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The information contained in this document is GKN Aerospace Sweden AB Proprietary information and it shall not either in its original or in any modifiedform, in whole or in part be reproduced, disclosed to a third party, or used for any purpose other than that for which it is supplied, without the written
consent of GKN Aerospace Sweden AB. Any infringement of these conditions will be liable to legal action.
1
Chemical Rocket Thrust Chambers23 Jan. 2013
Aerothermodynamics, Jan stlund
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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GKN was active when steel superseded iron during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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GKN Aerospace Engine Systems in Sweden
11
Military aircraft
engines
Sub systems for
rocket enginesEngine ServicesComponents 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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GKN Aerospace Engine Systems in Sweden
12
Military aircraft
engines
Sub systems for
rocket enginesEngine ServicesComponents 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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GKN Aerospace Engine Systems in Sweden
13
Military aircraft
engines
Sub systems for
rocket enginesEngine ServicesComponents 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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GKN Aerospace Engine Systems in Sweden
14
Military aircraft
engines
Sub systems for
rocket enginesEngine ServicesComponents 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
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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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21
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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22
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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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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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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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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30
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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34
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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CHEMICAL ROCKET THRUST CHAMBERS
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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40
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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41
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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47
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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48
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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49
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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52
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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56
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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57
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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CHEMICAL ROCKET THRUST CHAMBERS
INTERNAL AND EXTERNAL LOADS
58
Main jet pressureload may cause
buckling
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59
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=rowVdcnwJr87/28/2019 Chemical Rocket Thrust Chambers KTH (2)
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60
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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62
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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63
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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64
Radiatively cooled nozzles
No cooling flow
Very high material temperatures
1400K-2000KSimple structures
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65
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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66
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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67
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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68
Heat load
Vulcain 2 NE
Spiral tube wall:convective cooling
Sheet metal skirt wall: film-and radiation- cooling
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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70
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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71
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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72
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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73
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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74
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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75
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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76
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
Film coefficent flame side
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77
Film coefficent flame side
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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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
Conjugated Heat transfer analysis
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Conjugated Heat transfer analysis
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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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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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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