CMC Rocket Thrust Chamber Technology Status and Perspectives M. - PowerPoint PPT Presentation

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 CMC Rocket Thrust Chamber Technology Status and Perspectives M. Ortelt, H. Hald, D. Koch markus.ortelt@dlr.de German Aerospace Center (DLR) Institute of Structures and Design AIRBUS DS – Space Systems - 6 th R&T DAYS, Paris, 19.11.2015, Session 3, WG2 Technologies for Future Liquid Propulsion Knowledge for Tomorrow

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 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 3 CMC thrust chamber – Design concept Features - Decoupling of single components – no bonding - Decoupling of mechanical and thermal loads - Specific hybrid interface technologies - Selective inner liner design Transpiration cooled CMC thrust chamber – design principle

DLR-ST-IBT > M. Ortelt / H. Hald / D. Koch > Presentation > AIRBUS DS – Space Systems - 6th R&T DAYS > Paris > France, 19.11.2015 • Chart 4 Functional aspects - Standard CFD systems (FLUENT, CFX, …) are constructive (pure flow coupling) - Ongoing tool-development for ‚structure-flow-coupling‘ (TAU)  - Investigations on materials out-flow homogeneity

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 System analysis of transpiration cooling 12% - Comparison of chamber size (scaling) Tw = 800 K dc = 50 mm 10% - 50 mm chamber demonstration dc = 100 mm dc = 200 mm 8% Coolant ratio [ ‐ ] dc = 440 mm dc = 1000 mm 6% - O/F = 5.5 (injector) 4% - Contraction ratio 6.25 2% - Characteristic chamber length l*=1.84 m - 7 % coolant ratio 0% 1 10 100 1000 10000 100000 Vacuum thrust [kN] - Damage free operation 12% - Amount of coolant depends on Tw = 1200 K dc = 50 mm 10% dc = 100 mm - Hotgas conditions, A s , T dc = 200 mm 8% Coolant ratio [ ‐ ] - D + p  required coolant ratio dc = 440 mm 6% dc = 1000 mm - Further coolant ratio reduction potential 4% - Chamber length can be shortened 2% 0% 1 10 100 1000 10000 100000  High operational efficiency predicted Vacuum thrust [kN]

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 Processes for Manufacturing of Nonoxide CMC C, SiC fibers in C, SiC, SiC(N) matrix Polymer Infiltration Liquid Silicon Chemical Vapor Combi-Process and Pyrolysis Infiltration Infiltration (PIP+LSI, CVI+LSI) (PIP) (LSI) (CVI) Focus at DLR Stuttgart preform (e.g. fabrics, filament winding) preform (e.g. fabrics, filament winding) fibrer coating fiber-coating if necessary infiltration (e.g. RTM) infiltration (e.g. RTM) with Si-precursor (e.g. polysilazane) with C-precursor 3-6 times to pyrolysis (inert atmosphere) decrease  carbon matrix porosity siliconisation pyrolysis (inert gas, T>1420 ° C, Si+C  SiC) (inert atmosphere, T~1000 ° C, intermediate machining  stoichiometric SiC-matrix e.g. polysilazane  SiCN-matrix) joining finishing finishing PIP LSI Koch et al., DLR Werkstoffkoll. 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 7 Processing of Ceramic Matrix Composites (DLR-ST, BT) Autoclave • 30 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 8 Thrust chamber – potential CMC derivatives Initial 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) Open porosity  [%] (porosity + permeability k d / k f adaptable by manufacturing process) 10 35 18 7 Density kg/cm 3 2.3 2.6 1.6 1.6

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 – Components Porous metal injector Elements of oxide CMC for the LOX injection Integrated ‚BlackEngine‘ demonstrator,  cyl. 50 mm Porous CMC injector C/C-SiC face-plate Applied injector systems for  cyl. 50 mm Inner liner segment Co-axial injector

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 CMC thrust chamber – mechanical interfaces Characteristic 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 11 Thrust chamber - hot gas verification (LOX/LH2; LOX/GH2)  cyl. 50 mm Vulcain contour Structure tests  cyl. 80 mm P8 C/C damages Vulcain contour CMCs damage free C/C damage free near injector!  cyl. 50 mm P8 (2005) Component tests LOX / LH2, 65  70 bar P8, 2008 52 s, 5  6 kg/s, τ = 4.2 % 120 s, p c = 55 bar, LOX / LH2 operation,  = 15 % 90 bar tests Efficiency test Contraction 6.25 P6.1  cyl. 50 mm 2012 CMCs damage free P6.1 firing test Dec 2013 20 s, p c = 55 bar, LOX / GH2 (120 K),  = 9 % Demonstration of the integrated CMC TCA

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 Thrust chamber – pressure loads during hot-run Segmented chamber module  cyl. 50 mm Inner liner: Initial model material C/C O/F = 5.5 Adequate pressure drops at 8 % C/C porosity!

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 Thrust chamber – thermal loads during hot-run P6.1, 2012  cyl. 50 mm Cooling turned off Nominal hotrun-sequence O/F = 2.0 O/F = 5.5;  = 6.72 %; p c  55 bar Max. T surface  1800 K Temperature [K] 1800 � � ����� � 750 �/� U_T_1_8 [K] 1600 U_T_2_8 [K] { U_T_3_8 [K] 1400 U_T_4_8 [K] U_T_1_9 [K] 1200 U_T_2_9 [K] 1000 U_T_3_9 [K] U_T_4_9 [K] 800 600 400 200 0 0 10 20 30 Time [s] 40

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 Design CMC injector (‚ Cone Injector ‘) LH2O/GN2 principle Spray test Features, goals, results DLR ‐ Cone Injector test campaign 'IZ2' ‐ MASS FLOW CURVES 0,3 0,25 M ‐ GH2 ‐ COOL [kg/s] 0,2 0,15 M ‐ GH2 ‐ INJ [kg/s] 0,1 M ‐ LOX ‐ INJ [kg/s] Demonstrator 0,05 0 ‐ 6000 ‐ 1000 4000 9000 14000 Time [s] DLR ‐ Cone injector test campaign 'IZ2' ‐ PRESSURE CURVES 60 50 P_I_GH1 [bar] 40 P_I_GH2 [bar] 30 P_I_GH3 [bar] P_I_LOX [bar] 20 U_P_IGN [bar] 10 Channel morphology 0 Start-up Steady state ‐ 6000 ‐ 1000 4000 9000 14000 Time [s] Initially successful hotruns, P6.1, Dec 2013 Mechanical design Flow design

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 – Orbital propulsion size Hyperboloid geometry Hyperboloid chamber design Perfectly combined with ‚cone injector‘ technology Numerical comparison Comparison referred to typical 500 N class at similar performance - 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 16 Hyperboloid chamber contour – Comparison VINCI size Dynamic viscosity Without insert Including insert Heat flux Contour Mass flow Total heat flux Specific heat flux [MW/m 2 ] [kg/s] [MW] Classical 43 16 55 Hyperboloid without insert 42 16 56 Hyperboloid including insert 43 27 52 Performance

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 Future potential - Preburner Features - Standard injector technology concerning … - functional design - mixture ratio - Propellant overhead injected through chamber wall - Long life and light weight structures, similar to CMC Application principle thrust chamber design (oxide CMCs for ox-rich systems)