German Aerospace Center (DLR) Institute of Structures...

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


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


- 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


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


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


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)


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


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


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


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


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


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


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)


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


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


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


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


Thank you for your attention!


German Aerospace Center (DLR) Institute of Structures...

Documents

Transcript of German Aerospace Center (DLR) Institute of Structures...