Missiles Propulsion 1

8/10/2019 Missiles Propulsion 1

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Rocket Propulsion

Guided Weapons Systems

MSc Course

Missile Propulsion


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Rocket Thrust

Rockets are reaction engines.

Operating principle based on Newtons laws

of motion.

2nd law- rate of change of momentum isproportional to applied thrust (i.e. F = m x

a)

3rd law- every action has an equal and

opposite reaction.


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Conservation of Momentum

Example A spaceman of mass

80 kg throws a ball of

mass 0.4 kg forwardsat 20 m/s.

The spaceman willthen move

backwards at avelocity of (0.4 / 80) x20 = 0.1 m/s


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Rocket Thrust

Rocket ejects mass at a given momentumratefrom the nozzle and receives a thrustin

the opposite direction.

Momentum rate = x Ue= thrustWhere = propellant mass flow rate (kg/s)

Ue = exhaust velocity (m/s).

There may also be a thrust component due

to pressure field in nozzle (see later). Thrust may be increased by either increasing

propellant flow rate or exhaust velocity.

pmpm


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Rocket Principles

High pressure/temperature/velocityexhaust gases provided through

combustion and expansion through

nozzle of suitable fuel and oxidisermixture.

A rocket carries both the fueland

oxidiseronboard the vehicle whereasan air-breatherengine (e.g. turbojet or

turbofan) takes in its oxygen supply

from the atmosphere.


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History of Rockets

First reaction enginesoriginated in Greece,

around 400 BC, using

steam. Followed by the

aeolipile, designed by

Hero of Alexandria in

about 100 BC.


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Military History of Rockets

First military use ofproper rockets was byChinese in 1232 in Battle

of Kai-Keng v Mongols. Used gunpowder(saltpeter, sulphur,charcoal mixture) to fill

capped bamboo tubesattached to arrows -known as fire arrows.


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History of Rockets (Cont.)

Mongols then produced rockets of their ownand use spread across Europe via Arabs.

In England, Roger Baconimproved

gunpowder mixture to greatly increase

range.

In France, Jean Froissantimproved flight

accuracy by tube-launching (forerunner of

bazooka). In Italy, Joanes de Fontanadesigned surface-runningtorpedo to attack ships.


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History of Rockets (Cont.) By 16th century, rockets were only used forfireworks, though one breakthrough was

made by German Johann Schmidlap.

He was the first to use staging- a firework

with a large sky rocket (1st stage) jettisoned

after burn-out with a smaller 2nd stage

going to a higher altitude. Basis behind all of todays space rockets.


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By late 17th century, Newtons laws werebeing applied to rockets.

German and Russian rocket experimentersbuilt powerful rockets with masses above45 kg.

Military use again by Indian army in 1792 &1799 against British.

Led to British use, designs by Col WilliamCongreveused by British ships v FortMcHenry in war of 1814 (rockets redglare in Star-Spangled Banner).


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Rocket inaccuracy continued to be a bigbugbear but was significantly improved due

to Englands William Halesdiscovery of

spin stabilisation- using the exhaust gas tostrike small vanes and give the rocket spin.

Advances in breech-loaded cannon with

rifled barrels and exploding warheads (e.g.

by Prussians v Austrians) led to anotherdemise in military rocket use.


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Modern Rocketry

Probably began with Russias KonstantinTsiolkovsky(1857-1935) who proposed

idea of space exploration by rockets in

1903! Suggested use of liquid propellants for

increased range and stated that speed and

range were limited only by jet velocity of

escaping gas.

Also came up with mathematical range

equations (see later).


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Modern Rocketry (Cont.)

Next major pioneering work done by RobertGoddard(1882-1945) in USA, conducting

practical rocket experiments.

Began with solid propellantrockets in 1915 but then

produced worlds first liquid

propellant rocket in 1926

(liquid oxygen and gasoline).

Later improvements: gyroscope for flight

control, payload compartment and parachute

recovery.


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Modern Rocketry (Cont.)

Followed by Herman Oberth(1894-1989) inTransylvania.

Published an important book on the use of

rockets for space travel in 1923. His work led to further military development

of the rocket in the form of the infamous

German V-2 (known as A-4 in Germany),

used against London in WW2.


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V-2

Programme directed by Wernher Von Braun. Burnt mixture of liquid oxygen and alcohol at

rate of 130 kg/s for about a 70 s to develop

maximum thrust of about 725 kN - ballistic

coast to target.

Introduced too late to change outcome of war

but led to swift development of ICBMs.


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V-2 (Cont.)

Maximum speed - approx 1340 m/s.

Impact velocity - approx 1100 m/s (> Mach 3).

Typical range/altitude of 350/90 km

respectively. Carried 1 ton explosive warhead.

Launch mass about 13000 kg, impact mass

about 4040 kg. Length 14 m Diameter 1.65

m
http://www.v2rocket.com/start/makeup/v2_side_cutaway.jpg


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V-2 Propulsion System


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Ballistic Missiles

V-2 technology developedafter WW2 into ballistic

missile applications with

German rocket engineersworking on both US and

USSR programmes.

Eventually came ICBMs,

many also serving asspace launch vehicles

(e.g. Soviet R-7 and US

Atlas).

R-7 Sapwood ICBM


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Some US Ballistic Missiles

Missile Launch

Mass

(t)

Propellant Range

(km)

Deployed

Redstone

27 Liquid 400 1959Atlas 120 Liquid 14,000 1959

Titan 2 150 2 stage

liquid

15,000 1963

Minuteman

234 3 stage solid 12,500 1966

Polaris 14 2 stage solid 4,600 1964

Trident 59 3 stage solid 12,000 1990


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Criteria of Performance

Many have been described previously inPropulsion Parameterssection of course.

Covered in more detail here and specific torockets only.

Includes: thrust

specific impulse

total impulse effective exhaust velocity

thrust coefficient

characteristic velocity


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Thrust (F)

For a rocket engine:

Where:

= propellant mass flow rate

pe= exit pressure, pa= ambient pressure

ue= exit plane velocity, Ae= exit area

m

e e a eF mu p p A (1)


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Specific Impulse (I or Isp)

The ratio of thrust / propellant mass flow rate is used todefine a rocketsspecific impulse- best measure of overall

performance of rocket motor.

In SI terms, the units of I are m/s or Ns/kg.

In the US:

with units of seconds - multiply by g (i.e. 9.80665 m/s2)

in order to obtain SI units of m/s or Ns/kg.

Losses mean typical values are 92% to 98% of ideal values.

/spI F m

/spI F mg

(2)


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Total Impulse (Itot)

Defined as:

where tb= time of burning

If F is constant during burn:

0

bt

totalI Fdt (3a)

Thrusttime of burningtotal m b

I F t (3b)


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Total Impulse (Itot) (Cont.)

Thus the same total impulse may be obtained byeither:

high F, short tb (usually preferable), or

low F, long tb

Also, for constant propellant consumption rate:

(3c)

specific impulse total mass of propellant consumed

m

total b

F

I mtm


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Effective Exhaust Velocity (c)

Convenient to define an effective exhaust velocity (c),

where:

The terms effective exhaust velocityandspecific

impulseare therefore synonomous.

From equation (1) it can then be shown that:

F mc F

c Im

e a ee

p p Ac u

m

(4c)

(4b)(4a)


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Thrust Coefficient (CF)

Defined as:

where pc= combustion chamber pressure,

At= nozzle throat area

Depends primarily on (pc/pa) so a good

indicator of nozzleperformancedominated by

pressure ratio.

F

c t

FC

p A (5b)


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Characteristic Velocity (c*)

Defined as:

Calculated from standard test data.

It is independent of nozzle performance

and is therefore used as a measure ofcombustionefficiencydominated by Tc

(combustion chamber temperature).

* c tp Acm

(6)


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Thermodynamic Performance

of Rocket Engines Parameters mentioned above now covered in

greater depth, using following simplifyingassumptions:

combustion gases obey perfect gas laws. constant specific heat for combustion gases.

1-D flow.

no frictional losses.

no heat transfer to walls. combustion complete before gas enters nozzle.

process steady with respect to time.


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- Thrust Parameters affecting thrust are primarily:

mass flow rate

exhaust velocity

exhaust pressure nozzle exit area
http://www.fas.org/man/dod-101/sys/missile/agm-119-980721-N-5961C-001.jpg


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- ThrustMass flow rate

Most easily evaluated at throat, whereconditions will always be chokedand M = 1.

Substituting A = Atand M = 1 into GD eq (13):

i.e.

1

2 1

1

11

2

c

t c

m T

A p R

1

2 12

1t c

c

m A pRT

(7)


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- ThrustExhaust velocity

Several relationships may be derived:

(8a)

1 2

2

211

2

ee o

e

Mu RT

M

1

2 11

o ee

o

RT pup

(8b)

1 1

2 1 2 1e ee p oo o

p pu c T Q

p p

1

21

1

c ee

o

T pRu

M p

(8c)

(8d)


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- Thrust Equations (7) and (8) may thus be used to

obtain a useful overall equation for the rocketthrust:

(9)

1 1

2 1

22 11 1

c et c e a e

c o

RT pF A p p p ART p

1 211

2 12 211 1

et c e a e

o

pF A p p p Ap


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- Specific Impulse For a fully-expanded condition:

If not perfectly-expanded then I also

dependent upon Aeand pa.

(10)

1

21

1

c e

o

T pRI

M p


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- Specific ImpulseInfluence of

Pressure Ratio

&


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- Specific ImpulseVariable Parameters - Observations

Strong pressure ratio effect - but rapidly diminishing

returns after about 30:1.

High Tcvalue desirable for high I - but gives problemswith heat transfer into case walls and dissociation of

combustion productspractical limit between about

2750 and 3500 K, depending on propellant.

Low value of molecular weight desirablefavouringuse of hydrogen-based fuels.

Low values of desirable.


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- Thrust Coefficient May be theoretically represented as:

Thus independent of combustion temperature

and propellant composition.

mainly a function of pressure ratio and closelycontrolled by nozzle conditions therefore auseful measure of nozzle performance.

1 211

2 12 21

1 1

e e a eF

o c c t

p p p AC

p p p A

(11)


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Maximum thrust when exhausting into a vacuum

(e.g. in space), when: max

11 22 2 12 2

1 1FC

(11a)


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- Observations

More desirable to run a rocket under-expanded (to

left of optimum line) rather than over-expanded.

Uses shorter nozzle with reduced weight and

size.

Increasing pressure ratio improves performance

but improvements diminish above about 30/1.

Large nozzle exit area required at high pressure

ratiosimplications for space applications.


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- Characteristic Velocity May be shown to be theoretically represented

by:

Thus, in contrast to thrust coefficient, isindependent of pressure ratio but isdependent on chamber temperature.

Therefore used as indicator of combustionefficiency.

(12)

1 1

2 1 2 1* 1 1

2 2

c oRT ac


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Actual Rocket Performance

Performance may be affected by any of thefollowing deviations to simplifyingassumptions:

Properties of products of combustion vary with

static temperature and thus position in nozzle. Specific heats of combustion products vary with

temperature.

Non-isentropic flow in nozzle.

Heat loss to case and nozzle walls. Pressure drop in combustion chamber due to heatrelease.

Power required for pumping liquid propellants.

Suspended particles present in exhaust gas.


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Internal Ballistics

Liquid propellant enginesstore fuel andoxidiser separately - then introduced into

combustion chamber.

Solid propellant motorsuse propellant

mixture containing all material required forcombustion.

Majority of modern GW use solid propellant

rocket motors, mainly due to simplicity and

storage advantages. Internal ballisticsis study of combustion

process of solid propellant.


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Solid Propellant Combustion

Combustion chamber is highpressure tank containingpropellant charge at whosesurface burning occurs.

No arrangement made for itscontrolcharge ignited and left toitself so must self-regulateto

avoid explosion. Certain measure of control

provided by charge and

combustion chamber design and


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Solid Propellant Combustion

Fundamental property of combustion processis burn rate.

Burning recedes linearly in direction

perpendicular to surface by parallel layers,

sometimes known as rate of regression

(usually measured in mm/s)constant for

given charge under set conditions.


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Propellant Burn Rate

Propellant burn rate (r) is determined empirically fromburning of small slabs under standard conditions:

Propellant temperature = 294 K

Chamber pressure = 68.95 bar Mass flow rate from combustion given by:

Where: Ab= burning area and p= propellant density

b pm A r (13)


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Propellant Burn Rate

Burn rate (r) of the solid propellant is a function of:

Propellant composition

Combustion chamber conditions

combustion pressure on propellant (pc)

propellant initial temperature (Tp)

velocity of gaseous combustion products

combustion gas temperature

time since start of burn

motor motion


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Burning Rate versus

Combustion Pressure

These graphs can be approximated by:

n

cr ap (14)

where:

r = burning rate

pc = combustion pressure

a = empirical constant(influenced by Tp)

n = burn rate exponent


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Burning Rate versus

Combustion Pressure

Typical Double

Base

Propellant Burn

RateCharacteristics


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Self-Regulation of Combustion

Intersection of burn rate and propellant exit mass flow rate

curves gives equilibrium combustion pressure.

With n < 1, the combustion processself regulates, the lower

the value of n the more stable is the process.

With n > 1, the system will explode!


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Effect of Nozzle Throat Area (At)

on Combustion Stability

pc

r

n


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Propellant Area

(Restriction) Ratio (K)

Know that:

And for stability:

Restriction ratiodefined as:

So that (since r = a pcn):

K thus exerts very strong influence on equilibrium pc.

For n = 0.75 (say),pcK4, so very sensitive - hence

preferable to have low values of n for reduced sensitivity.

1

* 1 nc pp c a K

*

b c

t p

A pK

A c r

o b p bm m r A

*c t

o

p Am

c

(15)

(16)


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Effect of Burn Area (Ab)

on Burn Rate (r)

Since r = a pcn

and pc= (c* a pK)1/(1n)

Increase in K or Abfor

a fixed Atvalue modifies

the burn rate curve as

shown to shift the

operating point upwardsand increase thrust.

m


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Ballistic Additives Some substances (e.g. lead salicylate, lead stearate, etc.)

may be added to a solid propellant to reduce n and thus reduce

sensitivity and improve combustion stability - known as

platonisation.

Gives relatively flat curve of r versus pcover pcrange.

Propellant Initial


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Propellant Initial

Temperature Effect Affects burn rate coefficient (a) and thus burn rate (r).

Variations of up to 35% possible for pcand tb.

Total impulsehardly affected but thrustis and can give

problems.


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Temperature Sensitivity

Sensitivity expressed in form of temperaturecoefficients:

Burn rate:

Typically 0.2% per oC

Pressure:

Typically 0.15 to 0.35% per oC

ln 1

c c

p

p pp p

d r dr

dT r dT

ln 1c c

K

p c pK K

d p dp

dT p dT

(17b)

(17a)

T


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Temperature

Sensitivity (Cont.) Effect of temperature of burn rate and pressure

obtained from:

It may also be shown that:

(18b)

(18a)o pr r T

o Kp p T

1

1K p

n

B R t


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Burn Rate

Gas Velocity Effect Most solid

rocket motors

are in form of

perforated,cylindrical stick.

Gas produced by burning charge flows past

burning surface and out through nozzle.

Very high velocities producedaffects heat

transfer rate and increases burn rateerosive

burning.

B R t


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Burn Rate

Combustion Instabilities

Burning not necessarily smooth and regularprocess.

May produce sudden unpredictable pressure

peaks and result in burst cases, propellantlosses, reduced range and loss of accuracy.

Known as resonant burning.

Oppositelow pressure troughscan cause

intermittent stopping of burningchuffing. Particularly a problem with low initial

temperatures and over-sized nozzle throats.


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Rocket Propellants

Require a suitable mix of fueland oxidiser.

Four main possibilities:

Petroleum + oxidiser

Cryogenic

Hypergolic

Solid

Solid propellants generally favoured formilitary applications.


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Rocket Propellants

Petroleum Uses refined kerosene known as

RP-1(rocket propellant 1), burnt

with liquid oxygen (LOX) or, on

older rockets, with nitric acid asoxidiser.

Used, for example, on first stage

boosters of Delta, Atlas-Centaurand Saturn rockets with typical

Ispof 2600 m/s. Atlas-Centaur


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Rocket PropellantsCryogenic

Generally uses liquid hydrogen(LH2) as fuel

with liquid oxygen(LOX) as oxidiser.

Requires temperatures of -183oC for LOX and

-253oC for LH2, giving formidable engineeringproblems.

In liquid state, density is vastly increased so

that much smaller tanks are needed. Major storage problems so mostly unsuitable

for military rockets.


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Rocket Propellants

Cryogenic (Continued) Used on J-2 engines

on Saturn V 2nd/3rd

stages with Isp

of

about 4250 m/s and

also on Space Shuttle

(Isp= 4550 m/s).

LH2and LOX burnclean so by-product is

water vapour.

Saturn V

J-2 engine

Rocket Propellants
http://www.boeing.com/defense-space/space/rdyne/sightsns/images/j2grey.gifhttp://www.boeing.com/defense-space/space/rdyne/sightsns/images/j2grey.gifhttp://antwrp.gsfc.nasa.gov/apod/image/saturn5_apollo11.gif


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Rocket Propellants

Hypergolic

Fuels and oxidisers which ignite on contact giving

easy start/restart capabilities often needed for

spacecraft systems.

Much easier to store than cryogenics.

Fuel usually monomethyl hydrazine(MMH) with

oxidiser nitrogen tetroxide(N2O4) - both highly

toxic. Isptypically 3100 m/s.

Used on second stage of Delta, Titan and also on

Space Shuttle for orbital manoeuvres.


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Solid Propellant Selection

Desirable properties May be divided into those concerned with:

performance

satisfactory operation storage & handling

supply

Many are mutually conflictive in nature.


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Propellant Selection -

Performance Considerations By considering specific impulse (eq. 10), require:

Molecular weight of combustion products as low as

possible.

Temperature of combustion as high as possible.

Average propellant density as high as possible.

Specific heat ratio of combustion products as low aspossible.

Calorific value per unit mass as high as possible.

P ll t S l ti


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Operation Considerations Combustion temperature not too high otherwise mechanical

difficulties.

Chemically inert - oxidisers affect pumps, valves, seals, etc.

Similar expansion coefficient for propellant and case.

Good mechanical properties to prevent distortion from high

acceleration loads.

High thermal conductivity to minimise temperaturegradients.

High specific heat if used for cooling.

P ll t S l ti


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Storage/Supply ConsiderationsStorage

Low vapour pressure for liquid propellants.

Non-toxic propellants & products of combustion.

Low explosion & fire hazards - no detonation risk.

Supply

Readily available in peace and war time.

Low cost, though only small part of total R & D costs.

Solid Propellant


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Solid Propellant

Classifications

Double Base Propellants Homogeneous mixture of two explosives - usually

nitroglycerine (NG) dissolved in nitrocellulose (NC),

sometimes with additives.

Advantagesare:

Smokeless; low cost; low n value (about 0.3) and can beeasily platonised for good burning stability.

Disadvantagesare: Lower density than composites so low specific impulse;hazardous to manufacture; grain requires structural support.

Solid Propellant


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Solid Propellant

Classifications

Composite Modified Double Base

(CMDB) Propellants

Double-base propellants which include

addition of compounds, such as:

Ammonium perchlorate (AP)

Aluminium fuelSolid explosive nitramine compounds

(HMX, RDX).

Typical CMDB Ingredients


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Typical CMDB Ingredients

Solid Propellant


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Solid Propellant

Classifications

Composite Propellants

Heterogeneous mixture of powdered metal, crystalline

oxidiser and polymer binder.

Most common type used.Advantagesare:

Higher density & specific impulse than DB; easier tohandle, store & manufacture; more reliable combustion.

Disadvantagesare:

Smoky (depending on aluminium content); toxic exhaustgases.


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Solid Propellant Properties

n

Solid Propellant


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Solid Propellant

Applications Predominant in GW applications, mainly due to

ease of utilisation, straightforward handling, lack of

servicing equipment and simple firing.

Thrusts vary from 5 N to 10 MN.

No moving parts, unless TVC included.

Rarely able to turn thrust on/off and modify thrust

on demandthoughsolid propellant pulse rocket

motor may change this in the future.

Solid Propellant


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Solid Propellant

Applications - Profiles

Different profiles include:

Boost thrustfor anti-tank, ramjet boost, etc.

Boost-sustainanti-missile, anti-aircraft, etc.

Boost-coastair-to-air.

UK Missile Solid


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UK Missile Solid

Propellant Applications

Include:

Sea Cat (boost & sustain)

Sea Dart (boost)

Sea Wolf (boost)

VL Sea Wolf (TVC launch &

boost)

Starstreak (eject & boost)

Sea Wolf

Sea Dart

UK Missile Solid


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UK Missile Solid

Propellant Applications (Cont.)

Include:

Swingfire (dual boost & TVC

sustain)

Sea Skua (boost & sustain)

Rapier (dual boost & sustain)

Javelin (eject & boost)

Skyflash (boost)

Sea Skua

Javelin

Rocket Motor
http://www.army-technology.com/projects/javelin/javelin12.html


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Rocket Motor

Applications (Cont.)

Solid Propellant Rocket


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Solid Propellant Rocket

Motor Design

Cylindrical body good shape for pressure vessel - also easy

to manufacture, store & transport and good aerodynamically.

In this example, charge bonded to insulation layer which is

bonded to case.

Grain Design Charge


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Grain Design - Charge

Geometries

(a) cigarette (axial/end) burner - long burn time, big CGchange;

(b) slotted-tube radial burner; (c) star centre radial burner.

Different shapes

used to vary burn

rate/time and thus

thrust.


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Grain Design - Thrust

ProfilesThe grain design can be used to give thrust variation. Progressive

Burn during which thrust, pressure and burning area

increases.

Neutral


remain constant (15%). Regressive


decreases.


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Grain Burn

Characteristics

Grain Design (Cont )


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Grain Design (Cont.)

More possible grain geometries

Propellant Grain


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Propellant Grain

Definitions

Web thickness (b)

Minimum thickness from burning surface to case

wallfor an end-burner, equal to length of

grain.

Web fraction (bf)

ratio of web thickness (b) to grain radius

Volumetric loading

ratio of propellant volume (Vb) to chamber

volume (Vc)

C P ll t


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Common Propellant

Grain Configurations

C D i


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Case Design

Metal Cases Typical metals used are high strength steels and

titanium alloys.

Many advantages: Toughness - preventing damage during handling.

High melting temperature - allowing less insulation.

Good aging properties with time and resistance to

weather exposure.

Thin walls possible so more propellant may be packed in.

Case Design


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Case DesignFibre Reinforced Plastic Cases

Glass or carbon fibre laid into patterns and bonded

with epoxy resin.

Advantages:

High strength/weight ratio (up to 10 x metals); lay-

up may be tailored to suit stress requirements.

Disadvantages:

Lower melting temperatures (approx 180oC);reinforcements needed in mounting areas; high

thermal expansion coefficient; low thermal

conductivity.

P ll t G i M ti


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Propellant Grain Mounting

Secure mounting needed due to high loadsexperienced during manoeuvres.

Case bondedmethod

used for large grains

(diameter > 0.5 m, mass

> 300 kg).

Free standingmethod

used for smaller grain

types and when grain

has good stiffness

properties (DB types).

Case Bonding


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Case Bonding Major problem due to different expansion

coefficients of case and propellant.

Particular problem

in air-to-air

systems with big

temperature

fluctuations.

S lid R k t N l


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Solid Rocket NozzlesNozzle must provide thrust along rocket axis and

maximise it for given pressure ratio.

Three major

configurations

used.

Solid Rocket Nozzle Design


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Possibilities

S lid R k t N l D i


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Bell shape gives

maximum

performance and

lowest losses butalso the biggest

so only tends to

be used for

space

applications.

Nozzle Heat Transfer


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Nozzle Heat Transfer

Rates

Very high heat

transfer rates in

nozzle due to

combination of highgas temperatures

and high velocities

peak at throat.

Metals unsuitable in

such areas.


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Typical Nozzle Materials


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Typical Nozzle Temperatures

& Ablative Losses

Rocket Nozzle Cooling


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Rocket Nozzle Cooling If ablation rates are excessive, cooling may be

achieved by burning lower temperature propellant

on side of chamber - but reduces specific impulse.

Th t V t C t l (TVC)


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Thrust Vector Control (TVC)

Sometimes a requirement to change flight directionwithout using aerodynamic control methods - TVC

then used.

TVC methods may be placed into four main

categories:

Mechanical deflection of nozzle.

Insertion of heat resistant bodies into the main

jet.

Injection of fluid into the side of the diverging

nozzle section.

Separate thrust producing devices.



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Common Approaches to TVC

L = Liquid

S = Solid



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More Common Approaches to TVC

L = Liquid

S = Solid

Li id P ll t R k t


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Liquid Propellant Rockets Fuel & oxidant are stored outside the

combustion chamber with two possible

supply methods:

pressurised storage tanks (pressure-feed

system). use of pumps (turbopumpsystem)

complex.

Turbopump system only needs tanks to

sustain ambient pressure values -mainly space applications.

Lance

Pressure-feed systems have upper size/weight

limitation - mainly GWapplications.

Typical Liquid Propellant


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yp ca qu d ope a t

Turbopump Rocket

S lid Li id R k t


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Solid v Liquid Rockets

Solid Rocket Advantages High propellant density (volume-limited designs).

Long-lasting chemical stability.

Readily available, tried and trusted, well proven in

service.

No field servicing equipment & straightforward

handling.

Cheap, reliable, easy firing and simple electricalcircuits.

S lid Li id R k t


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Solid v Liquid Rockets

Solid Rocket Disadvantages

Lower specific impulses.

Difficult to vary thrust on demand.

Smoky exhausts.

Performance affected by ambient

temperature.

Advantages less distinct compared with

modern packaged liquid propellant rocket

(e.g. Lance).

Packaged Liquid-


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g q

Propellant Rocket Engine Some tactical missiles require total impulse

of < 500 kNs - range favourable to

pressure feed liquid-propellant system.

If hypergolic propellants used then startingand ignition systems can be simplified.

Concept further simplified by pre-packing

and sealing propellants into respectivetanks well before use.

Packaged Liquid Propellant v


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Solid Propellant

Packaged liquid propellant system advantages

include:

control of thrust on command, typically over 5:1

range. wide operating temperature limits.

no limitation on temperature cyclingno grain to

crack.

low smoke and flash emission levels.

reduced radio interference fromexhaust.

long term storage.

modular design.

Packaged Liquid Propellant v


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Solid Propellant (Cont.)

Packaged liquid propellant system disadvantages

include:

relatively new technology.

reduced reliability due to greater number of parts.

fire hazard if both tanks are ruptured.

toxic fumes.

more fragile and liable to handling damage.

MGM-52C Lance Packaged


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Liquid Propellant System

Pre-packaged bi-propellant liquid-rocket system using

unsymmetrical dimethylhydrazine (UDMH) as fuel and

red fuming nitric acid (IRNFA) as oxidizer.

Rocket Design Example


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A sea level rocket requires 10 kN of thrust for 20 sand is restricted in size to a maximum length of 1 m.

Size the nozzle throat, nozzle exit, propellant charge

and suggest a suitable charge shape given the burn

duration and required thrust.

Assume sea level pressure pa= 1 bar, propellant is

XLDB/AP where a (constant) = 3 x 10-6, = 1.25, R

= 0.325 kJ/kgK.



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Solution For XLDB/AP, using Table 1

p = 0.067 lb/in3= 1851 kg/m3, Isp= 269 s = 2639 m/s

To= 6060o

F = 3621 K, r = 0.35 in/s = 8.88 mm/s, n = 0.5

If fully-expanded,

Using the X-flow function for a choked throat and

M=1:

e spF mu mc mI

10000 / 2639 3.789 kg/sm

( 1) /2( 1)/( ) /(( 1) / 2) 0.658o c tm RT p A



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Solution (Cont.) Using r = a pc

n,

8.88 x 10-3= 3 x 10-6pc0.5and pc= 87.6 bar

From X-flow function, At = 3.789 (325 x 3621)/ (0.658 x 87.6 x 105)

= 7.131 x 10-4m2dt= 30.1 mm

pc/pa= 87.6 and, from charts, Ae/At= 10,

Ae= 7.131 x 10-3m2and de= 95.2 mm

/(0.658 )t o cA m RT p



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Solution (Cont.) For long range missile, assume use of end burner

propellant and = 3.789 x 20 = 75.78 kg

Also,

Ap= 3.789 / (1851 x 8.88 x 10-3) = 0.231 m2

dp= 0.54 m

and mp= Aplpp, lp= 75.78 / (0.231 x 1851) = 0.178m

Unrealistic grain dimensions so redo with another type.

p bm mt

p pm A r



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Solution (Cont.) Use slotted radial burner instead with inside and

outside diameters of dpiand dpo

Assume lp= 800 mm, leaving 200 mm for the nozzle.

Using mp= Aplpp, Ap= 75.78 / (0.8 x 1851) = 0.051

m2

0.051 = (/4)(dp22

dp12

) (dp22

dp12

) = 0.04 m

2

For dp1= 0.05 m, dp2= 206 mm(reasonable)

Though this will give a progressive thrust profile

Missiles Propulsion 1

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