Document
USA
Joseph E. Grady
for the American Ceramic Society’s
held June 2-7, 2013 in San Diego, CA
NASA Glenn Research Center
CMC Technology Advancements
for Gas Turbine Engine Applications
10th Pacific Rim Conference on Ceramic and Glass Technology
flexibility of the NAS.
in the capacity, efficiency and research needs for NextGen by will enable significant increases Airspace Systems Program capabilities, and technologies that developing revolutionary concepts, Directly address the fundamental ATM Systems research facilities.
Integrated Research Program Aviation Safety Program Aeronautics Test Program Conduct research at an integrated the benefits in a relevant environment comprehensive set of flight and ground-based system-level on promising concepts and Preserve and promote the testing capabilities of Conduct cutting-edge research that will produce one of the United States’ largest, most versatile and technologies and explore/assess/demonstrate innovative concepts, tools, and technologies to improve the intrinsic safety attributes of current and future aircraft.
NASA Aeronautics Programs
Fundamental Aeronautics Program Conduct fundamental research that will produce innovative concepts, tools, and technologies to enable revolutionary changes for vehicles that fly in all speed regimes.
F advantage SiC/SiC CMC offers at least at 1/3 density over superalloys Seals 400° Flaps & 2732 F Flowpath 2700 F CMC Turbine Frame Blades 2372 F 2400 F CMC Shrouds Vanes 2012 F liners Combustor
Turbine engine applications for ceramic composites
6.0% -3.1% -0.7% -0.95% -1.25% Fuel Burn Reduction -0.8% -3.6% -0.45% -4.85% Engine Wt Reduction SFC -0.5% -1.8% -3.0% -0.75% Reduction LPT CMC Overall F CMC Materials) Burner ΔP Reduction HPT Vanes (5% to 3%) o HPT Blades* “Technology” Vanes/Blades*
Summary of Fuel Burn Reduction (2700
F
°
high matrix cracking stress and
Environmental Barrier Coating
F
°
SiC/SiC composite with 20 ksi strength for 300 hours at 2700
NASA’s CMC research is focused on technology advancements needed for turbine applications • • Advanced fiber with increased creep resistance and sufficient high temperature strength • Advanced matrix with thermal conductivity • Durable • Joining & Integration • Life Prediction iBN coating Preform on each fiber between every fiber Super Sylramic-iBN for environmental protection) (in-situ grown BN surface layer (Boron removal for SOA creep-rupture resistance) Preform Treatment in High-Pressure N Preform Treatment Furnace Fiber treatment process improves on 2009 NASA patent for Sylramic-iBN fiber Sylramic” Fiber) “ Boron-Sintered (formed from commercial SiC Fiber Preform Blade Preform
Fabrication Process for 2700°F Fiber
and visualization tool fiber architecture analysis and increase thermal conductivity 3D fiber architectures suppress delamination
applications
3D-orthogonal fiber architecture
3D fiber architectures increase CMC durability for turbine
CVI BN
CVI SiC
PIP SiC
PIP SiC
etter oxidation resistance & off-axis properties
Reduced porosity; higher MCS and thermal conductivity Reduced matrix cracking B
• • •
Hybrid (CVI + PIP) SiC Matrix
• 2700°F • 20 ksi • 300 hours Requirement: CVI Full Hybrid Full PIP MI CVI of the annealed CVI SiC component Hybrid matrix performs better than PIP due to greater creep-resistance Hybrid matrix performs better than CVI due to better oxidation resistance Advanced fiber is needed to meet 2700°F turbine goal of PIP component • • •
Hybrid Matrix CMC: Durability Comparison
(g) Si(OH) SiC/SiC CMC O (g) H Environmental Barrier Coating (EBC) 1343 C 1252 C 1385 C Exposure Time (hrs) 1446 C Exposure conditions: 6 atm, 20 m/s airflow -5 -15 -10 Weight Loss of SiC in High Pressure Burner Rig SiC Wt. Loss (mg/cm )
Environmental Barrier Coating is needed for durability of SiC/SiC CMC
Weight loss rate too high for “Baseline EBC” BSAS Hafnia compounds show better resistance to recession in high pressure burner rig tests • • 0.0008 O (doped) 0.00075 Hf AS800 SN282 SiC/SIC CMC La HfO RE-Hf-luminosilicates BSAS Rare earth silicates 0.0007 -1 1200 BSAS Baseline 1/T, K 0.00065 0.0006 Temperature, °C 0.00055 SiC/SiC under high velocity Supersonics EBC stability development goal - 2005 0.0005 0.1 in High Pressure Burner Rig Testing 0.01 0.001 Specific weight change, mg/cm -h Weight Change of Candidate EBC Materials
Current State of 2700°F EBC Technology
PS-PVD Deposited Candidate Coating System for Turbine Engine CMCs Exterior of the PS-PVD Rig Same different material, processing parameters Plasma during vapor deposition to .
m in < 60sec μ x 10 0.5 m Variable microstructure Multilayer coatings Thin layers Temperatures up to 10,000K Non line-of-sight deposition Various coatings, solid oxide fuel cells, gas sensors, etc – – – – – – – system in U.S. Constructed at NASA GRC in 2008- 2010 via contract with Sulzer Metco GRC facility is second such • Bridges gap between plasma spray and vapor phase methods Low pressure (70-1400 Pa = 0.5 10 torr), high power (>100 kW) High deposition rate Spray material incorporated into gas stream Wide range of applications
Plasma Spray - Physical Vapor Deposition (PS-PVD) at NASA
tensile ISO tests for shear joint strength Joint seals & leakage Reduces part count, SiC/SiC SiC/SiC REABOND technology for crack-free & durable joints shapes Simplifies fabrication of complex
Joining capabilities can improve turbine performance
hybrid interlayer approaches Develop single, multiple and under engine conditions larger & complex components Joining technology for CMC turbine components APPROACH: • • Evaluate durability of joints • Scale-up joining processes for
to constituent properties and architectures
EBC cracking EBC / bond coat delamination
Objective: understand the sensitivity of CMC/EBC failure mechanisms
Failure defined as EBC spallation resulting from a combination of: • •
Multiscale Deformation and Life Modeling for CMC Turbine Components
F CMC
°
CMC materials
- Compare durability with current SOA
panels & coupons and test in a rig environment
• 2013: Demonstrate advanced SiC fiber • 2014: Fabricate and test 2700 • 2015: Fabricate turbine subelements
Near-Term Development Plans
turbine operating conditions High Pressure Burner Rig simulated tapes F temperature capability ° o silicon bond coat and rare earth silicate coatings EBC • 5-10 mil thick multilayer coating with hafnia- • 2700 • Plasma Spray / PVD application process Hi-Nic Type S fibers Hi-Nic Type S fibers - - BN interface coatings - 0/90/0/0/90/0 - 22% Fiber volume ratio - - BN interface coatings - 5 HS weave - 35% fiber volume ratio 1. Prepreg MI SiC/SiC 2. CVI SiC/SiC turbine vane subelement 30 hours of testing has been completed at 2500ºF, 10 atmospheres and 200 m/s gas velocity Durability testing of baseline turbine vane subelement in simulated engine environment has been initiated MI Vane Non-uniform wall thickness near ply joints Surface roughness • Delaminations in leading edge, trailing edge and rib Non-uniform wall thickness in leading edge • Porosity in fillet areas • • CVI Vane
CT scans detect subelement fabrication flaws
Isolated porosity in trailing edge 0.8 Bend Samples Subelements 0.6 0.4 Strain (%) 0.2 Panels CVI Material 100 150 200 250 300 350 400 450 500 Stress (Mpa) Panels 0.8 0.6 Subelements Strain (%) 0.4 and thermal conductivity of turbine vane subelements R 0.2 Porosity after infiltration of complex shape reduced strength, stiffness MI Material Stress (MPa) 100 200 300 400 500 600 700 800 900
Bend tests showed strength and modulus loss in fabricating turbine vane subelements
Summary
F CMC would enable a 6% reduction in
°
• A 2700 engine fuel burn due to the reduced need for turbine cooling • This requires development of an advanced fiber, fiber architecture, matrix and EBC, and is the basis of NASA’s CMC technology development program • In 2015, NASA will compare durability of an advanced 2700F CMC with current CMC systems in rig tests that simulate an engine environment