Chemical Rocket Thrust Chambers KTH (2)

7/28/2019 Chemical Rocket Thrust Chambers KTH (2)

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Chemical Rocket Thrust Chambers23 Jan. 2013

Aerothermodynamics, Jan stlund


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GKN Aerospace Sweden AB Proprietary Information. This information is subject to restrictions on first page.

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Outline of presentation

Overview of GKN

Space Propulsion at GKN- Ongoing Rocket Nozzle Projects

Chemical Rocket Thrust Chambers - Basic Concepts and Theory

Chemical Rocket Thrust Chambers -Nozzle Contour Design

Chemical Rocket Thrust Chambers -Internal and External Loads

Heat Transfer Methods used in Concept Design Phases

Heat transfer Methods used in Detail design and Verification

Phases

Manufacturing of Vulcain 1&2 NE and Vulcain 2+ NE DEMO


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OVERVIEW OF GKN

3


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GKN plc. at a glance

GKN is a global engineering groupEstablished in 1759, it has over 250 years of engineeringexperience

Its technologies and products are at the heart of vehicles and

aircraft produced by the worlds leading manufacturersGKN operates four divisions:

GKN Driveline (46%)

GKN Powder Metallurgy (14%)

GKN Aerospace (24%* excl Thn)

GKN Land Systems (14%)

4


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Some GKN history

GKN started as a tiny iron work on the Welsh

hillsides in 1759.

GKN was active when steel superseded iron during

the railway boom in the 1860s.

After the First World War, GKN moves in to the 20thcenturys greatest new industry automotive.

In 1988 Guest, Keen & Nettlefolds changed name

into GKN plc. and to took off into aerospace industy.

In October 2012 GKN bought Volvo Aero Corporation

in Trollhttan and renamed it GKN Aerospace Engine

Systems.

5


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GKN is a truly global company

GKN: ~44 000 people in more than 35 countries

GKN Aerospace: ~11000 people

GKN Aerospace Engine Systems in Trollhttan: ~2400 pepole

6


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GKN Aerospace sales and marketshare

7

GKN Aerospace sales:


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GKN Aerospace Airframe and niche products

8


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GKN Aerospace Engine products

9


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GKN Aerospace Engine Systems in Sweden

10

Military aircraft

engines

Sub systems for

rocket enginesEngine ServicesComponents for

aircraft engines

and gas turbines


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11

Military aircraft

engines

Sub systems for


aircraft engines

and gas turbines

90 % of all new large commercial aircraft engines useour components

Engine components, Engine technology

Engine technical support, Engine MRO* services*) Maintainence, Overhaul & Repair


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12

Military aircraft

engines

Sub systems for


aircraft engines

and gas turbines

RM12 in the Swedish Gripen fighter aircraft main contractor

We develop and produce components for several other

military engines, such as the F404 and F414 for the F18, and

F110 for the F16, and F135 for the Joint Strike Fighter


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13

Military aircraft

engines

Sub systems for


aircraft engines

and gas turbines

European Center of Excellence for nozzles and turbines

Patented sandwich technology

Turbines with extreme performance


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Military aircraft

engines

Sub systems for


aircraft engines

and gas turbines

Commercial engine overhaul experience

since 1966

On-site services and around-the-clock technical support

Lease/exchange engine support


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Center of Excellence for Aerothermodynamics

CoE forAerothermodynamics

Aeroacoustics

Aeromechanic

Low-observability

Heat transfer

Combustion

Aerodynamics

Work force

~25 persons

1 professor

3 companyspecialists

6 engineeringmethod specialists

13 PhD

1 Lic.Eng

12 MSc

15


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SPACE PROPULSION

ROCKET NOZZLE PROJECTS

16


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Projects we work in; Vulcain 2

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Projects we work in; SWEA/SWAN


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Projects we work in;

SCENE/Score-D within FLPP


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Cooperation withinnetworks, partners andcustomers in projects


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Trollhttan,

GAS

Vernon,

Snecma

Ottobrunn,

Astrium

Lampoldshausen,

DLR

Noordwijk,

ESA

Korou,

A5 launch siteParis,

CNES

Stockholm,

Swedish National Space Board

West Palm Beach,

Pratt&Whitney

Rocket nozzle development involves cooperationwithin a network of industries, agencies and

institutions


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for rocket engine nozzles

Nozzles From Concept to Flight


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CHEMICAL ROCKET THRUST CHAMBERS

BASIC CONCEPTS AND THEORY

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Main Functions and Technological Challanges

for Thrust Chambers

Contour optimization for minimum expansion losses andperformance prediction for expansion

Heat Transfer and Cooling

Mechanical design against internal (pressure/temperature)

and external loads (buffeting / booster radiation)Compromise design for overexpansion (on ground) and

underexpansion (in vaccum) for first stage application


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25

ALL propulsion systems work by exchanging momentum

sailboats exchange momentum with the air

automobiles exchange momentum with the Earth

jet engines take low-speed air and eject it at a higher speed

rocket engines exchange momentum with the propellant

momentum added to rocket = mass of ejected propellant x velocity

DP=mpropve

Rocket Basics : Momentum Exchange

Change inmomentum Propellant mass (Kg)

Exit velocity (m/s)

REACTIONACTION REACTIONACTION


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26

Thrust Chamber Fundamentals

A thrust chamber generates thrust

(F) by converting thermal energyof hot combustion gases(temperature) to kinetic energy(velocity).

F-engine thrust

Tc-combustion chamber temperature

pc-combustion chamber pressure

m-propellant mass flow

pe -nozzle exit pressure

ve -nozzle exit velocity

Ae -nozzle exit area

pa -ambient pressure

At -nozzle throat area

e= Ae/At -nozzle area ratio alt. Expansion ratio

Nozzle

OxidizerFuel

CombustionChamber

Convergent part

Divergent partThroat

Pc,Tc

m&

pa

pe ve, Ae

eaeee ApApvmF )( &


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27

Energy balance

RQhh 0102

cccc hvvhh }0{2

1 202

ehh 002

c

ecpeceTTTchhv 12)(2

/)1(

1)1(

2

c

ece

p

p

M

TRv

Assume1D isentropic expansion of acalorically perfect gas (Cp&R const.)

(QR - added heat fromchemical reaction)

Low Molecular weigth and high chambertemperature gives high exhaust velocity!

QR

Propellant

in

1

2

e

Propellant

out

h

s

1

2

e

pc

pe

Expansion

QR

For well expanded nozzle with high area ratiope0

ce TM

Rv

)1(

2max,

MRR

Rc

Tch

p

p

)1(


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28

Thrust chamber performance parameters

Isp =F/m=c*CF is the Specific Impulse [m/s]1

c* is the Characteristic Velocity [m/s]primarily a function of the combustion chamber properties

CF is the Thrust Coefficient (dimensionless)can be considered to be a function of nozzle geometry only

FFtcspeaeee CcmCApImApApvmF*)( &&&

]/[81.9];[ 200 smgsgmFIsp &1 To make it independent of the unit system following definition is also used


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Thrust chamber performance parameters

Isp =F/m=c*CF is the Specific Impulse [m/s]1

is a measure of how well a given flow rate of propellant is turnedinto thrust

It is an important performance parameter for the launcher

Recall the rocket equation: Du=Isp ln(m0/ mB)Du= launcher need, m0= initial mass, mB=dry mass

Consequences for engine design & layout :Maximize thrust chamber Isp and Minimize thrust chamber dry mass

FFtcspeaeee CcmCApImApApvmF*)( &&&

]/[81.9];[ 200 smgsgmFIsp &

1 To make it independent of the unit system following definition is also used


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Characteristic velocity

The Characteristic Velocity relates the combustion chamber pressure to thepropellants burned, and thus reflects the propellant energy and the

combustion efficiency

It is essential independent of the nozzle characteristics and may be used tocompare the characteristics of different propulsion systems and propellants

The Characteristic Velocity defines together with the chamber pressurethe size of the thrust chamber

&

)(

)1(2

)1(

1

2

c

tct

c

tc

c

tc

RT

ApAT

TRTm

)1/(

00

)1/(1

00

20

2

11

T

T

p

p

T

T

MT

T

Compute at sonic throat:uAm & Recall the isentropic relations

)()(

*

MRTRT

m

Apc

cctc

&


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Thrust coefficientThe Thrust Coefficient relates the created thrust to the stagnation pressureand thus reflects the expansion properties of the exhaust gas

The Thrust Coefficient gives the amplification of the thrust due to the gasexpansion in the supersonic nozzle compared to the thrust delivered if thechamber pressure only acted over the throat area

t

e

c

ae

c

e

tc

FA

A

p

pp

p

p

Ap

FC

ee

;11

2

1

21

1

1

t

e

c

aee

t

e

c

aee

tctc

FA

A

p

pp

c

v

A

A

p

ppvAp

m

Ap

FC

*

&

MRT

m

Apc

ctc

&

*

/1

11

2

c

ece

p

p

M

TRv


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Typical flow properties in a H2/O2 thrust

chamber @ vaccum conditions ( 1.2)Fcc=mvt+pt At FNE=m(ve-vt)+peAe-pt At

Stagnation CC Throat Nozzle exit

Area ratio 2.5 1 45

Pressure 110 bar 98 bar 56 bar 0.18 bar

Temperature 3550 K 3519 K 3218 K 1234 K

Velocity 0 m/s 395 m/s 1542 m/s 4128 m/s

Mach number 0 0.24 1 4.32c* 2277 m/s

Mass flow rate 237.5 kg/s

Thrust 671.5 kN 1024 kN

Isp 2827.4 m/s 4312 m/s

CF 1.24 1.89


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Operation of Nozzle at Off-Design Conditions

Optimum performance is obtained if the nozzle exit pressure is equalto the atmospheric pressure

Exit PressureBelow

AmbientPressure

Exit PressureEquals

AmbientPressure

Exit PressureAbove

AmbientPressure

Bell Nozzle at Sea Level:

The exhaust plume ispinched by high ambient

air pressure, reducing itsefficiency.

Bell Nozzle at Optimum Altitude:

The exhaust plume is column-shaped producing maximum

efficiency.

Bell nozzle at High Altitude:

The exhaust plume continuesto expand past the nozzle exit

reducing efficiency.

PlumeBoundary

PlumeBoundary

Shocks

Flowseparation

region


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Specific impuls of an Ideal thrust chamber

For an ideal thrust chamber with fixed chamber conditions theperformance is determined by the nozzle area ratio

FC

c

ae

c

e

c

c

spp

pp

p

pTMRI

e

1

1

1

*

11

2

1

2

1

1

2

2

11

1

21

e e

eee

tt

t

e MMv

v

A

A

12

2

11 ece Mpp


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Specific impuls of an Ideal thrust chamber

Tradeoffs in selecting the area ratio!

N l t f St A li ti


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Nozzle concepts forFirstStage Application

Key requirements:

Stable operation on ground

High performance (low losses)

at high altitude

Classical approach:

Bell nozzle

(e.g. Vulcain, SSME, )

Advanced nozzles withaltitude adaptation e.g.:

Dual bell nozzle

Extendiable nozzle

Aerospike nozzle

Plug nozzle


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NOZZLE CONTOUR DESIGN

37


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Performance losses in a nozzleThree main categories of loss mechanisms:

Geometric or divergence loss.Viscous drag loss.

Chemical kinetic loss.


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Losses

The optimum nozzle contour is a design compromise

that result in a maximum overall nozzle efficiency.

A long nozzle is needed to maximise the geometric

efficiency; But at the same time, nozzle drag and drymass is reduced if the nozzle is shortened.

If chemical kinetics is an issue, then acceleration of

exhaust gases at the nozzle throat should be slowed by

increasing the radius of curvature applied to the designof the throat region, at the cost of an increased nozzle

length and dry mass.


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Classical Nozzle Contour Types15o

cone TIC TOC TOP

Mach number distribution in different contours. The thickline indicates the approximate position of the internal

shock.

TIC - Truncated Ideal Contour

TOC - Thrust Optimised Contour

TOP - Thrust Optimised Parabola


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Conical Nozzle

evmF &

Assume conical flow at the exit of a nozzle and optimum expansion (pe= pa)

rv

m

dAvvv e

Axr

e&

cos

sin2

rx vv

drrdA

cos1

sin

sin2

cossin2cos 221

0

2

0

22

rr

Ar

Ar

e v

dr

dr

vdAv

dAvv

e

e

2

cos1

cos1

sin2

21

r

e

v

v

Reduction in thrust compared to an ideal nozzle with all the flow in axial direction

tan2

1 te

t

AA

D

L

For ideal conical flow, , vr

are constant overAe


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Nozzle Design with the use of MOC

In supersonic flow the Euler equations are hyperbolic, i.e. well

suited for the use of Method Of Characteristics (MOC).The most common method for generating rocket nozzle

contours.

In mathematics, the method of characteristics (MOC) is atechnique for solving partial differential equations.

For a first-order PDE, the method of characteristics discovers

lines (called characteristic lines or characteristics) along whichthe PDE becomes an ordinary differential equation (ODE). Once

the ODE is found, it can be solved along the characteristic lines

and transformed into a solution for the original PDE


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4317 04 2012,

Characteristic lines

The solution of the flow equation is found by constructing thecharacteristic curves/net

In a supersonic nozzle the characteristic lines is identical to the Machlines (small pressure perturbations in the flow are transported along the

Mach lines)

(-)-characteristic

(+

)-cha

racte

ristic

W

2

1

3

x

y

(-)-characteristic

(+

)-cha

racte

ristic

W

2

1

3

x

y With the conditions at point 1 and 2 thelocation and flow conditions of a new point 3can be numerical determined

Schlieren photograph of flow in a 2Dconical nozzle (after Busemann)

The Mach lines are made visible bysmall roughness on the nozzle walls


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Nozzle Design, initial expansion region

The basis in all MOC nozzle design methods is the Kernel.


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Ideal Nozzle DesignAn ideal nozzle produces an isentropic flow (i.e. without internal shocks)and gives a parallel and uniform exit flow.

With the condition that the last LRC KE is a uniform exit charactereisticMOC can be used to construct the inviscid nozzle contour.

After the inviscid design a boundary layer correction is added tocompensate for the viscous effects.


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Characteristics net in an ideal nozzle

MDesign

=4.6, =1.2.


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Truncated Ideal Contoured Nozzle

Ideal nozzle extremely long not suitable for rocket

applications.

Length necesarry to produce 1-D flow at exit.

Thrust contribution by the last part of nozzle is

negligible.

A pratically more feasable rocket nozzle is obtained

by truncating the nozzle.Such nozzles are called Truncated Ideal Contoured

(TIC) nozzles.


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Truncated Ideal NozzlesThe method can be outlined as follows:

A complete set of ideal nozzle contours is synthesised in a plot

together with lines representing constant surface area, exitdiameter, length and vacuum thrust coefficient respectively.

Within a given constraintsuch as expansion ratio

(or exit diameter), surfacearea, or length anoptimisation process canthen be used todetermine where totruncate the full nozzlecontour to obtain

maximum performance.

Length

Radius


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DA

C

B

Length

Radius

Truncated Ideal Nozzles

Max performance for a given:

Expansion ratio point ASurface area point BLength point CPoint D, the most thrustobtainable from any given

nozzle contour. Not ofpractical interest


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50

Ariane 4, Viking4

Viking4 (TIC)Propellants: N2O4/UDMH

Thrust(vac): 82 tons.Isp: 2960 m/s

Burn time: 125 sec.Mass Engine: 850 kg.Diameter: 2.6 m.Length: 3.5 m.

Chamber Pressure: 58.5 bar.Area Ratio: 30.80.

Oxidizer to Fuel Ratio: 1.70.


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Energia Buran, RD-0120

RD-0120 (TIC)Propellants: LOX/LH2Thrust(vac): 200 tons.

Isp: 4550 m/s.

Burn time: 600 sec.Mass Engine: 3,450 kg.Diameter: 2.4 m.Length: 4.5 m.

Chamber Pressure: 218 bar.Area Ratio: 85.7.

Oxidizer to Fuel Ratio: 6.00


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T

N

O K

Kernel

TN

Nrtd

O

rt

P

E

Control

surface

E

rE

L

C

Thrust Optimised Contoured Nozzles (TOC)

With the use of calculus of variation the conditions and shape of the controlsurface can be found that gives maximum performance.

Contour found by back construction of MOC net, similar as for an ideal contour


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Characteristics net in a TOC Nozzle

P b li B ll N l


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54

Parabolic Bell Nozzles

Rao proposed a parabolic-geometry approximation to the TOC nozzledimensions from the inflection point to the nozzle exit.

These types of nozzles are often referred as Thrust Optimised Parabolic(TOP) nozzles.

With a parabolic approximation the contour is defined by the equation

Where the constants B, C, D and E are given by rtd, N, LE and E.

A very large number of contours can be generated. However, only a few

of these are an approximation of a real TOC contour.Selecting the proper inputs can approximate the TOC nozzle veryaccurately without introducing a significant performance loss.

0)( 2 EDrCxBxr


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Parabolic Bell Nozzles

One main difference between TOC and TOP nozzle flows:

An internal shock, i.e. crossing of the right running characteristic lines, isformed upstream the last LRC line in a TOP nozzle.

The wall pressure at the nozzle exit becomes slighly higher in a TOPcompared with the TOC nozzle.

Shown to be useful for 1-stage nozzles where a margin against flow

separation is important.

TOP nozzle TOC nozzle

Ariane 5 Vulcain


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Ariane 5, Vulcain

Vulcain (TOP)Propellants: LOX/LH2.Thrust(vac): 110 tons.

Isp: 4310 m/s.Burn time: 605 sec.Mass Engine: 1,300 kg.Diameter: 1.8 m.Length: 3.0 m.Chamber Pressure: 102 bar.

Area Ratio: 45.Oxidizer to Fuel Ratio: 6.20.


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Space Shuttle, SSME

SSME (TOP)Propellants: LOX/LH2.Thrust(vac): 232 tons.Isp: 4550 sec.Burn time: 480 sec.

Mass Engine: 3,177 kg.Diameter: 1.6 m.Length: 4.2 m.Chamber Pressure: 204 bar.

Area Ratio: 77.5.

Oxidizer to Fuel Ratio: 6.00.


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INTERNAL AND EXTERNAL LOADS

58

Main jet pressureload may cause

buckling


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Challenges innozzle design

The LoadsVibration loads

may start cracks

Balance

betweenheat load

and cooling

gves wall-

temperature

buckling

Boosters radiation

Buffeting - pressure pulsationsaround the nozzle during accent High nozzle cooling rate can

cause the water steam inthe flame to condensate

http://www.youtube.com/watch?v=rowVdcnwJr8

I t l l d Fl S ti
http://www.youtube.com/watch?v=rowVdcnwJr8http://www.youtube.com/watch?v=rowVdcnwJr8http://www.youtube.com/watch?v=rowVdcnwJr8http://www.youtube.com/watch?v=rowVdcnwJr8


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Internal load - Flow SeparationSeparation shock pattern in a TOP nozzle at two

different feeding pressures

Mach Disk

Shock wave

Reverse Flow

Supersonic jet

Mach No.

GSTP FSC REFERENCE NOZZLE

FREE SHOCK SEPARATION AT P0=14 Bar

0.00

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

0.10

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

X/L

Pwall/P0

Experimental Data

K-OMEGA Calculation

Vacuum Pressure Contour

GSTP FSC REFERENCE NOZZLE

RESTRICTED SHOCK SEPARATION AT P0=16 Bar

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

X/L

Pwall/P0

Experimental Data

K-OMEGA Calculation

Vacuum Pressure Contour

Subsonic

Core

Mach Disc

Separation

Reattachment

Supersonic jet

Mach No.

Internal load - Flow Separation


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61

Design for optimal performance requires models for accurate prediction of

separation point location

side-load levels

Both at design and off-design conditions (start up and shut down transients)

The chosen nozzle contour influences theside load and separation behavior

First TC test with Vulcain NE

Internal load Flow Separation


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Internal Load - Heat load

The typical heat transfer rates of rocket propulsion system are higher than thosein jet engines and the combustion temperatures are usually two times the

melting point of steel.

A correct estimate of the expected critical temperatures and temperaturegradients of the different propulsion components is highly important in any stage

of the design process.

Only sufficient material should be built into the walls to absorb or transfer theheat without risking excessive erosion or heating of the walls and without a lossof structural strength at the heated conditions.

The maximum temperature obtained at nominal/extreme operational conditionsdictates the choice of material and also the cooling method/layout that can beused.


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Type Basic Mechanism Application

Radiationcooling

Wall made of heat resistant metal; heat radiates tosurroundings

For un-cooled nozzleextensions andmonopropellant thrusters

Regenerativecooling

Fuel is circulated through hollow-wall cooling jacketprior to injection and absorbs heat by convection

Good for large- and medium-sized thrust chambers

Dump mass

cooling

Small amount of fuel is circulated through hollow-

wall cooling jacket and absorbs heat by convectionprior to ejection in to the surroundings

For large- and medium-sized

thrust chambers

Film cooling Liquid fuel or cool gas boundary layer is injectedalong wall surface

Usually with large units

Ablative cooling Progressive endothermic decomposition of fiber-reinforced organic surface material forming an

insulating porous char for passage of pyrolysisgases

Small thrust chambers andnozzle extension of large

units

Heat sinkshielding

Ablative heat sink wall pieces and graphite throatinserts resistant to high temperatures, erosion, and

oxidation

Solid-Propellant thrustchambers


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Radiatively cooled nozzles

No cooling flow

Very high material temperatures

1400K-2000KSimple structures


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Dump cooled nozzles

Small amount of coolant flow, 5-7 % of

total fuel mass flow

Coolant with large heat capacity

Low pressure < 50 bar

Complicated structure

High material temperatures 1100K

Vulcain


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Regeneratively cooled nozzles

Large coolant flows, 20-100% of the fuel

flowModerate material temperatures

500-800K

High pressures 200-400 bar

Complicated structures

Fil l d l


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Film cooled nozzle

Complicated functionSimple wall structure

Small flows 3-4% of total main jet flow

High material temperatures 1300-2000K

TEG flow dumped at low pressure

gives performance loss

Heat load


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Heat load

Vulcain 2 NE

Spiral tube wall:convective cooling

Sheet metal skirt wall: film-and radiationcooling

Combustion chamber wall:regenerative cooling

Combination of cooling concepts frequently used


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HEAT TRANSFER METHODS USED IN

CONCEPT DESIGN PHASES

69

One dimensional steady heat transfer analysis


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O e d e s o a steady eat t a s e a a ys s

Heat transfer from hot gas through solid wall to a cool fluid

Tg (3000 K)

Twg (700 K)

Twl (530 K)

Main jet (hot gas)

Tl (420 K)

(340 K)

Ambient air (294 K)

Inner wallOuter wall

Atmosphere

Coolant fluid

Radial distance from centerline of thrust chamber

Temperature

lwg

lwlll

wlwgww

wgwagg

qqqq

TThq

TTtqTThq

)(

))(()(

-Convective heat transfer at hot side-Heat conduction through solid

-Convective heat transfer at cold side

q -heat transfer rate per unit areahg -hot gas film coefficientTwa -adiabatic wall temperatureTwg -hot side wall temperature -thermal conductivity of solid

tw -wall thicknessTwl -cold side wall temperaturehl -coolant film coefficientTl -coolant temperature

Adiabatic wall temperature (recovery temperature)


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Adiabatic wall temperature less

than the stagnation temperatureHeat transfer from low speed fluid

close to the wall to higher speed

fluid farther from wall

Adiabatic wall temperature (recovery temperature)Adiabatic wall

Tg

Twa

T0g

pc

u

2

2

gg

gwa

TT

TTr

0

200

2

11

1)1(

gg

g

g

wa

M

rrr

T

Tr

T

T

rT

TM

T

TM

g

wag

g

wag

0

0

10

nr 1(Pr) Laminar flow n=2Turbulent flow n=3

9.0rFor typical rocket propellants

Recovery factor

One dimesional heat transfer analysis


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y

With the use of the previous equations the heattransfer rate per unit area and wall temperatures canbe calculated as

lwg

lwa

hth

TTq

11

llgwwgg

lwag

hAAAtAh

TTq

1

wg

wl

xwA

xwA

gg

ll

D

D

The channel side walls acts as cooling fins

This fin effect and other geometry effects can be includedby defining effective wall areas that are larger than theactual area:

wwhwl

gwawg

qtTT

hqTT

llwwgg AqAqAq

Estimate of 2D effects


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Estimate of 2D effects

0 0.25 0.5 0.75 1

Hotsidewalltemperature

Location on hot side, z/a

1D

2D

z

qconstant@DLwh

Lwhm

TT

TTf

1)(

DLwhLwhm TTfTT 1)(

Twhm maximum center plane hot wall temperature

Notice that Twhm < Twh

Film coefficent coolant side


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All the design bureaus and companies apply Nusselt-typecorrelations to describe the heat transfer

A rather simple example of such a correlation may read as:

or a little more complex

4.08.0 PrRe026.0hD

hll

DhNu

(Dittus-Boelter relation for turbulent pipe flow)

These coefficients (a-m) should describe the influence of:

cooling channel geometry

curvature effects

catalytic surface reactions thermodynamic (real gas, cryogenic conditions, vicinity to the critical point,

varying fluid properties)

fluid mechanic (turbulence, stratification)

chemistry (pyrolysis) etc.

Coefficients for Nusselt correlations


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Coefficients for Nusselt correlations

coolant side heat transfer hydrocarbons fuel

(From: Liang, K., Yang, B., Zhang, Z., Investigation of heat transfer and cokingcharacteristics of hydrocarbons fuels, Journal of Propulsion and Power, Vol. 14, No. 5, 1998)


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Film coefficent flame side

Usually the hot gas side heat transfer is described

in form of a Bartz-type correlation

2.08.0

2.0

8.18.0

*2.0

026.0

T

Tc

D

D

c

p

Dh

g

eptc

t

g

2

whg TTT

ref

ref

T

There are various modifications around, almost every company workswith their one correlation, which try to account for local effects such as

curvatures of liner and throat, area ratio, Mach number which all havean influence on the heat transfer



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Several important trends and observations can be made:

a) Smaller throat diameter leads to larger heat flux (~1/Dt0.2)

b) Heat flux is almost linear in chamber pressure (~pc0.8). This limits the feasibility

of high chamber pressure, which are other wise very desirable.

c) Maximum heat flux occurs at throat (~(Dt/D)1.8). One critical design

consideration is therefore the thermal integrity of the throat structure.

d) Lighter gases lead to higher heat flux, through the combined effects of cp and c*

(hg~1/M0.6).

e) The factor (Te/)0.8-0.2 (Te/)

0.68 is greater than unity. This

enhancement of heat flux follows mainly from the fact that the gas in the

boundary layer is mostly cooler than in the core, hence denser, and that the

turbulent heat conductivity is proportional to density

2.08.0

2.0

8.18.0

*2.0

026.0

T

Tc

D

D

c

p

Dh

g

eptc

t

g


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HEAT TRANSFER METHODS USED IN DETAIL

DESIGN AND VERIFICATION PHASES

78

Conjugated Heat transfer analysis


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GKN Aerospace Sweden AB Proprietary Information. This information is subject to restrictions on first page.79

Vulcain 2 film cooled nozzle

Coolant domain:Sub-sonic flowReal-gas flow

Solid domain

Flow direction

Core flow domain:Supersonic flow

Multi-specie flowChemical reactions

3-D heat transfer and flow analysis that includessimultaneous simulation of the chemical reactingmainstream in the nozzle, heat conduction in the solid

and the cold real gas flow in the cooling channelsMainly used in detail design and verification phase

Method time consuming and not feasible inconcept and preliminary design phases

III

Main jet Chemicalreacting flow

MetaltubeCoolant flow, real gas



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Comparison of 3D CFD (colored lines) and 1D-heat transfer prediction(black line) versus measurements (symbols) and metallurgic data

W

allTemperature

TCCOOL Wall Temp (TCCOOL S2)

3DCFD NE207 (S2) T (max)

3DCFD NE207 (S2) T (mean)

3DCFD NE207 (S2) T (min)

Max Wall Temp (Expertise NE207)

Tube 194 cold side

Tube 194 hot side

Tube 175 cold side

Tube 175 hot side

Tube 152 cold side

Tube 99 cold side

Line indicating start of grain growth

Line indicating carbide separation

Axial location

3D CFD T(min)

3D CFD T(max)

3D CFD T(mean)

1D prediction


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MANUFACTURING OF

VULCAIN 1&2 NE AND VULCAIN 2+ DEMO NE

81

Launch Vulcain 2


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V1 / V2 Common Processes


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V1 / V2 Common Processes

Raw material:

Inconell 600 tubes

V1: 456 tubes

V2: 288 tubes

Process:

Twisting and bending of tubes

Packing of tubes

V2 NE tube welding


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V2 NE tube welding

Welding of tubes(C-weld)

Close-up of welding process

V2 NE Skirt and Stiffeners


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V2 NE Skirt and Stiffeners

Built from 8 form-pressed

sections of sheet metal 25 stiffeners from sheet metal

are welded on the skirt wall

An expanding fixture is usedduring welding of stiffeners tominimize weld shrinkage

No further forming of the contouris performed

Skirt mounting


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Skirt mounting

A large welding positioner(BODE) is used

The Skirt is welded to the TEGManifold both on the outside and

the inside.

Vulcain 2 NE


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Vulcain 2 NE

Main data:

Fuel LH2

Oxidator LOX

Thrust 1359 kN

Combustion chamberpressure 120 bar

Combustion chambertemperature 3550 K

The turbine exhaust gasesare used to film cool the skirt

Length 2.168 m

Outlet diameter 2.094 m

Area ratio, outlet/throat 58

Weight 449 kg

Vulcain 2+ Demo


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GKN Aerospace Sweden AB Proprietary Information. This information is subject to restrictions on first page.8817 04 2012,

Outlet Manifold

(Haynes 230)

Hook stiffener for

radial stability(Inconel 625)

Stiffeners for

radial stability

(Inconel 625)

Metal deposition jacket

for axial stability(Inconel 625)

Stiffeners for axial

stability (Inconel

625)

Box stiffeners for

increased ovalizationfrequency (Inconel

625)

Inlet Manifold

(Inconel 625)

Cover band

for Cone

Joint (In625)

Thermal Barrier

Coating

SandwichWalls

(Haynes 230)

Vulcain 2+ Demo


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GKN Aerospace Sweden AB Proprietary Information. This information is subject to restrictions on first page.8917 04 2012,

The lower

cone is laserwelded atFORCE

Weldeduppercone

Chemical Rocket Thrust Chambers

Vulcain 2+ Demo milling of Liner


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For past demonstrators the channelmilling has been performed using twin

disk milling for optimum efficiency.

Vulcain 2+ DEMO in test bench


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Chemical Rocket Thrust Chambers KTH (2)

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