Sdac Lec 9 10 Space Propulsion

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IN THE NAME OFALLAH

THE MOST BENEFICENT

THE MOST MERCIFUL



SPACERAFTDYNAMICS AND CONTROL

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[email protected] ; 0321-9595510

DR. QASIM ZEESHANBE, MECHANICAL ENGINEERING

NATIONAL UNIVERSITY OF SCIENCE AND TECHNOLOGY, NUST, PAKISTAN, 2000

MS, FLIGHT VEHICLE DESIGNBEIJING UNIVERSITY OF AERONAUTICS AND ASTRONAUTICS, BUAA, P.R.CHINA, 2006

PhD, FLIGHT VEHICLE DESIGNBEIJING UNIVERSITY OF AERONAUTICS AND ASTRONAUTICS, BUAA, P.R.CHINA, 2009



LECTURE # 9

SPACECRAFTPROPULSION



Introduction

Purpose of the propulsion subsystem

Transfer spacecraft from launch vehicle parkingorbit to spacecraft mission orbit

Maintain and control spacecraft orbit

Maintain and control spacecraft attitude Types of spacecraft propulsion systems

Chemical Liquid, Solid or Hybrid

Solar Electric

Nuclear Thermal or Electric



Introduction

Spacecraft propulsion is any method used to acceleratespacecraft and artificial satellites. There are many differentmethods.

Each method has drawbacks and advantages, and spacecraftpropulsion is an active area of research.

However, most spacecraft today are propelled by forcing agas from the back/rear of the vehicle at very high speedthrough a supersonic de Laval nozzle. This sort of engine iscalled a rocket engine.



Introduction

All current spacecraft use chemical rockets (bipropellant orsolid-fuel) for launch, though some (such as the Pegasus rocketand SpaceShipOne) have used air-breathing engines on their

first stage. Most satellites have simple reliable chemical thrusters (often

monopropellant rockets) or resistojet rockets for orbital station-keeping and some use momentum wheels for attitude control.

Soviet bloc satellites have used electric propulsion for decades,

and newer Western geo-orbiting spacecraft are starting to usethem for north-south stationkeeping and orbit raising.

Interplanetary vehicles mostly use chemical rockets as well,although a few have used ion thrusters and Hall effect thrusters

(two different types of electric propulsion) to great success.



Space Propulsion Applications

Launch Vehicles

Ballistic Missiles

Earth Orbiting Satellites

Upper Stages

Interplanetary Spacecraft

Manned Spaceflight

www.army-technology.com

en.wikipedia.org

www.britannica.

com blog.wired.com

www.psrd.hawaii.edu



Space Propulsion Functions

Primary propulsion

Launch and ascent

Maneuvering Orbit transfer, station keeping, trajectory correction

Auxiliary propulsion

Attitude control

Reaction controlMomentum management

www.nasm.si.edu



INTRO10



INTRO: DEVELOPMENT of ROCKET11

The real inventor of the rocket were Chinese [ Feng Jishen, (970 AD)

]

The invention of ROCKET was the practical result of experiments.

GUN POWDER & BAMBOO TUBES




Salahuddin

Ayyubi



Dardanelles Gun: The Cannon of Mehmud

Used during siege of CONSTANTINOPLE




Kublai Khan used it during his invasion of Japan in 1275

1300s; Rockets were used as bombardment weapons asfar west as Spain, brought west by the Mongol hordes,

and the Arabs

Tipoo Sultan, in the 1770s used against the

British army in India.




The Mysore rockets utilised effectively during the Anglo-Mysore Wars, and were

later updated by the British into the Congreve rockets, which were successively

employed during the Napoleonic wars and the War of 1812.




Goddard

Oberth

Von Braun

TsiolkovskyKorolev

Modern Rocket

Engineers, Mathematicians and Dreamers




Wan Hu

He is said to have attached 47 rockets to a bamboo chair, with the

purpose of ascending into heaven

There was a huge explosion. When the

smoke cleared, Wan and the chair were

gone, and was said never to have been seen

again.




Konstantin Tsiolkovsky (1857-1935)

, a mathematics teacher wrote

about:

Space travel, includingweightlessness and escape

velocity, in 1883

Artificial satellites in 1895.

Derived the rocket equation, and

dealt in detail with the use of rocket

propulsion for space travel;

Described multi-stage rockets in

1924.




Hennan Oberth (1894-1992)

He published his (rejected) doctoral thesis in 1923,

as a book

Examined the use of rockets for space travel

The design of liquid-fuelled engines using alcohol

and liquid oxygen




Robert Goddard (1882-1945) , a professor

Published , A Method of Reaching Extreme Altitudes

Goddard's inventions included the use of gyroscopes for guidance, the use

of vanes in the jet stream to steer the rocket.

First liquid-fuelled rocket from Auburn, Massachusetts, on 16 March 1926.

Goddard mentioned the possibility of sending an unmanned rocket to the

Moon, and for this he was ridiculed by the Press

Because of his rocket experiments he was later thrown out of

Massachusetts by the fire officer

In 1960 the US government bought his patents for two million dollars




Von Braun

Enthusiastic engineers , development of the A4 liquid-fuelled

rocket which became the notorious V2 weapon.

From its launch site in Gennany, carried a 1,000- pound bomb

into the centre of London.

US took von Braun and key members of his team.



INTRO: SPACE PROGRAM31

The Russian space program has

been the most active and focused in

history:

The first artificial satellite

The first man in space

The first spacecraft on the Moon

The first docking of two spacecraft

The first space station




The American space program :

Artificial satellite

Man in space

Man on the Moon

Docking of two spacecraft

International space station

Space Shuttle (24 ton to low Earth orbit)

Reuseablity




The Chinese space program :

NOT a RACE

Chang Zheng, or Long March.

China's first satellite in 1970

China (third in the world) to have launched a man into space



INTRO34

Other programs

Japan, India, and Pakistan

all have space programs



REVISION

LECTURE # 9



SPECIFIC IMPULSE

LECTURE # 9



SPECIFIC IMPULSE

The specific impulse of a rocket-propellant combination is analogous to "miles per

gallon" for an automobile. All other things being equal, the V we can obtain from a

rocket stage is directly proportional to its Specific Impulse.

Consequently, the specific-impulse provides us with a convenient measure of a

rocket's intrinsic efficiency.

Once we have chosen the fuel and the oxidizer to be used, a chemical rocket's

specific impulse is largely determined by the energy contained in its propellants.

The specific impulse of a rocket-propellant combination can be defined us the number

of seconds a pound of the Propellant will produce a pound of thrust, Generally

speaking, rocket scientists strive for the Highest Specific impulse they can achieve.

The specific impulse of A Rocket can be computed by dividing the Thrust it generates

by dot , the rate at which it consumes its propellants:

w

f I sp



SPECIFIC IMPULSE

This eqn is featured in Fig below together with another useful eqn for Specific

Impulse

w

f I sp

o

e sp

g

V I

f



SPECIFIC IMPULSE

The specific impulse of a rocket-propellant combination can be defined as the number of seconds a pound of

propellant can produce a pound of thrust.

In this conceptual diagram, a rocket-powered skateboard is attached with a string to a barber pole that is, inturn. connected to a spring scale.

To measure the specific impulse of the rocket-propellant combination, ignite the rocket and adjust its valves until

it is generating one pound of thrust, as indicated by the spring scale, and then count off the number of secondsduring which its one-pound-propellant load can continuously produce one pound of thrust.

w

f I sp

o

e

sp g

V I

f



SPECIFIC IMPULSE

To measure the specific impulse of a rocket-propellant combination, we first load the rocket with one pound of

propellant (fuel and oxidizer loaded at the proper mixture ratio).

Then we fire the rocket and adjust its propellant valves until the rocket is producing exactly one pound of thrust, as

indicated by the spring scale.

Finally, we count off the number of seconds until the rocket bums exactly one pound of propellant. The number of

seconds that elapses during the burning of one pound of propellant equals the specific impulse of that particular

rocket-propellant combination.

The resulting specific- impulse value, in turn, can be used in the rocket equation to calculate the velocity increment we

can obtain from the rocket when it is loaded with a specific payload and a specific load of propellants.

w

f I sp

o

e

sp g

V I

f



SPECIFIC IMPULSE

When expressed in units of seconds, the specific impulse can be interpreted in thefollowing ways: the impulse divided by the sea-level weight of a unit mass of propellant the time one kilogram of propellant lasts if a force equal to the weight of one kilogram

is produced, for example a hypothetical vehicle hovering over the Earth (imagine the fuel

to be supplied from outside, so that the mass on which the thrust is applied does notreduce by spending fuel) alternatively, for engines that can not produce a large thrust: approximately the time

one kilogram of propellant lasts if an acceleration of 0.01 g of a mass of one 100kilogram is produced

100 times the time an acceleration g can be produced (i.e. a thrust equal to the weighton Earth of the current mass) with a propellant mass of 1 % of the current total mass

(100 times the time it takes in this case to reduce the total mass by 1 %) the time an acceleration g can be produced with a propellant mass of 63.2 % of the

initial total mass (the time it takes in this case to reduce the total mass by a factor e, to36.8 %)

twice the net power to produce an acceleration of 1 m/s2 to a mass which at Earth has aweight of 1 N (i.e. a mass of 102 grams)

w

f I sp

o

e

sp g

V I



ROCKET EQUATION

LECTURE # 9



Tsiovosky

Konstantin Eduardovich Tsiolkovsky (17 September

1857 – 19 September 1935) was an ImperialRussian and Soviet rocket scientist and pioneer of

the astronautic theory.

Along with his followers the German HermannOberth and the American Robert H. Goddard, he

is considered to be one of the founding fathers of

rocketry and astronautics.

His works later inspired leading Soviet rocket

engineers such as Sergey Korolyov and ValentinGlushko and contributed to the success of the

Soviet space program.

http://en.wikipedia.org/wiki/File:Tsiolkovsky.jpg



Tsiovosky

Draft first space ship by Konstantin

Tsiolkovsky

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ROCKET EQUATION

Newton’s second law, manipulated in accordance with simple relationships from

integral calculus, is used in deriving the rocket equation. That famous equation,

which is also called Tsiovosky’s equation, is highlighted at the bottom of this figure.

Tsi- ovosk/s equation indicates that the maximum velocity we can obtain from a

load of propellant is directly proportional to the specific impulse multiplied by thenatural logarithm of the ratio of the weight of the rocket at ignition and the weight

of the rocket at burnout.

The Rocket Equation

Figure includes a simple derivation of the rocket equation, which is also called

Tsiolkovsky’s equation, in honor of the Russian schoolteacher who first derived it nearly 70years ago. Notice that the derivation hinges on the proper interpretation of Newton’s

second law (F = ma), where both the mass of the rocket, w, and its acceleration, a, are

constantly changing. The end result of this step-by-step derivation is the rocket equation:

f



SPECIFIC IMPULSE

w

f I sp

o

e

sp g

V I

W



ROCKET EQUATION

where V is the ideal (maximum) velocity the rocket can generate, I sp is the specific

impulse of the rocket-propellant combination, and W o and W f are the ignitionWeight and burnout weight respectively, at the beginning and end of the rocket

burn.

We can use the rocket equation to calculate the ideal velocity a particular rocketcan generate while burning a particular load ofpropellants.

In the real world, of course, trajectory losses, including gravity losses, drag losses,

and steering losses, must be subtracted from the ideal velocity to obtain a more

realistic estimate of the actual velocity the rocket can produce.

f

o sp

W

W gI V ln



The Tsiolkovsky Equation

mass M velocityV

dt

dM V

dt

V d M

dt

V M d

,

0

)(

Begin with momentum conservation for

An Isolated Body in free space

Exhaust

motion

Rocket

motion

No aerodynamic drag

and no gravity



The Tsiolkovsky Equation Cont.

dt dM V

dt V d M

Mass of the rocket is determined by what lies within

its mechanical envelope and rocket nozzle

Mechanical

envelope

Mass leaves

envelope at Vex

leading to mass

decrease withinenvelope



)ln(

100

I

Rex

t

ex

t

ex

M M V V

rocket of timeburnt

dt dt dM

M V dt

dt V d

dt

dM V

dt

V d M


Mass Initial to Massmaining of Ratio M

M

Nozzletheof out Velocity Exhaust V

VelocityinChange Final V

I

R

ex

Re



)ln(

R

I ex

M

M V V

Equation yTsilokovsk The


In order to maximize the V we must maximize Vex, the

exhaust velocity ( it must be very explosive fuel)

and the ratio of fully fueled or initial mass MI to final or remaining mass MR

( the rocket must be a fuel tank, composed of the lightest

substance possible that can withstand the stress)



Specific Impulse

Exhaust velocity is most

often given in terms of

Specific Impulse =Isp

sec)/5.4sec(450

sec8.9 2

kmat Oxygenand Hydrogen Liquid

is fuelschemical for I Highest

gravitytodueionaccellerat m g

g

V I

sp

ex sp



Analysis :

With time velocity increases

Natural logarithm of the ratio of initial to current mass

Exhaust velocity : How fast the mass is being expelled

The most advanced liquid-fuelled chemical rockets today produce an

exhaust velocity of : 4500m/s





It does not depend on the thrust

To achieve a high rocket velocity, the mass ratio has to be

largeThe mass ratio is defined as the ratio of vehicle-plus-propellant

mass, to vehicle mass

A mass ratio of, say, 5 indicates

80% of the initial mass of the rocket is fuel




Car, which has a typical empty mass of 1.5 ton, and a fuel mass of 40 kg

A mass ratio of 1.003

The rocket can travel faster than the speed of its exhaust.

The point at which the rocket speed exceeds the exhaust speed is

when the mass ratio becomes equal to e, the base of natural

logarithms

Tsiolkovsky calculated how fast a rocket needs to travel to reach

space.

and determined : there was a LIMITA mass ratio of 10 is almost impossible to achieve.



R

Fu elBurned ex

R

FuelBurned ex

R

I ex

R

Fu elBurned

R

Fu elBurned

R

FuelBurned R

R

I

M

M V

M

M V

M

M V V

Becomes EquationThe

M

M where

M

M

M

M M

M

M For

)1ln()ln(

1

1


For the case of small rocket burns on satelites in space,

where the mass of fuel burned is small relative to the satellite mass

The equation becomes simple

fV



SPECIFIC IMPULSE

w

f I sp

o

e

sp g

V I



SPECIFIC IMPULSE

• Specific Impulse rocket’s Ability to deliver a certain

(specific) impulse for a given weight of propellant

I sp Impulse

g 0 M propellant

F thrust 0

t

dt

g 0 m•

propellant dt 0

t

g 0 9.806m

sec2(mks)

Mean specific impulse

• At a constant altitude, with

Constant mass flow through engine

I sp Impulse

g 0 M propellant

F thrust 0

t

dt

g 0 m•

propellant dt

0

t

F thrust

g 0 m•

propellant

• Instantaneous specific impulse



Rocket Equation

V V final V 0 M 0 M final m propellant

V g 0 I sp ln 1m propellant

M final

g 0 I sp ln 1 P mf

P mf "propellant mass fraction"

• Sometimesm propellant

M final

m propellant

Is also called

propellant mass

Fraction or “load mass fraction”

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LECTURE # 9

SPACE PROPULSIONREQUIREMENT






PROPULSION REQUIREMENTS



Typical

Propulsion

Requirements

Maneuver ΔV, km/s



Typical

Propulsion

Requirements

Orbit transfer:

LEO to GEO

LEO to GEO

GTO to GEO (1)

GTO to GEO (2)LEO to Earth escape

LEO to translunar orbit

LEO to lunar orbit

GTO to lunar orbit

LEO to Mars orbit

LEO to solar escape

3.95 (no plane change required)

4.2 (including plane change of

28 deg)

1.5 (no plane change required)

1.8 (incl. plane change of 28

deg.)3.2

3.1

3.9

1.25-1.4

5.7

8.7

Orbit control: Station-keeping

(GEO)50-55 m/s per year

Orbit control: Drag

compensation

•alt.: 400-500 km

•alt.: 500-600 km

•alt.: >600 km

< 100 m/s per year max. (<25

m/s average)

< 25 m/s per year max. (< 5 m/s

average)< 7.5 m/s per year max.

Attitude control: 3-axis control2-6 m/s per year

Auxiliary tasks:

•Spin-up or despin

•Stage or booster separation

•Momentum wheel unloading

5-10 m/s per manoeuvre

5-10 m/s per manoeuvre

2-6 m/s per year



LECTURE # 9

SPACECRAFT PROPULSIONEQUATIONS



Equations

Thrust

Mass Flow

Impulse

Specific Impulse

Ideal Rocket Equation

Propellant Tank Mass



PERFORMANCE

Specific Impulse : Thrust produced per unit weight flow rate



How a Thermodynamic Rocket Works

eaee A p pvm F )(

m = mass flow rate (kg/sec)

ve = propellant exhaust velocity (m/sec)

pe = pressure at nozzle exit (Pa)

pa = ambient pressure (Pa)

Ae = area of nozzle exit (m2)

e

vm F

For Ideal Expansion (pe = pa):



Thrust Equation



Thrust Equation

P2

P

T

PeTe

Ae

F1 F2



Thrust Equation

Change of momentum of flux across surface

of CV

Sum of forces on

CV

Change of momentum of mass contained in

CV



Thrust Equation

No Body Forces

Steady Flow

Uniform flow at nozzle exit

One dimensional flow (only x)

Pe ≥ P2



Thrust Equation

F + Fb = +

F =

F = F1 – F2 F 2 = (pe-p2) Ae



Thrust Equation

F =

ρu ds = mass flow rate

F = m Ve

F = F1 – F2 m Ve = F1 – (pe-p2) Ae

F1 = m Ve + (pe-p2) Ae



Thrust Equation

If

Pressure ThrustMomentum Thrust

Characteristic

Thrust



Thrust Equation

Effective exhaust velocity

Characteristic Velocity



PERFORMANCE

MULTISTAGE ROCKETS

M 0i : The total initial mass of the ith stage prior to firing

including the payload mass

the mass of i, i+1, i+2, n stages.



PERFORMANCE

MULTISTAGE ROCKETS

Discarding, Inert mass (empty fuel tanks etc ) during the flight is

bound to improve the performance

The thrust remains the same, but after the tanks have been dropped

off, the mass of the rocket is smaller, so the acceleration will be

greater

The final velocity of an n stage launch system is the sum of the

velocity gains from each stage.

nn V V V V V ...........321



PERFORMANCE

The mass ratio of the single rocket

F S

P F S

o

M M

M M M R

The rocket is then divided into two rockets, each having half the fuel

and stacked one on top of the other

The first rocket is ignited, and burns until all its fuel is exhausted



PERFORMANCE

P F S

P F S

M M M

M M M R

2

11

The mass ratio of the first rocket

The lower rocket then drops off, and the upper rocket is ignited.

P S

P F S

M M

M M M R

2

12

1

2

1

2



PERFORMANCE

Compare the performance of a single and a two-stage rocket:

oee RvV log

2log1log Rv RvV eeee

Two stage stage

Single stage



PERFORMANCE

Calculation :

A rocket of total mass 100 ton

Mass of spacecraft of 1 ton

Exhaust velocity of 2,700 m/sStructural mass is 10% of the fuel mass.

Single stage

Ro = 9.09



PERFORMANCE

Two stage stage

velocity of the first stage

R1 = 9.09

velocity increment of the second stage

R1 = 8.42



PERFORMANCE

Total velocity = 7342 m/s ;

1383 m/s more than what was achieved for single stage

Perform the same calculation for 3 and 4 stage vehicles with same data

Definitions



Total Impulse is the force imparted integrated over the burning time

Mathematically

Definitions




Assumptions

Cannot be uniform over entire nozzle area

Measurement can be very difficult (velocity profile)1-D behavior is assumed (Uniform over entire area)

Average equivalent velocity at which Propellant is ejected

Mathematically .

m

F g I c s

Definitions



Thrust

)(.

aeeep p Aum F

Momentum Thrust Pressure Thrust

Definitions



Thrust

)(.

aeeep p Aum F Thrust equation

aep p

eum F

.

When

Characteristic Thrust

Definitions



Exhaust velocity

)(.

aeeep p Aum F Thrust equation

..

.

.

)(

m

p p A

m

um

m

F aeee

Dividing by mass flow rate

.

)(

m

p p Auc aee

e


Definitions



Altitude performance of rocket engines

.)(

m

p p Auc aee

e

Variable PartFixed Part

Definitions



Nozzle Exhaust velocity

The propellants burns inside the combustion chamber

The gas produced is heated by the chemical energy of the combustion

The gases expands through the nozzle

The exhaust velocity can be derived by setting the kinetic energy of

the exhaust gas equal to the change in enthalpy of the gas as it

expands through the nozzle.

The process is assumed to be under isentropic conditions

No heat escapes from the gas to the nozzle walls

The exhaust is assumed to behave like a perfect gas



Nozzle Expansion



NOZZLE

Isp versus Power



Lox hydrogen

Nuclear power required

High fuel

mass

ISP(Sec)

Thrust per Unit Power

N/kW

0.1 0.2

1000

50%eff.

500Resisto-jet

Arc-jet

MET

Ion

Thrusters



Orbital transfer: Earth to Mars

• Start in low circular orbit do rocket

burn to give V (add energy)

• This puts one into transfer orbit that

is an ellipse

• At highest point of ellipse

(aphelion) one performs rocket

burn to add V to “circularize” orbit

at Mars(add more energy)

• To return one reverses theprocedure (subtract energy)



LECTURE # 9

SPACECRAFT PROPULSIONFUEL TYPES



Rocket Fuel Types

Solid Fuel Liquid Fuel

(turbo pump fed)

Liquid Fuel

(pressure fed)

oxidizer

fuel

pressurant

Turbo pump

Thrust

chamber



ENERGY SOURCES



ENERGY SOURCES

Energy Source

Chemical Nuclear

Thermal

Electric

Solar

Thermal

Thermal



ENERGY SOURCES

Chemical

SolidPropellant

LiquidPropellant

Mono Propellant

Bi-propellant

Hybrid



LECTURE # 9

SPACECRAFT PROPULSIONSOLID FUEL



Solid Fuel Rockets

Solid Fuel

•Solid fuel rockets are oldest (China in

~200 AD)

•Simple and reliable

•Inexpensive, easy to Launch

•Part of Space Shuttle Booster

•Isp of solid fuel is low

Isp ~ 250 seconds compared to liquid

fuels Isp ~ 300-450 seconds



Solid fuels

•Solid fuels consist of a solid oxidizer plus a fuel example

•Oldest solid fuel Black Powder

•Saltpeter (KNO3) + Carbon+ Sulfur

•Oxidizer : KNO3 ( releases oxygen)

•Fuel : Carbon (burns with oxygen)

•Burn accelerator: Sulfur ( combines with left over potassium,

releasing more oxygen)



SRM



SRM

In solid propellant rockets , the word "motor" is used.

The propellant is contained and stored directly in the combustion

chamber

long-time storage (5 to 20 years).

Motors come in many different types and sizes, varying in thrust

from about 2 N to over 4 million N

Solid propellant rocket motors have been credited with having no

moving parts



SRM

The principal components of a solid rocket motor

1. Propellant grain

2. Igniter

3. Motor case

4. Exhaust nozzle

5. Thrust vector control

6. Mounting provisions



SRM



SRM Classification

Basis of Classification Examples of Classification

Application Satellite boosters, ballistic missiles, sounding

rockets

Diameter 0.025 to 6.6m

Length 0.025 to 45m

Propellant Composite, double-base, Composite-modified

double-base

Case design Steel monolithic, fiber monolithic, segmented

Grain installation Case-bonded, Cartridge-loaded

Grain configuration Cylindrical, Spherical, end burning , 3D



Grain Configuration

Grain Configuration

Grain configuration basically is a geometrical consideration that

is going to impose a certain thrust law, thrust versus time, which

in turn is going to satisfy the ballistic performance requirement

for a specific mission

Grain is the shape of propellant mass inside

the rocket motor. The propellant grain is a

cast, molded, or extruded body, once ignited,

burns on all exposed surfaces to form hot

gases

G i C fi i



Grain Configurations

A variety of grain configurations are available depending upon

the available propellant

Two dimensional as well as three dimensional configuration can

be selected depending on specific requirements

1. Star

2. Slotted tube

3. Wagon wheel

4. Three dimensional grain



LECTURE # 9

SPACECRAFT PROPULSION

LIQUID FUEL

LIQUID ROCKET ENGINE




The mission requirements can be translated into rocket enginerequirements in terms of

Thrust-time profile

Propellants

Number of thrust chambers

Total impulse

Number of restarts

Minimum reliability

Engine masses and their sizes or envelopes





The design of any propulsion system is tailored to fit a specific

application or mission requirement

Application Mission velocity

The desired flight trajectories

orbit transfer Vulnerability

Attitude control torques Duty cycle

Minimum life (during storage or in orbit)

Number of units to be built and delivered. They include

constraints on cost, schedule, operating conditions,

storage conditions, or safety rules





A liquid propellant rocket propulsion system is commonly

called a rocket engine

One or more thrust chambers

One or more tanks to store the propellants

A feed mechanism to force the propellants from the tanks into the thrust

chamber(s)

A power source to furnish the energy for the feed mechanism

Suitable plumbing or piping to transfer the liquids

A structure to transmit the thrust force

Control devices to initiate and regulate the propellant flow and thus the

thrust.





Many different types of rocket engines have been built and flown

Thrust size from less than 0.01 lbf to over 1.75 million pounds

One-time operation or multiple starts (some have over 150,000 restarts)

With or without thrust modulation (called throttling)

Single use or reusable

Arranged as single engines or in clusters of multiple units





The thrust chamber or thruster is the combustion device wherethe liquid propellants

Metered

Injected

AtomizedMixed

Burned to form hot gaseous reaction products

Hot gaseous are accelerated and ejected at a high velocity to

impart a thrust force

A thrust chamber has three major parts:

An injector

A combustion chamber

A nozzle





The propellants, which are the working substance of rocket

engines, constitute the fluid that undergoes chemical and

thermodynamic changes

Oxidizer (liquid oxygen, nitric acid, etc.)

Fuel (gasoline , alcohol, liquid hydrogen, etc.).

Chemical compound or mixture of oxidizer and fuelingredients, capable of self-decomposition





A bipropellant rocket unit has two separate liquid propellants, anoxidizer and a fuel

A monopropellant contains an oxidizing agent and combustiblematter in a single substance

A cold gas propellant is stored at very high pressure, gives a low

performance

A cryogenic propellant is liquefied gas at low temperature, such asliquid oxygen (-183°C) or liquid hydrogen (-253°C).

Li id F l (t b )



Liquid Fuel (turbo pump)

Liquid Fuel

(turbo pump fed)

oxidizer

fuel

Thrust

chamber

•Most commonly used space

launch vehicle

•Highest Isp and MI/MR

•Basic design unchanged since

German V-2

•LOX (liquid oxygen) is common

oxidizer

•Kerosene or Liquid Hydrogen are

common fuels

turbo

pump

Liquid Fuels



•Liquid fuels have lower average molecular

weight exhaust than solid fuels

•They can be burned in thrust chamber of

fixed geometry to maximize performance

•The liquid can be pure cryogenic gases

such as oxygen and hydrogen



LECTURE # 9


HYBRID

HYBRID FUEL



HYBRID FUEL



LECTURE # 9


OTHERS



Gas Acceleration Mechanism

Gas Acceleration Mechanism

Thermal Electrostatic Electromagnetic

PressureElectric

Field

Gasacceleration

Ions acceleration

Magnetic Field

Gasacceleration

ION THRUSTER



ION THRUSTER

Electromagnetic Propulsion



Electromagnetic Propulsion

NUCLEAR



NUCLEAR

While chemical and electric systems are

used for the propulsion of today’s

spacecrafts,

nuclear propulsion is still under study.



LECTURE # 9


ELECTRIC

Electric Propulsion



Electric Propulsion

Electric Propulsion



Electric Propulsion

• Electric propulsion heats or accelerates gases electrically to

achieve higher Isp than possible with chemical combustion

•Electro thermal uses electric arcs to heat gas to welding arc

temperatures

•Ion thrusters electrically accelerate ionized gas to high velocity

Electric Propulsion



Electric Propulsion

Electric Propulsion



Electric Propulsion

ELECTROTHERMAL SYSTEMS




ELECTROMAGNETIC SYSTEMS



ELECTROMAGNETIC SYSTEMS

ELECTROSTATIC



ELECTROSTATIC



ELECTRIC Thrusters



LECTURE # 9


COLD GAS

Cold Gas



Cold Gas

Cold Gas



Cold Gas

Cold gas



g

Cold gas



g

z1

x1

y1

0

II

III

IV

I

(a)

y1

II

I

III

IVX1

1

2 4

3

z1

(b)

II

I

III

IV

45

1

2

4

3

y1

z1

(c)



LECTURE # 9


HOT GAS



Liquid Propellant



q p



LECTURE # 9


HOT GAS

MONOPROPELLANT SYSTEMS




HOT GAS



LECTURE # 9


HOT GAS

BIPROPELLANT SYSTEMS




LECTURE # 9


HOT GAS

PROPELLANT TANKS

Propellant Tanks



Propellant Tanks



Sphere Right circular cylinder

Formulas for surface area and volume of specific tank geometries



Surface area: S

Volume: V

Radius: r

Diameter: D

Height: h

Spherical cap Spheroid

Torus



LECTURE # 9


APPLICATIONS

THRUST LEVEL



High thrust : for launch, missiles etc

Low thrust : for efficient in-space maneuvers

Spacecraft Propulsion



Characteristics ofSpace Propulsion Systems



Impulse bits

Impulse bit is the smallest change in momentum required to allow for e.g. fine

attitude and orbit control of a spacecraft.

Storable propellants

Storable Propellants are liquid (or gaseous) at ambient temperature and can

be stored for long periods in sealed tanks, e.g. monopropellant hydrazine

In contrast, cryogenic propellants, which are liquefied gases at low

temperature, such as liquid oxygen (-147 °C) or liquid hydrogen (-253 °C) are

difficult to be used for long space flight missions.

Note: at present only storable propellants are used for space flight

missions.




Impulse bits

Impulse bit is the smallest change in momentum required to allow for e.g. fine

attitude and orbit control of a spacecraft.

Storable propellants

Storable Propellants are liquid (or gaseous) at ambient temperature and can

be stored for long periods in sealed tanks, e.g. monopropellant hydrazine

In contrast, cryogenic propellants, which are liquefied gases at low

temperature, such as liquid oxygen (-147 °C) or liquid hydrogen (-253 °C) are

difficult to be used for long space flight missions.

Note: at present only storable propellants are used for space flight

missions.

INTELSAT V



Propellant mass of 168.9 kg required

Transfer orbit (7 kg) Spin up, reorientation

Drift orbit (29.9 kg) Reorientation, spin down

GEO (132 kg) NS Station Keeping (106 kg)

EW Station Keeping (11.7 kg)

Attitude Maintenance (12.3 kg)

Disposal (2 kg)

INTELSAT V



ORBIT TRANSFER INTELSAT V satellite has a Thiokol AKM that produces an average thrust of 56

kN(12,500 lbf) and burns to depletion in approximately 45 seconds.

STATIONKEEPING AND ATTITUDE CONTROL

Array of four 0.44 N (0.1 lbf) thrusters for roll control,

Array of ten 2.0 N (0.45 lbf) thrusters for pitch and yaw control and E/Wstationkeeping,

Array of two 22.2 N (5.0 lbf) thrusters for repositioning and reorientation.

Four 0.3 N (0.07 lbf) EHTs are used for N/S stationkeeping.

The nominal mass of the spacecraft at beginning of life (BOL) is 1005kg and the dry mass at end of life (EOL) is 836 kg. The difference of169 kg represents the mass of the propellant for a design life of 7years.



LECTURE # 9

SPACECRAFT PROPULSIONCHEMICAL PROPULSION

Chemical Propulsion



Chemical reaction produces energy Liquid

Bipropellant Two reactants

Fuel and Oxidizer

MMH, UDMH, O2, HNO3, N2O4

Monopropellant Single reactant

Catalyst

N2H4, H2O2

Solid Fuel and oxidizer combined in a solid mixture (grain)

Hybrid Typically a solid fuel and a liquid or gaseous oxidizer



LECTURE # 9

SPACECRAFT PROPULSIONSOLAR ELECTRIC PROPULSION

Solar Electric Propulsion



Several Classifications

Electrothermal Resistojet

Arcjet

Electrostatic Ion engine

Electromagnetic Pulsed Plasma Thruster

Hall Effect Thruster MPD



LECTURE # 9

SPACECRAFT PROPULSIONNUCLEAR PROPULSION

Nuclear Propulsion



Fission or Radioactive Isotope Decay

Nuclear Thermal Propulsion (NTP)

Transfers heat produced by nuclear process into propellant

gas Propellant heating increases thrust and specific impulse

Nuclear Electric Propulsion (NEP)

Uses heat produced by nuclear process to produce electric

power

Electric power used to ionize and accelerate propellant



LECTURE # 9

SPACECRAFT PROPULSIONPROPELLANT LESS PROPULSION

Propellant-less Propulsion



Solar Sail

Uses solar pressure to generate thrust

Large, reflective surface area required

Electrodynamic TetherUses Earth’s (or other planet’s) magnetic field to

generate a force (with an electric current) Atmospheric Drag

AerobrakingRe-entry




LECTURE # 9

SOLAR SAILS

SOLAR SAILS





Small solar sails, (devices that produce thrust as a reaction force induced by

reflecting incident light) may be used to make small attitude control and

velocity adjustments.

This application can save large amounts of fuel on a long-duration missionby producing control moments without fuel expenditure.

For example, Mariner 10 adjusted its attitude using its solar cells andantennas as small solar sails

SOLAR SAILS



Mariner 10 was an American robotic space

probe launched by NASA on November 3,

1973, to fly by the planets Mercury and Venus.

Mariner 10 was launched approximately two

years after Mariner 9 and was the last

spacecraft in the Mariner program (Mariner 11

and 12 were allocated to the Voyager

program and redesignated Voyager 1 and

Voyager 2).

The mission objectives were to measure

Mercury's environment, atmosphere, surface,

and body characteristics and to make similar

investigations of Venus. Secondary objectives

were to perform experiments in the

interplanetary medium and to obtain

experience with a dual-planet gravity assist

mission.

There currently is a spacecraft mission doing a

more in-depth survey of Mercury, MESSENGER.

The planning of the mission was dependent on

Mariner 10's data sets.

http://nssdc.gsfc.nasa.gov/nmc/spacecraft

Display.do?id=1973-085A

http://history.nasa.gov/SP-423/mariner.htm

MARINER 10

http://nssdc.gsfc.nasa.gov/nmc/spacecraftDisplay.do?id=1973-085A












MARINER 10



MARINER 10

Mariner 10 image showing a scarp on the plains of

Mercury. The SW-NE trending scarp, which may have

been formed by compressional stresses is radial to



been formed by compressional stresses, is radial to

the Caloris Basin, to the southwest (north is up). Theimage is about 240 km across. (Mariner 10, Atlas of

Mercury,

Location & Time Information

Date/Time (UT): 1974-03-30

Distance/Range (km): N/A

Central Latitude/Longitude (deg): +52.50/171.00

Orbit(s): Flyby

Imaging Information

Area or Feature Type: scarp, crater, plains

Instrument: GEC 1-inch vidicon tube (TV) camera

Instrument Resolution (pixels): 700 x 832, 8 bit

Instrument Field of View (deg): 0.38 x 0.47

Filter: N/A

Illumination Incidence Angle (deg): N/A

Phase Angle (deg): N/A

Instrument Look Direction: N/ASurface Emission Angle (deg): N/A

MARINER 10

Mariner 10 mosaic of Mercury taken as the spacecraft

was outbound after the first flyby. This mosaic was

made up of 18 pictures taken about 6 hours after



made up of 18 pictures, taken about 6 hours after

closest approach. The north pole is at the top, and theequator is about two-thirds down from the top. The

resolution is about 2 km. Half of the Caloris Basin is

visible at the terminator on the left, just above the

middle of the image. The bright ray crater at the upper

right is the 45 km diameter Degas

Location & Time Information

Date/Time (UT): 1974-03-30

Distance/Range (km): 200,000Central Latitude/Longitude (deg): N/A

Orbit(s): Flyby

Imaging Information

Area or Feature Type: Global view

Instrument: GEC 1-inch vidicon tube (TV) camera

Instrument Resolution (pixels): 700 x 832, 8 bit

Instrument Field of View (deg): 0.38 x 0.47

Filter: N/AIllumination Incidence Angle (deg): 75

Phase Angle (deg): 75

Instrument Look Direction: N/A

Surface Emission Angle (deg): 0

SOLAR SAILS



Solar sails (also called light sails or photon sails)are a form of spacecraft propulsion using theradiation pressure (also called solar pressure) ofa combination of light and high speed ejectedgasses from a star to push large ultra-thin

mirrors to high speeds. Light sails could also be driven by energy beams

to extend their range of operations, which isstrictly beam sailing rather than solar sailing.

Solar sail craft offer the possibility of low-costoperations combined with long operatinglifetimes.

Since they have few moving parts and use nopropellant, they can potentially be usednumerous times for delivery of payloads.



LECTURE # 9

SPACE PROPULSION

CASE STUDY

SPACE PROPULSION – CASE STUDY






Summarizes key features and performance characteristics of existing andplanned (near future) propulsion systems for use on spacecraft such as

satellites

This study demonstrates the design of a baseline orbital propulsion systemfor a spacecraft with a total loaded mass of 5000 kg (including the

propulsion subsystem). In more detail, the propulsion system should allow

for the spacecraft to perform a coplanar LEO-GEO orbit transfer mission.

Three options are considered:

1. Minimum energy, high thrust (Impulsive shot)

2. Low thrust chemical

3. Low thrust (spiral transfer)




The main difference is in the velocity change required, the thrust level and themission duration.

We determine subsequently the various propulsion options for use, their main

advantages and disadvantages with respect to amongst others functionality,operation and cost, and calculate propulsion subsystem and net vehicle mass

(vehicle mass excluding propulsion system).

The results are compared and the various systems assessed for their missionsuitability.

No attempts are made to determine the effects of varying total vehicle mass

on the outcome of this assessment. Also we neglect the need for a margin on required velocity change to account

for mission uncertainties.




Propulsion Options

The following propulsion options are considered as candidate for the mission at

hand:

Chemical:

a. Liquid bipropellant i. Cryogenic

ii. Storable

b. Solid

c. Hybrid

Other propulsion options like chemical monopropellant, laser-thermal, nuclear-

thermal, plasma, and tethered propulsion are left out of consideration to limit

the amount of analysis to be performed.




System Concept Analysis and Design

Many missions desire minimal mass to reduce the mission cost, which may be

achieved in part by choosing the appropriate propulsion system.

Some military missions and manned missions, however, may prefer minimal TOF over

minimal cost associated with minimal mass.

At times, the choices for the orbit transfer and propulsion system may be obvious,

but, the appropriate design is not obvious, given that designing a propulsion system

depends on the transfer orbit chosen, and vice versa.

For such design problem, a tool that designs the orbit transfer and propulsion

system may alleviate the difficulty in solving the coupled problem





We could go through a detailed analysis to determine the mass of propulsion

system, but, it becomes very complex and tedious.

There are many detailed design choices required to compute the propulsion system

mass, we take a very simple approach to approximating the propulsion system

mass.

To develop a propulsion budget based on a given V budget, for PreliminaryDesign, we can estimate the cost of the space mission by using the rocket equation

to determine the total required spacecraft plus propellant mass, in terms of the dry

mass of the spacecraft, the total required V

)ln(lnln R gI m

m gI

mm

m gI V sp

f

o sp

po

o sp

1

)/(

g I V

f p spemm





mf mo - mp is the Final Vehicle Mass

mo = Initial Vehicle Mass

mp =Mass of the Propellant Consumed

R =Mass ratio.

It assumes zero losses due to Gravity and Drag, and is thus the Limiting Ideal case.In practice the V achieved will be somewhat smaller

)ln(lnln R gI m

m gI

mm

m gI V sp

f

o sp

po

o sp

1

)/(

g I V

f p spemm





Isp is the most important and first figure of merit.

Note that the Isp is dependant on the nature of propellant, nozzle design, ambient

pressure, and combustion efficiency.

Given a structural mass fraction rs for the propulsion system, the total mass of the

propulsion system is

)ln(lnln R gI m

m gI

mm

m gI V sp

f

o sp

po

o sp

1

)/(

g I V

f p spemm

s

p

total r

mm

1





To evaluate Empty Vehicle Mass (EVM) a 1% ~ 5% margins areconsidered to account for ullage, propellant boil off in case of liquidand sliver in case of solid propellant.

In this section, the various propulsion options available are analyzedto a level of detail considered fit for conceptual analysis anddesign. Table provides an overview of the various options as well astheir typical performances, and advantages and disadvantages with

respect to amongst others thrust control, restartability, reliability,flight status and cost.

The information in the table is not considered to be all inclusive, butis given for the reason of demonstrating how propulsion options maybe mapped and characterized in a comparative way.

Courtesy Barry Zandbergen, TU Delft

)ln(lnln R gI m

m gI

mm

m gI V sp

f

o sp

po

o sp

1

)/(

g I V

f p spemm

s

p

total r

mm

1





Various options are analysed with respect to their effect on totalpropulsion system mass.

We will select appropriate values for thrust and whenapplicable also

Propellant

specific impulse

oxidiser-to-fuel ratio

propellant density

tank pressure, sail material, etc.

and perform an initial sizing of the system.

The numbers are not optimized to obtain the best performance.

)ln(lnln R gI m

m gI

mm

m gI V sp

f

o sp

po

o sp

1

)/(

g I V

f p spemm

s

p

total r

mm

1


No System ConceptTypical (vacuum)

performances

Advantage/

DisadvantageRemark

1

Cryogenic

chemical

bipropellant

Propellant: Liquid hydrogen &

liquid oxygen Isp: 400-450 sec

Typical oxidiser/fuel mass ratio

is 5. Thrust level: High (~kN)

Advantages: • Space qualified • Used today mostly for

launcher stages including some upper stages • Extensive in

flight heritage • Highest average Isp with respect to other pure

chemical systems Disadvantages: • Expensive and complex

compared to other chemical systems due to low temperature

storage of propellants (hydrogen at 20 K and oxygen at 80 K) •

high g-load on s/c due to relatively high thrust (necessary to

reduce storage time)

Presently only used for

short duration launcher

missions (Ariane 5, Space

Shuttle, Centaur)

P ll t MMH/h d iAdvantages: • Space qualified • Extensive in flight heritage •



2Storable chemical

bipropellant

Propellant: MMH/hydrazine

&N2O4 Isp: 300-340 sec. Typical

oxidiser/fuel mass ratio is 1.65.

Thrust level: Moderate thrust

(~400 N)

g p q g g

Propellants can be stored for longer periods of time (up to manyyears) • Average Isp compared to other pure chemical systems •

Hydrazine can also be used as monopropellant offering an

increase in reliability Disadvantages: • Expensive and complex

compared to solid and hybrid systems • Moderate g-load on s/c

Used for almost all current

GEO communication

satellites

3Solid chemical

propellant

Propellant: Composites +

aluminium Isp: 285-300 sec.

Thrust level: High (~kN)

Advantages: • Space qualified • Extensive in flight heritage •

Low cost • High reliability Disadvantages: • Limited

performance (Isp) • Not restartable • inaccurate injection •

high g-load on s/c due to relatively high thrust

Has been used in the very

beginning of GEO

commerce. Only used

seldom today

4Hybrid chemical

propellant

Propellant: Aluminized poly-

ethene & N2O4 Isp: 310-320

sec. Typical oxidiserto fuel mass

ratio is 2. Thrust level: High

(~kN)

Advantages: • Restartable • Cost in between those of solid

and liquid chemical systems • Reliability in between reliability

of solid and pure liquid systems Disadvantages: • Not space

qualified • Little flight heritage • inaccurate injection (not

staged) • high g-load on s/c

Low-cost replacement for

liquid systems.

5 Solar thermalPropellant: Liquid hydrogen Isp:

800 sec Thrust level: ~100N

Advantages • High performance Disadvantages • Low

development status (first flight planned by in 2004, but no

definite plans have emerged) • Requires separate power source

to provide thermal power

Candidate system for LEO-

GEO transfer missions and

interplanetary cargo

missions

6 Ion propulsion

Propellant: Xenon gas Isp:

Better than 2000 sec Thrustefficiency: 75% Thrust level:

Thrusters of 200 mN are

available.

Advantages • Space qualified • Very high performance •

Control Disadvantages • Full transfer takes long time (> 1 year)due to low thrust level • Little heritage in low thrust transfer

mission. • Autonomous transfer necessary • High cost •

Requires separate power source to provide electrical power

Candidate system forinterplanetary cargo

missions

7 Solar sailingSail (assembly) loading of 2.5

gram/square meter. Thrust

level: ~0.9 |iN/m2 at 1 AU

Advantages: • Requires no propellant mass • Essentially

unlimited operation time • Low acceleration loads due to low

thrust (~0.9 |iN/m2) Disadvantages: • No flight heritage •

Relatively long flight times due to low acceleration levels •

Complex system due to large sail and need for sail steering

Candidate system for

interplanetary cargo

missions and interstellar

travel

PropertyLiquid

Cryogenic

Liquid

Storable



y g

Solid HybridV [km/s] 3.895 3.950 3.895 3.895

Thrust level 60 kN 1.6 kN 50 ~ 20 kN 50 kN

Effective exhaust velocity; Isp*go [m/s] 4500 3250 3000 3250

On board propellant mass [kg] 2983 3552 3817 3666

Propulsion system dry mass [kg] 377 426 382 424Propulsion system total mass [kg] 3360 3978 4199 4080

Vehicle dry mass [kg] 2017 1448 1183 1334

Vehicle dry mass excl. propulsion system [kg] 1640 1022 801 910

Propellant storage volume [m3] 9.60 3.4 2.62 3.12

Acceleration (end of mission) [go

] 3.2 0.11 2.3 3.8

Burn time ~3.5min 120 min ~3.5 min ~4 min.

Time in transfer orbit 5-6 hours Several days 5-6 hours 5-6 hours

Property

Liquid

Cryogeni

c

Liquid

StorableSolid Hybrid

Solar

Thermal

Ion

Prop

Solar

Sail

V [km/s] 3.895 3.950 3.895 3.895 4.611 4710 4710

Thrust level 60 kN 1.6 kN 50 ~ 20 kN 50 kN 100 N 1.4 N 0.56



Effective exhaust velocity; Isp*go [m/s] 4500 3250 3000 3250 8000 20000 -

On board propellant mass [kg] 2983 3552 3817 3666 2300 1060 -

Propulsion system dry mass [kg] 377 426 382 424 2180 2181 -

Propulsion system total mass [kg] 3360 3978 4199 4080 4480 3241 2500

Vehicle dry mass [kg] 2017 1448 1183 1334 2700 3940 5000

Vehicle dry mass excl. propulsion system

[kg]1640 1022 801 910 520 1759 2500

Propellant storage volume [m3] 9.60 3.4 2.62 3.12 33.3 1.96 -

Collector/solar array/sail area [m2] - - - - 396.8 95.2 1x106

Acceleration (end of mission) [go] 3.2 0.11 2.3 3.8 0.002 0.00003 0.000016

Burn time ~3.5min 120 min ~3.5 min ~4 min. ~45 hour ~173 days ~1 year

Time in transfer orbit 5-6 hoursSeveral

days5-6 hours

5-6

hours

Multiple

days

~173 days

(~1/2 year)~2 year

Characteristics PAM-D PAM-DII TOS IUS Centaur H-10 D-M L-9

Stage: Manufacturer

Boeing Boeing Lockheed Martin Boeing

Lockheed

Martin Ariane-space RSC Energia Ariane-spaceLength (m) 2.04 2.00 3.30 5.20 9.0 9.9 6.8 4.5

Diameter (m) 1.25 1.62 3.44 2.90 4.3 2.6 4.1 5.4

Engine: ManufacturerThiokol Thiokol CSD CSD

Pratt-

WhitneySNECMA Isayev DASA

Type (Star 48) ISTP SRM-1 SRM-1, SRM-2 RL 10A-3-3A HM7B 11 DM 58 Aestus

Number 1 1 1 1, 1 2 1 1 1

Fuel Solid Solid Solid Solid LO2LH2 LO2LH2 LOX RP1 N2O4/ MMH



Composition TP-H-3340 — HTPB HTPB 5.5:1 4.77 2.6 2.05Total Thrust (N)

66,440 78,300 200,000200,00081,200

147,000 62,700 84,000 29,000

Specific Impulse (s)292.6 281.7 294

292.9

300.9442 444.2 361 324

Burn Time (s)54.8 121 150

153.0

104.0488 725 680 1100

Stage: Pad Mass (kg) 2,180 3,490 10,800 14,865 18,800 12,100 18,400 10,900

Impulse Propel. Mass (kg) 2,000 3,240 9,710 9,7102,750

16,700 10,800 15,050 9,700

Burnout Mass (kg)189 250 1,090

1,255

1,1502,100 1,300 2,140 1,200

Airborne Support Equip. Mass

(kg)1,140 1,600 1,450 3,350 4,310 — — —

Illustration:

Schedule: Start Date 1975 1980 1983 1978 1982 1986 New 1988

Operational Date 1982 1985 1986 1982 1990 New 1996

Type of Development Commercial Commercial Commercial Gov't Gov't ESA ILS ESASponsor Boeing Boeing OSC USAF USAF ESA ILS ESA

CSD — Chemical System Division, United Technologies; ESA — European Space Agency; ILS — International Launch Services; OSC — Orbital Sciences Corporation;

RSC — Rocket Space Corporation; DASA — Daimler Chrysler Aerospace; SNECMA — Socie'te' Nationale d'Etudes et de Constructions de Moteurs d'Avion




LECTURE # 9

SPACE PROPULSION

COMPARISON

COMPARISON






COMPARISON



The above curves of m PS /m S/C, plotted as a function of Δv gives the first and

most important indication for the pre-selection of propulsion systems. If we

assume m PS /m S/C < 0.30, we can read directly from the curves of m PS /m S/C :

for low Δv ≤ 150 m/s, compressed cold gas and vaporising liquid propulsion

systems seem to be the best choice, because they meet the requirement andhave the lowest cost;

for 150 < Δv ≤ 650 m/s, monopropellant hydrazine fed propulsion systems are

the best choice, because of their inherent simplicity (reliability) and potential low

cost, while still meeting the requirement;

for high Δv > 650 m/s, bipropellant systems, monopropellant hydrazine fedresistojet systems (power-augmented thrusters, arcjets), and electrostatic

(electromagnetic) systems will satisfy the Δv-requirements best.




LECTURE # 9

SPACE PROPULSION

FUTURE REQUIREMENTS






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