german aerospace center (dlr) institute of structures...
TRANSCRIPT
Knowledge for Tomorrow
CMC Rocket Thrust Chamber TechnologyStatus and Perspectives
M. Ortelt, H. Hald, D. [email protected]
German Aerospace Center (DLR)Institute of Structures and Design
AIRBUS DS – Space Systems - 6th R&T DAYS, Paris, 19.11.2015, Session 3, WG2 Technologies for Future Liquid Propulsion
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 1
Outline
- Conceptional aspects of the transpiration cooled CMC TCA
- Development status- Structural components- Materials- Test data
- Some future perspectives for CMC in space propulsion components
- Summary & outlook
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 2
CMC thrust chamber – Design concept
Transpiration cooled CMC thrust chamber – design principle
- Decoupling of single components – no bonding- Decoupling of mechanical and thermal loads- Specific hybrid interface technologies- Selective inner liner design
Features
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 3
- Standard CFD systems (FLUENT, CFX, …) are constructive (pure flow coupling)- Ongoing tool-development for ‚structure-flow-coupling‘ (TAU)- Investigations on materials out-flow homogeneity
Functional aspectsDLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 4
System analysis of transpiration cooling
0%
2%
4%
6%
8%
10%
12%
1 10 100 1000 10000 100000
Coolan
t ratio [‐]
Vacuum thrust [kN]
Tw = 800 K dc = 50 mm
dc = 100 mm
dc = 200 mm
dc = 440 mm
dc = 1000 mm
0%
2%
4%
6%
8%
10%
12%
1 10 100 1000 10000 100000
Coolan
t ratio [‐]
Vacuum thrust [kN]
Tw = 1200 K dc = 50 mm
dc = 100 mm
dc = 200 mm
dc = 440 mm
dc = 1000 mm
- Comparison of chamber size (scaling)- 50 mm chamber demonstration
- O/F = 5.5 (injector)- Contraction ratio 6.25- Characteristic chamber length l*=1.84 m- 7 % coolant ratio- Damage free operation- Amount of coolant depends on
- Hotgas conditions, As, T- D + p required coolant ratio
- Further coolant ratio reduction potential- Chamber length can be shortened
High operational efficiency predicted
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 5
Processes for Manufacturing of Nonoxide CMC
C, SiC fibers in C, SiC, SiC(N) matrix
Polymer Infiltrationand Pyrolysis
(PIP)
Liquid SiliconInfiltration
(LSI)
Combi-Process(PIP+LSI, CVI+LSI)
Chemical VaporInfiltration
(CVI)Focus at DLR Stuttgart
preform (e.g. fabrics, filament winding)
fiber-coating if necessary
infiltration (e.g. RTM)with C-precursor
pyrolysis (inert atmosphere) carbon matrix
finishing
siliconisation(inert gas, T>1420°C, Si+CSiC) stoichiometric SiC-matrix
preform (e.g. fabrics, filament winding)
fibrer coating
infiltration (e.g. RTM)with Si-precursor (e.g. polysilazane)
pyrolysis(inert atmosphere, T~1000°C,
e.g. polysilazane SiCN-matrix)
finishing
3-6 times todecreaseporosity
PIP
intermediate machiningjoining
Koch et al., DLR Werkstoffkoll. 2013LSI
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 6
Processing of Ceramic Matrix Composites (DLR-ST, BT)
• Autoclave30 bar, 350 °C
• Warm Press 350°C
• RTM 300°C• Pyrolysis, LSI, 2000°C• Machining Center
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 7
Thrust chamber – potential CMC derivativesInitial C/C model material LOX-sensitive!Other derivatives damage free after efficiently cooled and non-cooled operation:
Oxipol AvA-Z-ISC C/SiCN C/C (CVI)
10 35 18 7
2.3 2.6 1.6 1.6
Density kg/cm3
Open porosity [%] (porosity + permeability kd / kf adaptable by manufacturing process)
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 8
CMC thrust chamber – Components
Co-axial injector
Applied injectorsystems forcyl. 50 mm
Porous metal injector
Porous CMC injector
Elements of
oxide
CMC
for
the
LOX
injectionIntegrated ‚BlackEngine‘ demonstrator, cyl. 50 mm
C/C-SiC face-plate
Inner liner segment
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 9
CMC thrust chamber – mechanical interfacesCharacteristic hybrid interface types
Bolt interface for CFRP-metal joining Load-de-coupling double-shell nozzle extension
with keyed joint elements for CMC-metal joining
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 10
Thrust chamber - hot gas verification (LOX/LH2; LOX/GH2)
120 s, pc = 55 bar, LOX / LH2 operation, = 15 %
Structure testsP8
Efficiency testP6.12012
20 s, pc = 55 bar, LOX / GH2 (120 K), = 9 % Demonstration of the integrated CMC TCA
P6.1 firing testDec 2013
P8 (2005)
LOX / LH2, 65 70 bar52 s, 5 6 kg/s, τ = 4.2 % 90 bar tests
Vulcain contour
cyl. 50 mm
Contraction 6.25
cyl. 50 mm
Vulcain contour
cyl. 80 mm
Component tests
P8, 2008
cyl. 50 mm
C/C damages
near injector!C/C damage free
CMCs damage free
CMCs damage free
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 11
Thrust chamber – pressure loads during hot-runSegmented chamber module
Adequate pressure drops
at 8 % C/C porosity!
cyl. 50 mmInner liner:
Initial modelmaterial
C/C
O/F = 5.5
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 12
0
200
400
600
800
1000
1200
1400
1600
1800
0 10 20 30 40
Temperature [K]
Time [s]
U_T_1_8 [K]
U_T_2_8 [K]
U_T_3_8 [K]
U_T_4_8 [K]
U_T_1_9 [K]
U_T_2_9 [K]
U_T_3_9 [K]
U_T_4_9 [K]
Nominal hotrun-sequence
O/F = 5.5; = 6.72 %; pc 55 bar
Thrust chamber – thermal loads during hot-run
Cooling turned off
O/F = 2.0
Max. Tsurface 1800 K
750 /{
P6.1, 2012
cyl. 50 mm
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 13
CMC injector (‚Cone Injector‘)Features, goals, results
Mechanical design Flow designStart-up Steady stateChannel morphology
Demonstrator
LH2O/GN2
Spray test
Design
principle
0
0,05
0,1
0,15
0,2
0,25
0,3
‐6000 ‐1000 4000 9000 14000
M‐GH2‐COOL [kg/s]
M‐GH2‐INJ [kg/s]
M‐LOX‐INJ [kg/s]
DLR ‐ Cone Injector test campaign 'IZ2' ‐MASS FLOW CURVES
Time [s]
0
10
20
30
40
50
60
‐6000 ‐1000 4000 9000 14000
P_I_GH1 [bar]
P_I_GH2 [bar]
P_I_GH3 [bar]
P_I_LOX [bar]
U_P_IGN [bar]
Time [s]
DLR ‐ Cone injector test campaign 'IZ2' ‐ PRESSURE CURVES
Initially successful hotruns, P6.1, Dec 2013
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 14
Hyperboloid chamber contour – Orbital propulsion size
Comparison referred to typical 500 N class
Hyperboloid geometry
Numerical comparison
at similar performance
Hyperboloid chamber design
Perfectly combined with ‚cone injector‘ technology
- Advantages for- film cooling- transpiration cooling
- Composite affine structure manufacturing (winding technique)
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 15
Hyperboloid chamber contour – Comparison VINCI size
Including insert
Without insert
Dynamic viscosity
Heat flux
Performance
Contour Mass flow Total heat flux Specific heat flux[kg/s] [MW] [MW/m2]
Classical 43 16 55
Hyperboloid without insert 42 16 56
Hyperboloid including insert 43 27 52
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 16
Future potential - Preburner
Application principle
(oxide CMCs for ox-rich systems)
- Standard injector technologyconcerning …
- functional design- mixture ratio
- Propellant overhead injectedthrough chamber wall
- Long life and light weightstructures, similar to CMC thrust chamber design
Features
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 17
Further activities
- ALM / SLM technology for injector systems
- CMC components for hybrid propulsion systems
- Paraffin / N2O: Hybrid thruster
- CMC thrust chamber for an ADN orbital propulsion system
- Investigations on alternative propellants for CMC high performance thrust chamber application
- LOX / LCH4
- LOX / kerosene
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 18
Summary- The transpiration cooled CMC thrust chamber principle could be demonstrated successfully
- a simplified and de-coupled design principle has been proven
- no critical material degradation under efficient operation, considering scaling aspects
- mechanically safe structures
- First firing tests of the ceramic ‚cone injector‘ concept successful and promising
- Adequate hybrid mechanical interfaces demonstrated
- New hyperboloid thrust chamber contour numerically validated
- CMC application potential for ADN thruster (orbital propulsion)
- CMCs principally interesting for hybrid propulsion systems
- Preburner application (in particular ox-rich)
- Up-coming SLM technology
- Alternative propellants
Outlook
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 19
Thank you for your attention!
DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 20