Simulation of Combustor Damage Mechanisms and Material System Performance via a Sub-Element Configured Specimen Test

2011 ◽  
Author(s):  
Nagaraja Rudrapatna ◽  
Benjamin H. Peterson
Keyword(s):  
Gas Turbine ◽  
Thermal Fatigue ◽  
Bond Coat ◽  
Test Method ◽  
Material System ◽  
Engine Operation ◽  
Damage Mechanisms ◽  
Liner Material ◽  
Material Systems ◽  
Combustor Liner

Modern gas turbine combustors are made of high temperature alloys, employ effusion cooling and are protected by a Thermal Barrier Coating (TBC). Standard material characterization tests such as creep, oxidation and low cycle fatigue are indicators of a material’s potential performance but they neither fully represent the combustor geometric/material system nor fully represent the thermal fatigue conditions a combustor is subjected to during engine operation. Combustor rig tests and/or engine cyclic endurance tests to determine the suitability of new material systems for combustors are time consuming and costly. Therefore, a simple test method for screening material systems under representative combustor conditions is needed. This experimental system was recently developed at Honeywell Aerospace to characterize various gas turbine combustor damage mechanisms and assess state-of-the-art and developmental materials. A configured specimen is fabricated using materials and processes similarly to actual combustor hardware, including sheet metal forming, welding, TBC coating, and effusion hole laser drilling. The configured specimen is cyclically exposed to hot spot thermal gradients typically experienced by fielded hardware using a jet-fueled burner and heated cooling air. Damage mechanisms simulated include bond coat oxidation, TBC spallation, thermal fatigue and distortion. A summary of these damage mechanisms and lessons learned from test development are presented. Results from recent combustor liner, bond coat, and top coat material modifications are also discussed. The effect of combustor liner material creep and thermal fatigue resistance, bond coat composition and processing, and TBC composition and structure on combustor durability is presented.

An Experimental System for Assessing Combustor Durability

2010 ◽  
Author(s):  
Nagaraja S. Rudrapatna ◽  
Benjamin H. Peterson ◽  
Daniel Greving
Keyword(s):  
Gas Turbine ◽  
Thermal Fatigue ◽  
Failure Modes ◽  
Low Cycle Fatigue ◽  
Experimental System ◽  
Test Method ◽  
Material System ◽  
Engine Operation ◽  
Standard Material ◽  
Material Systems

Modern gas turbine combustors are made of high temperature alloys, employ effusion cooling and are protected by a Thermal Barrier Coating (TBC). Gas turbine combustor failure modes, such as TBC spallation, cracking and distortion resulting from oxidation, creep and thermal fatigue, are driven by hot spot peak temperature and the associated thermal gradient. Standard material characterization tests such as creep, oxidation and low cycle fatigue are indicators of a material’s potential performance but they neither fully represent the combustor geometric/material system nor fully represent the thermal fatigue conditions a combustor is subjected to during engine operation. Combustor rig tests and/or engine cyclic endurance tests to determine the suitability of new material systems for combustors are time consuming and costly. Therefore, a simple yet efficient test method for screening material systems under representative combustor conditions is needed. An experimental system has been developed to fill this gap. This paper discusses the configured specimen geometry, test methodology, observed test results and a comparison with typical failure modes observed in combustors.

An Experimental System for Assessing Combustor Durability

2010 ◽  
Vol 133 (4) ◽  
Author(s):  
Nagaraja S. Rudrapatna ◽  
Benjamin H. Peterson ◽  
Daniel Greving
Keyword(s):  
Gas Turbine ◽  
Thermal Fatigue ◽  
Failure Modes ◽  
Low Cycle Fatigue ◽  
Experimental System ◽  
Test Method ◽  
Material System ◽  
Engine Operation ◽  
Standard Material ◽  
Material Systems

Modern gas turbine combustors are made of high temperature alloys, employ effusion cooling, and are protected by a thermal barrier coating (TBC). Gas turbine combustor failure modes, such as TBC spallation, cracking, and distortion resulting from oxidation, creep, and thermal fatigue, are driven by hot spot peak temperature and the associated thermal gradient. Standard material characterization tests, such as creep, oxidation, and low cycle fatigue are indicators of a material’s potential performance but they neither fully represent the combustor geometric/material system nor fully represent the thermal fatigue conditions a combustor is subjected to during engine operation. Combustor rig tests and/or engine cyclic endurance tests to determine the suitability of new material systems for combustors are time-consuming and costly. Therefore, a simple yet efficient test method for screening material systems under representative combustor conditions is needed. An experimental system has been developed to fill this gap. This paper discusses the configured specimen geometry, test methodology, observed test results, and a comparison with typical failure modes observed in combustors.

Ceramic Stationary Gas Turbine Development

1993 ◽  
Author(s):  
Mark van Roode ◽  
William D. Brentnall ◽  
Paul F. Norton ◽  
Gregory P. Pytanowski
Keyword(s):  
Silicon Carbide ◽  
Silicon Nitride ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Ceramic Materials ◽  
Ceramic Composites ◽  
Department Of Energy ◽  
Engine Operation ◽  
Component Design ◽  
Combustor Liner

A program has been initiated under the sponsorship of the Department of Energy (DOE), Office of Industrial Technology, to improve the performance of stationary gas turbines in cogeneration through the selective replacement of metallic hot section parts with uncooled ceramic components. It is envisioned that the successful demonstration of ceramic gas turbine technology, and the systematic incorporation of ceramics in existing and future gas turbines will enable more efficient engine operation, resulting in significant fuel savings, increased output power, and reduced emissions. The program which started in September, 1992, takes an engine of the Solar Centaur family of industrial gas turbines, and modifies the design of the hot section to accept ceramic first stage blades and first stage nozzles, and a ceramic combustor liner. The ceramic materials selected for the blade are silicon nitride, for the nozzle silicon nitride and silicon carbide, and for the combustor liner silicon carbide as well as two continuous fiber reinforced ceramic composites, one with a silicon carbide matrix and another with an oxide matrix. This paper outlines the approach, conceptual component design, and materials selection for the program.

Durability of Oxide/Oxide CMCs in Gas Turbine Combustors

2012 ◽  
Author(s):  
Mark van Roode ◽  
Arun K. Bhattacharya
Keyword(s):  
Water Vapor ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Rupture Strength ◽  
Aluminum Silicate ◽  
Inlet Temperature ◽  
Liner Material ◽  
Barrier Coating ◽  
Liner System ◽  
Combustor Liner

An integrated creep rupture strength degradation and water vapor degradation model for gas turbine oxide-based CMC (Ceramic Matrix Composite) combustor liners was expanded with heat transfer computations to establish maximum TRIT (Turbine Rotor Inlet Temperature) for gas turbines with 10:1 pressure ratio. Recession rates and average CMC operating temperatures were calculated for an existing baseline N720/A (N720/Al2O3) CMC combustor liner system, with and without protective Al2O3 FGI (Friable Graded Insulation) for 30,000-h liner service life. The potential for increasing TRIT by YAG (Y3Al5O12) substitution for the fiber, matrix and FGI constituents of the CMC system was explored, because of the known superior creep and water vapor degradation resistance of YAG compared to Al2O3. It was predicted that uncoated N720/A can be used as a combustor liner material up to a TRIT of ∼1200°C, offering no TRIT advantage over a conventional metal + TBC (Thermal Barrier Coating) combustor liner. A similar conclusion was previously reached for a SiC/SiC CMC liner with BSAS-type EBC (Barium Strontium Aluminum Silicate Environmental Barrier Coating). The existing N720/A + Al2O3 FGI combustor liner system can be used at a maximum TRIT of ∼1350°C, a TRIT increase over metal + TBC and uncoated N720/A of ∼150°C. Replacing the Al2O3 with YAG is predicted to increase the maximum allowable TRIT. Substitution of the fiber or matrix in N720/A increases TRIT by ∼100°C. A YAG FGI improves the TRIT of the 720/A + Al2O3 FGI by ∼50°C, enabling a TRIT of ∼1400°C, similar to that predicted for SiC/SiC CMCs with protective rare earth monosilicate EBCs.

Durability of Oxide/Oxide Ceramic Matrix Composites in Gas Turbine Combustors

2013 ◽  
Vol 135 (5) ◽  
Author(s):  
Mark van Roode ◽  
Arun K. Bhattacharya
Keyword(s):  
Water Vapor ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Ceramic Matrix ◽  
Inlet Temperature ◽  
Oxide Ceramic ◽  
Liner Material ◽  
Barrier Coating ◽  
Liner System ◽  
Combustor Liner

An integrated creep rupture strength degradation and water vapor degradation model for gas turbine oxide-based ceramic matrix composite (CMC) combustor liners was expanded with heat transfer computations to establish the maximum turbine rotor inlet temperature (TRIT) for gas turbines with 10:1 pressure ratio. Recession rates and average CMC operating temperatures were calculated for an existing baseline N720/A (N720/Al2O3) CMC combustor liner system with and without protective Al2O3 friable graded insulation (FGI) for 30,000-h liner service life. The potential for increasing TRIT by Y3Al5O12 (YAG) substitution for the fiber, matrix, and FGI constituents of the CMC system was explored, because of the known superior creep and water vapor degradation resistance of YAG compared to Al2O3. It was predicted that uncoated N720/A can be used as a combustor liner material up to a TRIT of ∼1200  °C, offering no TRIT advantage over a conventional metal + thermal barrier coating (TBC) combustor liner. A similar conclusion was previously reached for a SiC/SiC CMC liner with barium strontium aluminum silicate (BSAS)-type environmental barrier coating (EBC). The existing N720/A + Al2O3 FGI combustor liner system can be used at a maximum TRIT of ∼1350  °C, a TRIT increase over metal + TBC, and uncoated N720/A of ∼150  °C. Replacing the Al2O3 with YAG is predicted to increase the maximum allowable TRIT. Substitution of the fiber or matrix in N720/A increases TRIT by ∼100  °C. A YAG FGI improves the TRIT of the N720/A + Al2O3 FGI by ∼50  °C, enabling a TRIT of ∼1400 °C, similar to that predicted for SiC/SiC CMCs with protective rare earth monosilicate EBCs.

Ceramic Stationary Gas Turbine Development Program: Second Annual Summary

1995 ◽  
Author(s):  
Mark van Roode ◽  
William D. Brentnall ◽  
Paul F. Norton ◽  
Bryan D. Edwards
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Fatigue Testing ◽  
The United States ◽  
United States Department ◽  
Ceramic Component ◽  
Support Structure ◽  
Engine Operation ◽  
Ceramic Components ◽  
Combustor Liner

A program is being performed under the sponsorship of the United States Department of Energy, Office of Industrial Technologies, to improve the performance of stationary gas turbines in cogeneration through the selective replacement of hot section components with ceramic parts. Solar Turbines Incorporated leads a team that includes major U.S. and offshore suppliers of ceramic components, recognized test laboratories and a cogeneration enduser to develop and demonstrate ceramic insertion in a stationary gas turbine with the objectives of more efficient engine operation, resulting in significant fuel savings, increased output power, and reduced emissions. The engine selected for the program, the Centaur 50 is being retrofitted with first stage ceramic blades, first stage ceramic nozzles, and a ceramic combustor liner. The engine hot section is being redesigned to accommodate the ceramic parts to the existing metallic support structure. Detailed design of the ceramic components and of the interfacing metallic support structure has been completed. Two blade designs with different attachments and a nozzle design with a modified airfoil geometry have been developed. Three combustor liner designs are being evaluated based on monolithic tiles or rings, or integral cylinders of continuous fiber-reinforced ceramic matrix composites (CFCC). Fabrication of first generation prototype blades and nozzles is in progress. Fabrication of subscale combustor hardware has been completed. Materials property data are being gathered in support of the ceramic component design and life prediction. Fast fracture and dynamic fatigue testing were performed for the candidate blade and nozzle materials. Creep and oxidation testing is in progress. Nondestructive methodologies are being applied to test specimens, simulated components, subscale hardware and prototype components. A Centaur 50 engine was procured and has been modified for ceramic component testing in a full-size engine configuration.

Performance Prediction of Gas Turbine Under Different Strategies Using Low Heating Value Fuel

2013 ◽  
Author(s):  
Edson Batista da Silva ◽  
Marcelo Assato ◽  
Rosiane Cristina de Lima
Keyword(s):  
Natural Gas ◽  
Gas Turbine ◽  
Equivalence Ratio ◽  
Safe Operation ◽  
Engine Operation ◽  
Heating Value ◽  
Gasification Process ◽  
Surge Margin ◽  
And Performance ◽  
Industrial Gas Turbine

Usually, the turbogenerators are designed to fire a specific fuel, depending on the project of these engines may be allowed the operation with other kinds of fuel compositions. However, it is necessary a careful evaluation of the operational behavior and performance of them due to conversion, for example, from natural gas to different low heating value fuels. Thus, this work describes strategies used to simulate the performance of a single shaft industrial gas turbine designed to operate with natural gas when firing low heating value fuel, such as biomass fuel from gasification process or blast furnace gas (BFG). Air bled from the compressor and variable compressor geometry have been used as key strategies by this paper. Off-design performance simulations at a variety of ambient temperature conditions are described. It was observed the necessity for recovering the surge margin; both techniques showed good solutions to achieve the same level of safe operation in relation to the original engine. Finally, a flammability limit analysis in terms of the equivalence ratio was done. This analysis has the objective of verifying if the combustor will operate using the low heating value fuel. For the most engine operation cases investigated, the values were inside from minimum and maximum equivalence ratio range.

Assessment of Solar Steam Injection in Gas Turbines

2016 ◽  
Author(s):  
C. Kalathakis ◽  
N. Aretakis ◽  
I. Roumeliotis ◽  
A. Alexiou ◽  
K. Mathioudakis
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Fuel Efficiency ◽  
Thermal Power ◽  
Steam Injection ◽  
Power Production ◽  
Engine Operation ◽  
Efficiency Increase ◽  
Solar Receiver ◽  
Steam Production

The concept of solar steam production for injection in a gas turbine combustion chamber is studied for both nominal and part load engine operation. First, a 5MW single shaft engine is considered which is then retrofitted for solar steam injection using either a tower receiver or a parabolic troughs scheme. Next, solar thermal power is used to augment steam production of an already steam injected single shaft engine without any modification of the existing HRSG by placing the solar receiver/evaporator in parallel with the conventional one. For the case examined in this paper, solar steam injection results to an increase of annual power production (∼15%) and annual fuel efficiency (∼6%) compared to the fuel-only engine. It is also shown that the tower receiver scheme has a more stable behavior throughout the year compared to the troughs scheme that has better performance at summer than at winter. In the case of doubling the steam-to-air ratio of an already steam injected gas turbine through the use of a solar evaporator, annual power production and fuel efficiency increase by 5% and 2% respectively.

Siemens SGT-800 Industrial Gas Turbine Enhanced to 50MW: Combustor Design Modifications, Validation and Operation Experience

2013 ◽  
Author(s):  
Daniel Lörstad ◽  
Annika Lindholm ◽  
Jan Pettersson ◽  
Mats Björkman ◽  
Ingvar Hultmark
Keyword(s):  
Combustion Chamber ◽  
Gas Turbine ◽  
Cooling System ◽  
Good Condition ◽  
Combined Cycle ◽  
Combustion Stability ◽  
Combustion System ◽  
Design Work ◽  
Engine Operation ◽  
Cooling Air

Siemens Oil & Gas introduced an enhanced SGT-800 gas turbine during 2010. The new power rating is 50.5MW at a 38.3% electrical efficiency in simple cycle (ISO) and best in class combined-cycle performance of more than 55%, for improved fuel flexibility at low emissions. The updated components in the gas turbine are interchangeable from the existing 47MW rating. The increased power and improved efficiency are mainly obtained by improved compressor airfoil profiles and improved turbine aerodynamics and cooling air layout. The current paper is focused on the design modifications of the combustor parts and the combustion validation and operation experience. The serial cooling system of the annular combustion chamber is improved using aerodynamically shaped liner cooling air inlet and reduced liner rib height to minimize the pressure drop and optimize the cooling layout to improve the life due to engine operation hours. The cold parts of the combustion chamber were redesigned using cast cooling struts where the variable thickness was optimized to maximize the cycle life. Due to fewer thicker vanes of the turbine stage #1, the combustor-turbine interface is accordingly updated to maintain the life requirements due to the upstream effect of the stronger pressure gradient. Minor burner tuning is used which in combination with the previously introduced combustor passive damping results in low emissions for >50% load, which is insensitive to ambient conditions. The combustion system has shown excellent combustion stability properties, such as to rapid load changes and large flame temperature range at high loads, which leads to the possibility of single digit Dry Low Emission (DLE) NOx. The combustion system has also shown insensitivity to fuels of large content of hydrogen, different hydrocarbons, inerts and CO. Also DLE liquid operation shows low emissions for 50–100% load. The first SGT-800 with 50.5MW rating was successfully tested during the Spring 2010 and the expected performance figures were confirmed. The fleet leader has, up to January 2013, accumulated >16000 Equivalent Operation Hours (EOH) and a planned follow up inspection made after 10000 EOH by boroscope of the hot section showed that the combustor was in good condition. This paper presents some details of the design work carried out during the development of the combustor design enhancement and the combustion operation experience from the first units.

The Study of Thermal Properties and Micro Structures of YSZ

2007 ◽  
Author(s):  
S. M. Guo ◽  
M. B. Silva ◽  
Patrick F. Mensah ◽  
Nalini Uppu
Keyword(s):  
Thermal Conductivity ◽  
Thermal Properties ◽  
Gas Turbine ◽  
Bond Coat ◽  
Sintering Temperature ◽  
Heating Process ◽  
Thermal Barrier ◽  
Inlet Temperature ◽  
Low Thermal Conductivity ◽  
Gas Tight

Thermal barrier coatings (TBCs) are used in gas turbine engines to achieve a better efficiency by allowing increased turbine inlet temperature and decreasing the amount of cooling air used. Plasma spraying is one of the most reliable methods to produce TBCs, which are generally comprised of a top coating of ceramic and a bond-coat of metal. Usually, the top coating is Yttria-Stabilized-Zirconia (YSZ), providing the thermal barrier effect. The bond-coat is typically a layer of M-Cr-Al-Y (where “M” stands for “metal”), employed to improve the attachment between the ceramic top-coat and the substrate. Due to the extreme temperature gradient presented in the plasma jet and the wide particle size distribution, during the coating process, injected ceramic powders may experience a significantly different heating process. Different heating history, coupled with the substrate preheating temperature, may affect the thermal properties of the YSZ layers. In this paper, four sets of mol 8% YSZ disks are fabricated under controlled temperatures of 1100°C, 1200°C, 1400°C and 1600°C. Subsequently the thermal properties and the microstructures of these YSZ disks are studied. The results indicate a strong microstructure change at a temperature slightly below 1400°C. For a high sintering temperature, a dense YSZ layer can be formed, which is good for gas tight operation; At low sintering temperature, say 1200°C, a porous YSZ layer is formed, which has the advantage of low thermal conductivity. For gas turbine TBC applications, a robust low thermal conductivity YSZ layer is desirable, while for Solid Oxide Fuel Cells, a gas-tight YSZ film must be formed. This study offers a general guideline on how to prepare YSZ layers, mainly by controlling the heating process, to form microstructures with desired properties.

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