US20180340687A1 - Refractory ceramic component for a gas turbine engine - Google Patents

Refractory ceramic component for a gas turbine engine Download PDF

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Publication number
US20180340687A1
US20180340687A1 US15/603,898 US201715603898A US2018340687A1 US 20180340687 A1 US20180340687 A1 US 20180340687A1 US 201715603898 A US201715603898 A US 201715603898A US 2018340687 A1 US2018340687 A1 US 2018340687A1
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Prior art keywords
liner
gas turbine
turbine engine
bricks
component
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Abandoned
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US15/603,898
Inventor
Jay A. Morrison
Jerry Klopf
Alexander Ralph Beeck
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Siemens AG
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Siemens AG
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Filing date
Publication date
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Priority to US15/603,898 priority Critical patent/US20180340687A1/en
Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: Klopf, Jerry
Assigned to SIEMENS ENERGY, INC. reassignment SIEMENS ENERGY, INC. ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: MORRISON, JAY A., BEECK, ALEXANDER RALPH
Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SIEMENS ENERGY, INC.
Publication of US20180340687A1 publication Critical patent/US20180340687A1/en
Application status is Abandoned legal-status Critical

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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/007Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00Stators
    • F01D9/02Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23MCASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M5/00Casings; Linings; Walls
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23MCASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M5/00Casings; Linings; Walls
    • F23M5/02Casings; Linings; Walls characterised by the shape of the bricks or blocks used
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002Wall structures
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00Materials; Properties thereof
    • F05D2300/20Oxide or non-oxide ceramics
    • F05D2300/21Oxide ceramics
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23MCASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M2900/00Special features of, or arrangements for combustion chambers
    • F23M2900/05004Special materials for walls or lining

Abstract

A refractory ceramic component for a gas turbine engine is employed that is more cost effective than typical components used in the gas turbine engine. The refractory ceramic component may be a refractory ceramic liner that is easily replaceable. The refractory ceramic liner may be a unitary construction or made of numerous bricks that are interlocked. The ceramic used is a refractory oxide material.

Description

    BACKGROUND 1. Field
  • Disclosed embodiments are generally related to components in gas turbine engines.
  • 2. Description of the Related Art
  • Gas turbines comprise a casing or cylinder for housing a compressor section, a combustion section, and a turbine section. A supply of air is compressed in the compressor section and directed into the combustion section. The compressed air enters the combustion inlet and is mixed with fuel. The air/fuel mixture is then combusted to produce high temperature and high pressure (working) gas. This working gas then travels through the transition and into the turbine section of the turbine.
  • The turbine section may comprise rows of vanes which direct the working gas to the airfoil portions of the turbine blades. The working gas travels through the turbine section, causing the turbine blades to rotate, thereby turning a rotor associated therewith.
  • Higher efficiency of a combustion turbine can be achieved by increasing the temperature of the working gas flowing through the combustion section to as high a temperature as is practical. The aggressive hot gas, however, can degrade various metal turbine components, such as the combustor, transition ducts, vanes, ring segments, and turbine blades as it flows through the turbine.
  • For this reason, strategies have been developed to protect turbine components from extreme temperatures, such as the development and selection of high temperature materials adapted to withstand these extreme temperatures and cooling strategies to keep the components adequately cooled during operation. Superalloys with additional protective coatings are commonly used for hot gas path components of gas turbines. In view of the substantial and longstanding development in the area of superalloys, further increases in the temperature capability of superalloys has become more difficult.
  • Ceramic matrix composite (CMC) materials have been developed and increasingly utilized in gas turbine engines. Typically, CMC materials include a ceramic or a ceramic matrix material, either of which hosts a plurality of reinforcing fibers. The fibers may have a predetermined orientation to provide the CMC materials with additional mechanical strength. Generally, (fiber reinforced) ceramic matrix composites are manufactured by the infiltration of a matrix slurry (e.g., alumina, mullite, silicon-containing polymers, molten silicon, or the like) into a fiber preform. While these materials may offer a higher temperature resistance than superalloys, fiber grains of the CMC may coarsen and result in reduced strength over time. In addition, matrix grain coarsening can result in CMC embrittlement leading to a propensity for cracking and crack propagation as firing temperatures increase.
  • While both of the above strategies are frequently employed, their usage can necessitate additional cooling strategies and can be undesirably expensive for certain applications.
  • SUMMARY
  • Briefly described, aspects of the present disclosure relate to a component for a gas turbine engine.
  • An aspect of present disclosure may be a gas turbine engine comprising a combustor basket. A cone may be connected to the combustor basket. The cone comprises a liner; and a shell surrounding the liner, wherein the liner is formed from a refractory oxide ceramic material and the shell is formed from metal.
  • Another aspect of the present disclosure may be a component for a gas turbine engine. The gas turbine engine may comprise a liner; and a shell surrounding the inner layer, wherein the inner layer is formed from a refractory oxide ceramic material and the shell is formed from metal.
  • BRIEF DESCRIPTION OF THE DRAWINGS
  • FIG. 1 is a side view of a gas turbine engine employing a cone using refractory ceramic bricks to form the liner.
  • FIG. 2 is a cut away view of a cone formed with the refractory ceramic bricks.
  • FIG. 3 is another view of the refractory ceramic bricks illustrating the mating of the ceramic bricks.
  • FIG. 4 is a view of a refractory ceramic brick and an interlayer.
  • FIG. 5 is a view of the cone and its connection to the combustor.
  • FIG. 6 is cross sectional view of a cone that uses a metal spring located between the shell and the liner.
  • DETAILED DESCRIPTION
  • To facilitate an understanding of embodiments, principles, and features of the present disclosure, they are explained hereinafter with reference to implementation in illustrative embodiments. Embodiments of the present disclosure, however, are not limited to use in the described systems or methods.
  • A ceramic liner is proposed for use in the gas turbine engines that is able to provide the protection desired for the gas turbine engine. This is a ceramic liner that avoids the costs associated with CMC. FIG. 1 shows a portion of gas turbine engine 100 having a cone 10 that utilizes the proposed ceramic liner 12. The cone 10 is connected to an integrated exit piece (IEP) 14 and transmits the working fluids through the gas turbine engine 100. While the application discusses the proposed ceramic liner 12 with reference to a cone 10, it should be understood that the ceramic liner 12 may be used for other gas turbine engine components, such as transitions, blade tip seals, and interstage turbine ducts.
  • FIG. 1 shows the cone 10 connected to the combustor basket 8. The cone 10 is formed by the liner 12 and the shell 13. Combustor 6 produces working gases within the combustor basket 8. The working gases then flow into the cone 10. As discussed above these working gases are very hot. The liner 12 is formed from a refractory ceramic material. Shell 13 surrounds the liner 12. The shell 13 is made out of metal. An air gap 9 may be present between the liner 12 and the shell 13. If an air gap 9 is utilized, air may flow through the air gap 9, which can advantageously assist in cooling the liner 12 and the shell 13. Alternately, the air gap can be maintained as a stagnant gap and utilized to promote radiant heat transfer between the liner 12 and shell 13.
  • FIG. 2 is a cut away view of the cone 10. The cone 10 is formed with refractory ceramic bricks 15. While a plurality of bricks 15 are shown and discussed throughout the application, it is also possible to form the liner 12 as a unitary piece. The bricks 15 shown in FIG. 2 are assembled and each of the bricks 15 formed to obtain the desired shape of the cone 10. The desired shape of the cone 10 may change depending on the type of gas turbine engine in which the cone 10 is being used.
  • Forming the liner 12 with a plurality of bricks 15 makes it easier to service and/or replace the liner 12. For example, in the event that a portion of the liner 12, such as one of the bricks 15, has undergone some type of damage or has simply outlived its natural life span that particular brick 15 may be serviced. The liner 12 may provide further savings in life cycle cost with only the liner 12 needing to be replaced at combustion/transition intervals, instead of the entire cone 10.
  • Another advantage of using a plurality of bricks 15 to form the liner 12 relates to costs of using the bricks 15. The costs of the bricks 15 made from refractory oxide ceramic material are relatively inexpensive compared with other ceramic materials, or forming the liner 12 out of CMC.
  • Additional savings can be achieved by using bricks 15 formed from refractory oxide ceramic material. The refractory oxide ceramic material is able to handle higher temperatures without cooling. When used in the gas turbine engine 100 savings can be achieved by not having to provide additional cooling features that would be needed to cool metal components and part.
  • The refractory ceramic bricks 15 form a liner 12 that is heavier than typically used liners. Each of the bricks 15 has an inner surface 17 and an outer surface 19. The distance d1 (i.e. the thickness) between a point on the inner surface 17 and a point on the outer surface 19 may be greater than 20 mm, and preferably greater than 25 mm. A range for the distance d1 may be between 20-30 mm. The thicker or greater the distance d1, the more heat protection liner 12 will provide. The thickness or distance d1 also compensates for the lower durability that the bricks 15 made of oxide ceramic may have. The overall thickness of the refractory ceramic bricks 15 also distinguishes the liner 12 from other types of liners.
  • As indicated above, the use of bricks 15 reduces the need for cooling air since the bricks 15 can withstand higher operating temperatures, such as those greater than 1400° C. This greater than the temperatures that other types of materials can typically withstand. By using materials that permit the gas turbine engine 100 to operate at higher temperatures NOx emissions can be reduced. The liner 12 made of refractory oxide ceramic material can replace current metal designs, which are cooled by impingement and film cooling and need high temperature turbine alloys. The cooling of the ceramic liner 12 may be via radiation to the metallic shell 13.
  • As discussed above, the bricks 15 are made of a refractory oxide material. Some examples of the oxide ceramic materials that can be used to construct the bricks 15 are zirconium oxide, titanium oxide, aluminum oxide, mullite, combinations thereof, and the like. For example brick 15 may be composition of SiO2 and Al2O3. Preferably the brick 15 is a conglomerate of multiple phases, such as mullite and aluminium oxide. This type of oxide ceramic material is capable of being easily cast in order to form the bricks 15 necessary for the formation of the gas turbine engine component.
  • FIGS. 3 and 4 show a close up view of bricks 15 that form the liner 12. The liner 12 is shown being used in conjunction with an interlayer 11 that is located between the liner 12 and the shell 13. The interlayer 11 may be a ceramic fiber mat 11. The ceramic fiber mat 11 is used in order to provide the shell 13 additional protection from the heat as well providing a contact buffer zone between the liner 12 and the shell 12. The fiber mat 11 may also be used to dampen vibration. A further purpose of the fiber mat 11 is to minimize hot gas flow between the liner 12 and the shell 13. Another purpose of the fiber mat 11 is to provide an inward spring force on the liner segments bricks 15 to ensure positive contact under operating conditions.
  • The ceramic fiber mat 11 may be made of ceramic materials such as alumina, mullite, aluminosilicate, yttria alumina garnet, silicon carbide, silicon nitride, silicon carbon nitride, molydisicilicide, zirconium oxide, titanium oxide, combinations thereof, and the like. The fiber material used in the ceramic fiber mat 11 may comprise a non-oxide material. The fiber material may comprise ceramic fibers sold under the trademark Nextel, such as Nextel 610, and 720 fibers. In addition, fiber material may be in any suitable form, such as a straight filament, a bundle or a roving of multiple fibers, a braid, or a rope. The fiber material may comprise non-ceramic materials, including but not limited to carbon, glass, polymeric, metal, or any other suitable fiber materials.
  • Each of the bricks 15 has a first mating side 16 and a second mating side 18. The first mating side 16 and the second mating side 18 are contoured so that they complement each other. First mating side 16 is contoured outwards (i.e. convex) so that it is shaped to engage second mating side 18, which is contoured inwards (i.e. concave). The pressures exerted by adjacent bricks 15 when assembled cause the bricks 15 to mate with each other and remain in place. While the bricks 15 may mate with each other in this manner, other means for engaging each of the bricks 15 may be used, such as pins, grooves and other interlocking assemblies.
  • FIG. 5 shows a cross-sectional view of a pin 21 that can be used with the liner 12 and shell 13 in order to securely attach the cone 10 to the gas turbine engine 100. The pin 21 may be secured at different receiving locations 22 around the combustor basket 8. In the embodiment shown in FIG. 5 there may be three receiving locations 22 that are located 120° apart around the circumference of the gas turbine engine 100. The locations of the receiving locations 22 provide distribution for the circular shaped cross section of the cone 10. The receiving location 22 may be threaded so to permit the pin 21 to be threaded into the combustor basket 8.
  • Other attachment means can be used to connect the pins 21, such as bolts, etc. Preferably the pins 21 are removable so as to permit easy repair and replacement of the bricks 15. The pins 21 should be able to accommodate thermal growth of the cone 10 while limiting movement of the cone 10 in the axial and circumferential directions. Additionally the pins 21 should be able to accommodate the weight that the cone 10 may have due to the use of the bricks 15, which may be heavier than other materials typically used.
  • FIG. 6 is cross sectional view of a cone 10 that uses a metal springs 23 between the shell 13 and the liner 12. The metal springs 23 can be used in place of the air gap 9 or the interlayer 11. However, in some embodiments combinations of the air gap 9, the interlayer 11 and the metal springs 23 may be used in the area between the liner 12 and the shell 13. The springs 23 are able to accommodate thermal fluctuations of each of the bricks 15 as the impacted by the heat of the working gases. In addition to being able accommodate movement of the entire liner 12 as it thermally fluctuates, the use of the springs 23 can also isolate movement of one brick 15 versus movement of another brick 15 within the liner 12. The springs 23 may also be used to dampen vibration loads. Another purpose of the springs 23 to provide an inward spring force on the liner segments bricks 15 to ensure positive contact under operating conditions. The use of a metal spring 23 may also require cooling air to be passed between the liner and the shell. As an alternative springs 23 may be wave springs made of CMC. Having the springs made of CMC can enable the springs 23 to handle higher temperatures without the need for cooling air. The springs 23 may also be used to dampen vibration. Another purpose of the springs 23 to provide an inward spring force on the liner segments bricks 15 to ensure positive contact under operating conditions.
  • While embodiments of the present disclosure have been disclosed in exemplary forms, it will be apparent to those skilled in the art that many modifications, additions, and deletions can be made therein without departing from the spirit and scope of the invention and its equivalents, as set forth in the following claims.

Claims (20)

What is claimed is:
1. A gas turbine engine comprising:
a combustor basket;
a cone connected to the combustor basket, the cone comprising;
a liner; and
a shell surrounding the liner, wherein the liner is formed from a refractory oxide ceramic material and the shell is formed from metal.
2. The gas turbine engine of claim 1, wherein the liner has an outer surface and an inner surface, wherein the distance between the outer surface and the inner surface is greater than 20 mm.
3. The gas turbine engine of claim 1, wherein the liner is comprised of a plurality of bricks.
4. The gas turbine engine of claim 3, wherein each of the plurality of bricks has a first mating side and a second mating side, wherein the first mating side of the plurality of bricks mates with the second mating side of another of the plurality of bricks.
5. The gas turbine engine of claim 1, wherein the refractory oxide ceramic material comprises SiO2 and Al2O3.
6. The gas turbine engine of claim 1, wherein the refractory oxide ceramic material is a conglomerate of mullite and aluminum oxide.
7. The gas turbine engine of claim 1, further comprising an interlayer located between the shell and the liner.
8. The gas turbine engine of claim 7, wherein the interlayer is made of a ceramic fiber mat.
9. The gas turbine engine of claim 1, further comprising a metallic spring located between the shell and the liner.
10. The gas turbine engine of claim 1, wherein the cone is connected to the combustor basket with a plurality of pins.
11. A component for a gas turbine engine comprising:
a liner; and
a shell surrounding the liner, wherein the liner is formed from a refractory oxide ceramic material and the shell is formed from metal.
12. The component of claim 11, wherein the liner has an outer surface and an inner surface, wherein the distance between the outer surface and the inner surface is greater than 20 mm.
13. The component of claim 11, wherein the liner is comprised of a plurality of bricks.
14. The component of claim 13, wherein each of the plurality of bricks has a first mating side and a second mating side, wherein the first mating side of the plurality of bricks mates with the second mating side of another of the plurality of bricks.
15. The component of claim 11, wherein the refractory oxide ceramic material comprises SiO2 and Al2O3.
16. The component of claim 15, wherein the refractory oxide material is a conglomerate of mullite and aluminum oxide.
17. The component of claim 11, further comprising an interlayer located between the shell and the liner.
18. The component of claim 17, wherein the interlayer is made of a ceramic fiber mat.
19. The component of claim 11, further comprising a metallic spring located between the shell and the liner.
20. The component of claim 11, wherein the component is adapted to be connected to a cone combustor basket with a plurality of pins.
US15/603,898 2017-05-24 2017-05-24 Refractory ceramic component for a gas turbine engine Abandoned US20180340687A1 (en)

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US15/603,898 US20180340687A1 (en) 2017-05-24 2017-05-24 Refractory ceramic component for a gas turbine engine
PCT/US2018/032533 WO2018217485A1 (en) 2017-05-24 2018-05-14 Refractory ceramic component for a gas turbine engine

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Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2915877A (en) * 1954-03-03 1959-12-08 Parsons & Marine Eng Turbine Cylindrical furnaces
US4030875A (en) * 1975-12-22 1977-06-21 General Electric Company Integrated ceramic-metal combustor
US5957067A (en) * 1997-07-28 1999-09-28 Abb Research Ltd. Ceramic liner
US6571560B2 (en) * 2000-04-21 2003-06-03 Kawasaki Jukogyo Kabushiki Kaisha Ceramic member support structure for gas turbine
US20090260364A1 (en) * 2008-04-16 2009-10-22 Siemens Power Generation, Inc. Apparatus Comprising a CMC-Comprising Body and Compliant Porous Element Preloaded Within an Outer Metal Shell
US20100260960A1 (en) * 2003-04-25 2010-10-14 Siemens Power Generation, Inc. Damage tolerant gas turbine component
US8522559B2 (en) * 2004-12-01 2013-09-03 Siemens Aktiengesellschaft Heat shield element, method and mold for the production thereof, hot-gas lining and combustion chamber
US8552559B2 (en) * 2004-07-29 2013-10-08 Megica Corporation Very thick metal interconnection scheme in IC chips
US20150251376A1 (en) * 2012-09-28 2015-09-10 General Electric Company Layered arrangement, hot-gas path component, and process of producing a layered arrangement
US20170138597A1 (en) * 2015-07-22 2017-05-18 Rolls-Royce North American Technologies Inc. Combustor tile with monolithic inserts

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EP1126221A1 (en) * 2000-02-17 2001-08-22 Siemens Aktiengesellschaft Padded refactory tile as liner for a gas turbine combustor
US7291407B2 (en) * 2002-09-06 2007-11-06 Siemens Power Generation, Inc. Ceramic material having ceramic matrix composite backing and method of manufacturing
EP2169311A1 (en) * 2008-09-29 2010-03-31 Siemens Aktiengesellschaft Material mixture for producing a fire-retardant material, fire-retardant moulding body and method for its manufacture
EP2233450A1 (en) * 2009-03-27 2010-09-29 Alstom Technology Ltd Multilayer thermal protection system and its use
US9422865B2 (en) * 2013-03-14 2016-08-23 Rolls-Royce Corporation Bi-metal fastener for thermal growth compensation

Patent Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2915877A (en) * 1954-03-03 1959-12-08 Parsons & Marine Eng Turbine Cylindrical furnaces
US4030875A (en) * 1975-12-22 1977-06-21 General Electric Company Integrated ceramic-metal combustor
US5957067A (en) * 1997-07-28 1999-09-28 Abb Research Ltd. Ceramic liner
US6571560B2 (en) * 2000-04-21 2003-06-03 Kawasaki Jukogyo Kabushiki Kaisha Ceramic member support structure for gas turbine
US20100260960A1 (en) * 2003-04-25 2010-10-14 Siemens Power Generation, Inc. Damage tolerant gas turbine component
US8552559B2 (en) * 2004-07-29 2013-10-08 Megica Corporation Very thick metal interconnection scheme in IC chips
US8522559B2 (en) * 2004-12-01 2013-09-03 Siemens Aktiengesellschaft Heat shield element, method and mold for the production thereof, hot-gas lining and combustion chamber
US20090260364A1 (en) * 2008-04-16 2009-10-22 Siemens Power Generation, Inc. Apparatus Comprising a CMC-Comprising Body and Compliant Porous Element Preloaded Within an Outer Metal Shell
US20150251376A1 (en) * 2012-09-28 2015-09-10 General Electric Company Layered arrangement, hot-gas path component, and process of producing a layered arrangement
US20170138597A1 (en) * 2015-07-22 2017-05-18 Rolls-Royce North American Technologies Inc. Combustor tile with monolithic inserts

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