Experimental study of multi-hole cooling for integrally-woven, ceramic matrix composite walls for gas turbine applications

Solar Turbines Incorporated, under U.S. government sponsored programs, has been evaluating ceramic matrix composite combustor liners in test rigs and Solar’s Centaur® 50S gas turbine engines since 1992. The objective is to evaluate and improve the performance and durability of CMCs as high-temperature materials for advanced low emissions combustors. Field testing of CMC combustor liners started in May of 1997 and by the end of 2004, over 67,000 operating hours had been accumulated on SiC∕SiC and oxide∕oxide CMC liners. NOx and CO emissions have been consistently <15ppmv and <10ppmv, respectively. Maximum test durations of 15,144h and 13,937h have been logged for SiC∕SiC liners with protective environmental barrier coatings. An oxide∕oxide CMC liner with a Friable Graded Insulation coating has been tested for 12,582h. EBCs significantly improve SiC∕SiC CMC liner life. The basic three-layer EBC consists of consecutive layers of Si, mullite, and BSAS. The durability of the baseline EBC can be improved by mixing BSAS with mullite in the intermediate coating layer. The efficacy of replacing BSAS with SAS has not been demonstrated yet. Heavy degradation was observed for two-layer Si∕BSAS and Si∕SAS EBCs, indicating that the elimination of the intermediate layer is detrimental to EBC durability. Equivalent performance was observed when the Hi-Nicalon fiber reinforcement was replaced with Tyranno ZM or ZMI fiber. Melt infiltrated SiC∕SiC CMCs have improved durability compared to SiC∕SiC CMCs fabricated by Chemical Vapor Infiltration of the matrix, in the absence of an EBC. However, the presence of an EBC results in roughly equivalent service life for MI and CVI CMCs. Results to date indicate that oxide∕oxide CMCs with protective FGI show minor degradation under Centaur® 50S gas turbine engine operating conditions. The results of, and lessons learned from CMC combustor liner engine field testing, conducted through 2004, have been summarized.

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Damage to an A-N720 ceramic matrix composite under simulated gas turbine static component conditions using laser heating

Journal of Composite Materials ◽

10.1177/0021998318775472 ◽

2018 ◽

Vol 52 (30) ◽

pp. 4127-4138

Author(s):

Larry Lebel ◽

Rachid Boukhili ◽

Sylvain Turenne

Keyword(s):

Stress Relaxation ◽

Gas Turbine ◽

Matrix Composite ◽

Laser Heating ◽

Ceramic Matrix Composite ◽

Ceramic Matrix ◽

Operating Conditions ◽

Specimen Thickness ◽

Air Cooling ◽

Static Component

A ceramic matrix composite material constructed from a porous alumina matrix and Nextel™ 720 fibers was subjected to a novel type of experiment designed to reproduce the operating conditions of a gas turbine static component. Material specimens were exposed to cyclic laser heating on one side and active air cooling on the other side while being constrained in their bending deflection. For most specimens, accumulation of damage under the constant-amplitude heat load resulted in a temperature increase above the material limit of 1200℃ at the heated surface. Stress relaxation was observed due to inflicted sudden damage, like surface fiber buckling or ply delamination, and due to gradual permanent deformation of the material. The amount of damage was quantified by the variation of a stiffness parameter based on the measured reaction force and through-thickness temperature difference. While the level of damage was significant at the hot face, with the depth of cracking and delamination reaching up to half of the specimen thickness, the specimen back face remained intact. Stabilization of the damage was observed due to stress relaxation and decoupling of the damaged plies from the non-affected ones.

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Ceramic Matrix Composite Turbine Vanes for Gas Turbine Engines

Volume 1: Turbo Expo 2005 ◽

10.1115/gt2005-68229 ◽

2005 ◽

Cited By ~ 7

Author(s):

Venkat Vedula ◽

Jun Shi ◽

David Jarmon ◽

Scott Ochs ◽

Lola Oni ◽

...

Keyword(s):

Silicon Carbide ◽

Gas Turbine ◽

Matrix Composite ◽

Thermal Stresses ◽

Ceramic Matrix Composite ◽

Ceramic Matrix ◽

Component Reliability ◽

Turbine Engine ◽

Component Design ◽

Turbine Vanes

Ceramic matrix composite (CMC) turbine vanes, due to their high temperature capability, allow significantly higher firing temperatures with minimal cooling. Turbine vanes were designed for a gas turbine engine with special attention to attachment methods that minimize thermal stresses due to large differences in coefficients of thermal expansion between the CMC airfoil and metal platforms. Detailed aerodynamic, thermal and structural analyses were performed to ensure component reliability. The paper describes the component design, analysis, fabrication, and rig testing of a silicon carbide fiber reinforced silicon carbide matrix (SiC/SiC) turbine vane.

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Experimental Study of Relationships between Ultrasonic Attenuation and Dispersion for Ceramic Matrix Composite

Physics Procedia ◽

10.1016/j.phpro.2015.08.100 ◽

2015 ◽

Vol 70 ◽

pp. 171-174

Author(s):

A.A. Naumenko ◽

S.A. Shcherbinin ◽

D.I. Makariev ◽

A.N. Rybyanets

Keyword(s):

Experimental Study ◽

Matrix Composite ◽

Ultrasonic Attenuation ◽

Ceramic Matrix Composite ◽

Ceramic Matrix ◽

Attenuation And Dispersion

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Design and Testing for Ceramic Matrix Composite Turbine Vane

Volume 6: Ceramics; Controls, Diagnostics and Instrumentation; Education; Manufacturing Materials and Metallurgy ◽

10.1115/gt2017-63264 ◽

2017 ◽

Cited By ~ 3

Author(s):

Fumiaki Watanabe ◽

Takeshi Nakamura ◽

Yousuke Mizokami

Keyword(s):

Gas Turbine ◽

Matrix Composite ◽

Thermal Cycle ◽

Ceramic Matrix Composite ◽

Ceramic Matrix ◽

Gas Turbine Engine ◽

Inlet Temperature ◽

Turbine Engine ◽

Turbine Vanes ◽

Turbine Vane

Ceramic matrix composite (CMC) have higher temperature capability and lower density than nickel based alloys which have been used for hot section components of gas turbine engines. These properties are expected to bring many benefits, such as higher turbine inlet temperature (TIT), reduction of cooling air, and reduction of weight, when it is used as the material for hot section components of gas turbine engine. The authors have been developing CMC turbine vane for aircraft engines. In this paper, the authors present the summary of design, manufacturing, and testing, which were conducted from 2010 to 2012. The purpose of this work was to verify that the SiC-SiC CMC which IHI has developed has the applicability to aircraft turbine vanes. The concept was planned for CMC hollow turbine vanes, in which the airfoil and the platform are fabricated in CVI process. As the demonstration of this concept, the first stage turbine vane was designed with CMC for IHI IM270 that is the 2MW-class small industrial gas turbine engine. Bending rig test was conducted at room temperature in order to check the structural feasibility of the airfoil-platform joint. The outer platform of vane was fixed in the same way with the engine parts, and the load simulating the aerodynamic force was applied at the airfoil portion. The fracture load was higher than the load which the vanes receive in the actual engine. Burner rig test was conducted in order to check the durability against thermal cycle. A CMC vane was set between dummy metal vanes, and cyclically heated by gas burner. The maximum airfoil surface temperature was set to 1200 degree C, and the maximum temperature difference between airfoil and platform was about 700 degree C. The minimum airfoil temperature at the interval of heating was about 300 degree C. The time of one thermal cycle was 6 minutes that consisted of 3 minute heating and 3 minute natural cooling. The test was conducted for 1,000 cycles. In post-test inspection there was no defect like a crack. Engine test for CMC vanes was conducted using IHI IM270. The four CMC vanes were mounted into the first stage turbine nozzle assembly in place of the normal metal vanes. The test was conducted for 400 hours. The inlet temperature of CMC vanes were measured by thermocouples installed at the leading edge, and the measured temperature was about 1050 degree C at the steady state. From this work, the applicability of the design concept for the CMC vane to actual engine was verified in which airfoil-platform are fabricated in CVI process.

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133 Development of the Ceramic Matrix Composite(CMC) for Gas Turbine

The Proceedings of the Materials and processing conference ◽

10.1299/jsmemp.2001.9.283 ◽

2001 ◽

Vol 2001.9 (0) ◽

pp. 283-284

Author(s):

Ken-ichiroh Igashira ◽

Hyoe Ono ◽

Naohumi Akikawa ◽

Yoshihiro Matsuda

Keyword(s):

Gas Turbine ◽

Matrix Composite ◽

Ceramic Matrix Composite ◽

Ceramic Matrix

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Oxide Ceramic Matrix Composite Materials for Aero-Engine Applications: A Literature Review

10.3233/atde210029 ◽

2021 ◽

Author(s):

George Karadimas ◽

Konstantinos Salonitis ◽

Konstantinos Georgarakis

Keyword(s):

High Temperature ◽

Composite Materials ◽

Gas Turbine ◽

Matrix Composite ◽

Ceramic Matrix Composite ◽

Ceramic Matrix ◽

Turbine Engines ◽

Oxide Ceramic ◽

Matrix Composites ◽

Gas Turbine Engines

The development of aircraft gas turbine engines has extensively been required for the development of advanced materials. This complex development process is however justified by the system-level benefits in terms of reduced weight, higher temperature capability, and/or reduced cooling, each of which increases efficiency. This is where high-temperature ceramics have made considerable progress and ceramic matrix composites (CMCs) are in the foreground. CMCs are classified into non-oxide and oxide-based ones. Both families have material types that have a high potential for use in high-temperature propulsion applications. Typical oxide-based ones are based on an oxide fiber and oxide matrix (Ox-Ox). Some of the most common oxide subcategories, are alumina, beryllia, ceria, and zirconia ceramics. Such matrix composites are used for example in combustion liners of gas turbine engines and exhaust nozzles. However, until now a thorough study on the available oxide-based CMCs for such applications has not been presented. This paper focus on assessing a literature survey of the available oxide ceramic matrix composite materials in terms of mechanical and thermal properties.

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