diffusion and

turbulent mixingSource: G.P. Sutton, Rocket Propulsion Elements


12

Choosing a Propellant

• Many factors go into choosing a propellant combination

– Performance

– Cost

– Density

– Corrosiveness

– Toxicity

– Stability

– Storability

– Self igniting (hypergolic)

– Lubricity (for turbopumps)

• The choice of mixture ratio to operate the engine at is often a

compromise

– Launch vehicles like O/F that yields higher density for smaller tanks

– Propulsion engineer may choose a lower O/F in order to lower flame

temperature or to accommodate a large amount of film cooling

– O/F may be chosen for equal volume to use identical fuel and ox tanks or to

better package in spacecraft structure


13

Liquid Oxidizers

• LOX: cryogenic (90 K)

• Hydrogen peroxide (H2O2)

– 70 to 99%, remaining is water

– Can be used as monopropellant as well (H2O2 → H2O + 1/2 O2 + heat)

• Nitric acid (HNO3)

– Corrosive, toxic

– Used between 1940 and 1965 – No longer used

• Nitrogen tetroxide (N2O4)

– High density yellow/brown liquid

– Most common storable oxidizer

– Narrow liquid temperature range, easily frozen or vaporized

– Contaminated by water

– Hypergolic with many fuels (e.g. MMH on OMS of STS)

• Nitrous oxide (N2O), a.k.a. laughing gas

– “Self-pressurized” but low performance

– Not used much in liquid systems, mostly hybrids


14

Liquid Fuels

• Hydrocarbons

– RP-1 – Rocket-Propellant grade of Kerosene(CH1.97)

• High performance (ISP = 350 s)

• Bright yellow plume

• Storable

– Methane – CH4

• High performance (ISP = 360 s)

• Cryogenic (100 K)

• 0.4 specific gravity, i.e. larger than LH2 (<0.1), but lower than RP-1 (0.8)

– Propane (C3H8) and propylene (C3H6)

• LH2

– Highest performance – ISP = 450 s

– Cryogenic (20 K)

• Hydrazine (N2H4) and derivatives

– Toxic; Used as mono- or bipropellant

– MMH (monomethylhydrazine, CH3NHNH2)

– UDMH (unsymmetrical dimethylhydrazine, (CH3)2NNH2): More stable than hydrazine


Liquid Oxidizers and Fuels

• Properties of Some Liquid Oxidizers

• Properties of Some Liquid Fuels

15

Source: G.P. Sutton, Rocket Propulsion

Elements


16

Vacuum Specific Impulse

Source: D.K. Huzel and D.H. Huang, Modern Engineering

for Design of Liquid-Propellant Rocket Engines


Thrust Chamber Assembly


18

Injector Types

• Introduce, meter the propellant flow

– major influence on the performance, chamber heat load, and combustion stability of the engine

– “Black art”: Despite advances with CFD modeling, much of injector design is empirical

• Impinging Elements - form fans of mixed droplet sprays

– Like (e.g. 2 ox.) and unlike doublets

– Like and unlike (e.g. fuel/ox.) triplets

– Splash plate and sheet forming elements

– Pintle injector

– Swirl elements

• Non-impinging elements - Rely on diffusion and turbulence to mix propellants

– Showerhead injector

• Shear Mixing - Rely on velocity difference to mix, usually “gas-liquid” injection

– Coaxial elements - common on LOX/LH2 engines (SSME, RL10, etc.)


Elements


Liquid Rocket Injector Design – SSME Injector

19


Elements


20

Acoustic Combustion Instability

• Types of acoustic (high frequency) instability:

– Intrinsic acoustic mode:

• Small propellant flow oscillations

• Determined by coupling between combustion process and chamber acoustics

– Injection coupled acoustic mode

• Coupling between propellant injection process and chamber acoustics

• Frequency of acoustic oscillations (> 1000 Hz) is function of gas speed of

sound and chamber geometry (see Fig. from Huzel and Huang)

• Can result in engine damage in a fraction of a second

Source: D.K. Huzel and D.H. Huang,

Modern Engineering for Design of Liquid-

Propellant Rocket Engines


21

Thrust Chamber Cooling

• The thrust chamber must contain the intense combustion heat

and pressure of the burning rocket propellants

• Core combustion temperatures over 5000 °F are achieved with

most propellant combinations

• Thrust Chamber Cooling Methods

– Film or boundary layer cooling

– Transpiration cooling

– Heat sink

– Adiabatic wall

– Radiation

– Ablative

– Regenerative


22

Regenerative Cooling, 1

• Thrust chamber cooled by

propellant flowing in a jacket in

chamber wall

• Heat loss to chamber wall is

recovered as heat to propellant

• Velocity of coolant is designed to be

approximately proportional to the

heat load [K] volume leaving fluid of temp.

[K] volume entering fluid of temp.

[J/kg/K] coolant of capacity heat

[kg/s] coolant offlow mass

[W] convection from addition heat

[W] radiation from addition heat

)()()(

out

in

coolant

conv

rad

inoutcoolantoutradinconvrad

T

T

c

m

Q

Q

TTcmQQQ



23

Regenerative Cooling, 2

0

1

2

3

4

5

Dia

m R

ati

o

0

2

4

6

8

10

He

at

Flu

x (

BT

U/i

n2

-s

ec

)

0 20 40 60 80 100 120

DISTANCE (IN.)

Diameter Ratio Heat Flux

Rocket Engine Chamber ModelPredicted Wall Heat Flux

• Types

– Tubular - chamber is constructed from a tube

bundle that is brazed or welded together

– Milled channel - cooling slots are milled into a

solid copper chamber wall

• channels closed out with electro-deposited

nickel

– Dual wall jacket with helical cooling channels

• Common for low-thrust chambers


24

Film Cooling

• A.k.a boundary layer

cooling

• One of the propellants,

usually the fuel, is

injected near the wall to

provide a lower

temperature gas layer

• Film cooling can be

achieved by direct

injection of fuel at wall, or

by using fuel-rich injector

elements near the wall

• Film cooling is usually

required in addition to

other cooling schemes,

particularly at high

chamber pressures

)T (assumed vol. the leaving coolant vaporized of etemperatur

volume entring coolant of etemperatur

[kg/s] coolant of rateflow mass

[W] wallthe intoflow heat

[K] fluid the of temp. onvaporizati

[J/kg] fluid of onvaporizati of heat

[J/K/kg] fluid of capacity heat

[J/K/kg] P ct. at gases vaporized of capacity heat

][)(

C

out

in

film

wall

vap

fg

p

T

T

invapfgpfilmwallinconvrad

T

T

m

Q

T

h

c

c

TTchdTcmQQQout

vap



25

Radiation Cooling

• Commonly used for low

thrust/low pressure spacecraft

propulsion

• Hot chamber walls radiate heat

to space

• Wall zone gas temperatures 200

- 2000 °F hotter than chamber

wall due to radiation heat flux

• Typical chamber materials

– C103 Niobium alloy with

oxidation protection coating

• 2300 - 2700 °F operating temp

– Platinum/rhodium alloy

chambers

• No protective coatings required


– Rhenium chamber with iridium

oxidation coating


wallnear temp. exterior

temp. wall

factor shape

const. Boltzmann-Stefan

][m wallof area sectional-cross

1)(0 emissivity

)(

2

44

tenvironmen

w

tenvironmenwinconvrad

T

T

F

A

TTFAQQ



26

Ablative Cooling

[kg/s] material ablation of enthalpy

[kg/s] material ablative offlow mass

)(

ablative

ablative

ablativeablativeinconvrad

h

m

hmQQ

• Silica fibers in a phenolic resin

matrix that chars and erodes as

it absorbs heat

• Silica fibers provide structural

strength to char layer

– Molten glass layer helps block

oxidation of carbon char layer

• Phenolic matrix pyrolizes into a

carbonaceous char layer,

releases gases that remove heat

and create gas film at surface



27

Real Nozzles

• Losses:– Divergence loss (shape)

– Viscous drag loss (B.L.)

– Chemical kinetics loss (non-equilibrium)

– Typical 98.5% efficiency

Low altitude

Over-expanded

High altitude

Under-expanded


28

Ignition Systems

• Proper ignition of a rocket engine is one of the most important aspects

of a rocket test

• Types of Ignition Systems

– Hypergolic

• Use of propellants that self ignite on contact

• Nitrogen Tetroxide and amines (hydrazine, MMH, UDMH, etc.)

– Pyrotechnic Ignition

• Use of a solid propellant squib or grain

• Common in early days of rocketry

– Augmented Spark Igniter (Precombustor ignition)

• Main chamber lit by a separate torch-like combustor fed by some fuel and ox

that is ignited by a spark plug

– Hypergolic slug

• A cartridge of a hypergolic fluid is installed in one feedline so that it is pushed

into chamber by main propellant flow

• Often a pyrophoric fluid is used in the fuel line - ignites with oxygen


29

Typical Configurations

• Large launch vehicle first stage and second stage

– Pump-fed, i.e. light tanks

– Regen engine: metal, such as inconel…

– Possible ablative nozzle extension

– Typ. Pc: 5.5 MPa (800 psia) – 8.5 MPa (1,200 psia) or higher

• Small launch vehicle or upper stage

– Pressure-fed, large portion of blow-down (50% or more), heat exchangers

– Ablative engine: silica/phenolic with carbon over-wrap

– Typ. Pc: 1.0 MPa (150 psia) – 2 MPa (300 psia) or higher

• Spacecraft

– Low chamber pressure

– Pressure-fed

– Typ. Pc: 1.0 MPa (150 psia) – 2 MPa (300 psia)


Mission Analysis


31

Basics of propulsion

• Thrust: Generated by expelling propellant at high speed (momentum

exchange)

Propellants

m

Chamber Nozzle

Vex

pex

patm

pc

exatmexexT AppVmF )(

SL) sec.(@ per burnt propellant of Weight

Thrust

0

0

gm

F

dtmg

dtF

I T

t

t

T

SP

mass final

mass initialln SPIgV

• Specific impulse: total impulse per unit weight of propellant (measures

energy content of propellant and how it is transferred in thrust)

• V & rocket equation:


Mission Delta-V Budget

• Suborbital Delta-V ~ 1.4 km/sec

• Launch to LEO:

– Not only increase in velocity to 7.8

km/sec

• Additional 1.5-2 km/sec required for

gravity loss and atmospheric drag

• At the equator, the velocity of

rotation of the Earth is ~ 450 m/s

– Orbital rockets needs to be

considerably larger than suborbital

• Stationkeeping (low thrust)

32

• Earth – Moon – Mars (high thrust)

https://en.wikipedia.org/wiki/Delta-v_budget


33

Mission Analysis, 1

• Payload: usually refers to the

part of the s/c which performs

the mission; for propulsion,

usually refers to the dry mass

“carried” by the propulsion

system

• Wet s/c = payload (or dry s/c) +

prop. syst. + propellant +

pressurant

• Pressurant neglected for

requirements def.

• Depending on mission/type of

propulsion, prop. system mass

(minert) may be small or large

compared to that of the payload

ln

initial mass

final mass

where:

with

s/c dry or payload mass

propellant mass

inert mass of propulsion system

iSP

f

i

f

i pay inert prop f prop

pay

prop

inert

mV gI

m

m

m

m m m m m m

m

m

m


34

Mission Analysis, 2

• Inert mass large for large ΔV mission (carry a

lot of propellant)

• Inert mass fraction:

• Propellant mass fraction:

• For known ΔV, propellant mass as fn. of finert

• Single Stage to Orbit Calculation (example)

– Assume for a LEO mission ΔV = 7500 m/sec

and Isp = 350 sec

– The Rocket Equation yields the mass ratio

mi/mf = 8.88

– Only 11% of the vehicle is structures &

payload (inert mass fraction)

1

inertinert

prop inert

prop

prop inert

prop inert

mf

m m

mf f

m m

1 1

1

SP

SP

V

gI

pay inert

prop V

gI

inert

m e f

m

f e


35

References

• R.H. Humble, G.N. Henry, W.J, Larson, Space Propulsion Analysis and

Design [SPAD], New York : McGraw-Hill, 1995.

• G.P. Sutton and O. Biblarz, Rocket Propulsion Elements, 8th Ed., John

Wiley & Sons, 2010

• V. Yang, M. Habiballah, J. Hulka and M. Popp, Eds., Liquid Rocket

Thrust Chambers; Aspects of Modeling, Analysis , and Design,

Progress in Astronautics and Aeronautics, Vol. 200, 2004

• D.K. Huzel and D.H. Huang, Modern Engineering for Design of Liquid-

Propellant Rocket Engines, AIAA, 1992

• NASA SP-8080 to NASA SP-8025: reports addressing specific

components

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