Abstract: Ceramic cowlings that are used as the connection between a hot gas source and a Stirling engine or a turbine. The ceramic cowling is designed and fabricated with non-dusting, high temperature, dense, low thermal expansion ceramic. It must also be highly resistant to thermal shock. Also, a combination of a ceramic cowling and a shroud for covering and holding the ceramic cowling on a Stirling engine or turbine such that hot gases can flow through the ceramic cowling and into the heat exchanger coil of the Stirling engine and exhaust in a controllable manner. A method of enhancing the power efficiency of a Stirling engine and with systems including the use of at least one enhanced power Stirling engine.

Abstract: Disclosed is a thermal shield element comprising a hot side that is to face a hot medium, a cold side which is to face away from the hot medium, circumferential areas that connect the hot side to the cold side, and a material volume which is delimited by the hot side, the cold side, and the circumferential areas. The inventive thermal shield element is characterized in that the material volume encompasses at least two material zones which differ regarding the thermal expansion coefficient thereof.

Abstract: The gas turbine comprises an annular combustion chamber having inner and outer walls made of ceramic matrix composite material, and a high pressure turbine nozzle secured to a downstream end of the combustion chamber and comprising a plurality of stationary airfoils extending between the inner and outer walls of an annular flow path through the nozzle for the gas stream coming from the combustion chamber. The turbine nozzle is made of ceramic matrix composite material and it is connected to the downstream end of the combustion chamber by brazing.

Abstract: The high pressure turbine nozzle is mechanically connected to a downstream end portion of the combustion chamber and inner and outer connection ferrules connect the combustion chamber and turbine nozzle assembly to inner and outer metal shrouds in order to hold said assembly between the shrouds. Locking members are provided to prevent the turbine nozzle from turning about its axis relative to at least one of the metal shrouds so as to prevent the forces that are exerted on the turbine nozzle vanes by the flow of gas coming from the chamber being taken up by the portions of the connection ferrules that extend between the combustion chamber and turbine nozzle assembly and the metal shrouds.

Abstract: Aspects of the invention relate to a system for attaching a ceramic liner to a non-ceramic combustor head-end component while accommodating different rates of thermal expansion of these components. One end of the liner can be received within a slot formed by the combustor head-end component. The liner can be held in place by a plurality of pins with each pin passing through a pair of aligned openings in the liner and the head-end component. The end of the liner can be spaced from each of the walls of the slot so as to form a series of gaps in fluid communication. The gaps and the pins can accommodate differential thermal expansion between the liner and combustor head-end component. If desired, at least some of the gaps can be tightly controlled so as to regulate air leakage around the end of the liner.

Abstract: The annular combustion chamber having ceramic matrix composite material walls is mounted inside a metal casing by linking members fastened to the chamber by brazing. The linking members comprise a plurality of inner linking tabs and a plurality of outer linking tabs connecting the chamber to the inner and outer metal shrouds of the casing, each linking tab has a first portion fastened to the outside surface of a wall of the combustion chamber by brazing, the first portions of the linking tabs being spaced apart from one another circumferentially so that the brazed connections between the chamber and the linking members occupy a set of limited zones that are spaced apart from one another.

Abstract: A method of cooling a gas turbine engine component having a perforate metal wall includes providing a plurality of pores in the wall, wherein the pores extend substantially perpendicularly through the wall, and wherein the pores are covered and sealed closed at first ends thereof by a thermal barrier coating disposed over a first surface of the wall, and providing a plurality of film cooling holes in the wall, wherein the holes extend substantially perpendicularly through the wall and the thermal barrier coating. The method also includes providing cooling fluid to the plurality of pores and the plurality of film cooling holes along a second surface of the wall, channeling the cooling fluid through the pores for back side cooling an inner surface of the thermal barrier coating, and channeling the cooling fluid through the holes for film cooling an outer surface of the thermal barrier coating.

Abstract: A wall element (29) for a wall structure (21) of a gas turbine engine combustor (15). The wall element (29) comprises a main member (36) with an upstream edge region (30) and a downstream edge region (31). A plurality of heat removal members (38) are provided on the main member (36). The downstream edge (35) of the wall element and/or the downstream facing surface of the heat removal members closest to the downstream edge (35) are provided with a thermally resistant coating.

Abstract: A gas turbine combustion chamber includes combustion chamber tiles 3 attached to a supporting structure 6 of the gas turbine combustion chamber, with each tile possessing at least one dilution air hole 4 which is flush with a dilution air hole of the supporting structure 6, wherein a diameter of the dilution air hole of the supporting structure 6 is considerably larger than a diameter 14 of the dilution air hole 4 of the combustion chamber tile 3.

Abstract: A combustor liner includes panels joined together at a cooling nugget including a bridge and a lip extending therefrom. The lip defines a slot terminating in an outlet. Thermal barrier coating covers inboard surfaces of the panels and lip with a nominal thickness. The lip has a distal end at the slot outlet which is spaced from the coating aft of the slot less than about the coating nominal thickness.

Abstract: A high temperature splash plate for use in the combustor of a gas turbine engine that also is a specular optical reflector. A thin layer of a high temperature reflector is applied to the surface of the splash plate of the component that forms a boundary for hot combustion gases. The component typically includes a thermal barrier coating overlying the high temperature metallic component that permits the component to operate at elevated temperatures. The thermal barrier coating must be polished in order to provide a surface that can suitably reflect the radiation into the gas flow path. A thin layer of the high temperature reflector then is applied over the polished thermal barrier coating by a process that can adequately adhere the reflector to the polished surface without increasing the roughness of the surface. The surface reflects radiation in the direction of the turbine back into the hot gas flow path. The reflected radiation is not focused onto any other hardware component.

Abstract: A high temperature gas turbine component for use in the gas flow path that also is a specular optical reflector. A thin layer of a high temperature reflector is applied to the flow path surface of the component, that is, the surface of the component that forms a boundary for hot combustion gases. The component typically includes a thermal barrier coating overlying the high temperature metallic component that permits the component to operate at elevated temperatures. The thermal barrier coating must be polished in order to provide a surface that can suitably reflect the radiation into the gas flow path. A thin layer of the high temperature reflector then is applied over the polished thermal barrier coating by a process that can adequately adhere the reflector to the polished surface without increasing the roughness of the surface. The high temperature reflector can be applied to any surface aft of the compressor, such as on a combustor wall. The surface reflects radiation back into the hot gas flow path.

Abstract: A mounting assembly for an aft end of a liner of a gas turbine engine combustor including a support member, wherein a longitudinal centerline axis extends through the gas turbine engine. The mounting assembly includes a pin member extending through each one of a plurality of circumferentially spaced openings in a portion of the support member for the combustor and into a plurality of partial openings formed in the aft end of the liner, with each pin member including a head portion at one end thereof, and a device positioned within each opening in the support member so as to retain the pin members therein. The pin members and the support member are able to slide radially and/or axially with respect to the liner aft end as the support member experiences thermal growth greater than the liner.

Abstract: An assembly for providing a seal at an aft end of a combustor liner for a gas turbine engine including a longitudinal centerline axis extending therethrough. The sealing assembly includes a substantially annular first sealing member positioned between an aft portion of a support member and the liner aft end so as to seat on a designated surface portion of the liner aft end and a substantially annular second sealing member positioned between the support member aft portion and a turbine nozzle located downstream of the liner aft end so as to seat on a designated surface portion of the support member aft portion. Accordingly, the first sealing member is maintained in its seated position as the support member aft portion moves radially with respect to the liner aft end and the second sealing member is maintained in its seated position as the support member aft portion moves axially with respect to the turbine nozzle.

Abstract: Disclosed is a lining for inner walls of combustion chambers having essentially plate-shaped shielding elements arranged on the inner walls so as to leave a gap. To introduce efficient sealing off against the penetration of hot gases, porous flow barriers capable of being inserted into the gap between adjacent shielding elements are used. The flow barriers are effective in reducing the air demand. The reduction in the air demand has a positive effect on the stability of the burner flames, the effectiveness of the machine and the pollutant emissions and makes it possible to have a further increase in performance, while adhering to maximum given material temperatures.

Abstract: The invention relates to a heat-shield brick, in particular for lining a combustion chamber wall, comprising a hot side that can be exposed to a hot medium, a wall side that lies opposite said hot side and a peripheral side that lies adjacent to the hot side and the wall side and that has peripheral lateral face. A tensioning element, pre-stressed in the peripheral direction is provided on the peripheral side, whereby a compressive stress generated perpendicularly to the peripheral lateral face. The invention also relates to a combustion chamber comprising a combustion chamber lining, which has heat-shield bricks of this type and to a gas turbine comprising a combustion chamber.

Abstract: An assembly that includes a CMC article, and a method of forming the assembly, so as to reduce the likelihood during a transient thermal condition of the CMC article becoming interlocked with the worn shank of a fastener used to secure the article to its support structure. The CMC article has oppositely-disposed first and second surfaces, a hole through the article and intersecting the first and second surfaces so as to define oppositely-disposed first and second openings at the first and second surfaces, respectively, and continuous chamfers along the entirety of the first and second openings. The assembly further includes a support structure adjacent the article, and a fastener received in the hole of the CMC article and securing the CMC article to the support structure. The location of the chamfers at the perimeter of each opening eliminates a relatively sharp edge that could interlock with the worn shank of the fastener during a thermal excursion.

Abstract: A fastener (31) is disclosed for a first article, for example a tile (29), to a second article for example an outer wall (27) of a wall structure (21, 22) of a combustor (15) of a gas turbine engine (10). The fastener (31) comprises an elongate main part, which may be in the form of a bolt (34). The main part can extend through the first and second articles. The main part defines a cooling pathway (46) therethrough to allow the passage of a cooling gas through the main part.

Abstract: The invention relates to a heat shield block, in particular for lining a combustion chamber wall, having a hot side which can be subjected to a hot medium, a wall side opposite the hot side, and a peripheral side adjoining the hot side and the wall side. A tension element which can be prestressed to a prestress (Fz) is attached to the peripheral side, release of a fragment, formed during a fracture, of the heat shield block being reliably prevented by the prestress (Fz) of the tension element. The invention also relates to the use of a heat shield block, in particular for lining a combustion chamber wall of a gas turbine.

Abstract: A turbomachine has inner and outer annular shells of metal material containing, in a gas flow direction F, a fuel injector assembly, an annular combustion chamber of composite material, and an annular nozzle of metal material forming the fixed-blade inlet stage of a high pressure turbine. Provision is made for the combustion chamber to be held in position between the inner and outer metal annular shells by a plurality of flexible metal tabs having first ends interconnected by a metal ring fixed securely to each of the annular shells by first fixing means, and second ends fixed by second fixing means on a ring of composite material fixed securely to said composite material combustion chamber, the flexibility of said metal fixing tongues allowing expansion to take place freely in a radial direction at high temperatures between said composite material combustion chamber and said metal annular shells.

Abstract: A turbomachine comprises an annular shell of metal material containing in a gas flow direction F a fuel injection assembly, an annular combustion chamber of composite material, and an annular nozzle of metal material forming the fixed-blade inlet stage of a high pressure turbine, said nozzle being supported by the annular shell and being fixed thereto by first releasable fixing means, and provision being made for the combustion chamber to be mounted in floating manner inside the annular shell and held in position solely by the nozzle to which it is fixed in resilient manner by second releasable fixing means.

Abstract: A turbomachine has inner and outer annular shells of metal material containing, in a gas flow direction F, a fuel injector assembly, a composite material annular combustion chamber, and an nozzle of metal material forming the fixed-blade inlet stage of a high pressure turbine. Provision is made for the combustion chamber to be held in position between the inner and outer metal annular shells by a plurality of flexible metal tabs having first ends interconnected by a flange-forming metal ring fixed securely to each of the annular shells by first fixing means, and second ends fixed by second fixing means firstly to the composite material combustion chamber and secondly to one end of a composite material wall whose other end forms a bearing plane for a sealing element secured to the nozzle and providing sealing for the gas stream between the combustion chamber and the nozzle.

Abstract: In a turbomachine comprising inner and outer annular shells of metal material containing, in a gas flow direction F: a fuel injection assembly; an annular combustion chamber of composite material having a plurality of air feed orifices in a primary zone and/or in a dilution zone; and an annular nozzle of metal material forming the inlet stage with fixed blades for a high pressure turbine, provision is made for the combustion chamber to be held in position between the inner and outer annular metal shells by means of a plurality of flexible metal tongues having first ends interconnected by respective metal rings fixed securely to said inner and outer annular shells by first fixing means, and second ends fixed securely to the composite material combustion chamber at it air feed orifices by second fixing means.

Abstract: A wall element (29) for a wall structure (21) of a gas turbine engine combustor (15). The wall element (29) comprises a main member (36) with an upstream edge region (30) and a downstream edge region (31). A plurality of heat removal members (38) are provided on the main member (36). The downstream edge (35) of the wall element and/or the downstream facing surface of the heat removal members closest to the downstream edge (35) are provided with a thermally resistant coating.

Abstract: A seal structure for sealing a gap between a combustor liner and a neighboring structure adjacent to the combustor liner, includes an annular sealing member mounted on the neighboring structure so as to be in contact with an annular outer surface of the combustor liner to seal the gap between the combustor liner and the neighboring structure. The annular sealing member includes a plurality of sealing segments which are arranged in an annular form as a whole.

Abstract: An annular combustion chamber has outer and inner axially-extending side walls of composite material and an end wall of metal material held in position on the outer and inner side walls by fixing means. Provision is made for the fixing means to pass through annular cavities that are designed to receive cylindrical end portions of the outer and inner side walls, and that are created between peripheral edges of the end wall and facing portions folded downstream, a determined amount of clearance J being provided between the peripheral edges and the facing faces of the outer and inner side walls in such a manner as to allow expansion to take place freely, in operation, in a radial direction between said end wall and said side walls.

Abstract: In an annular combustion chamber having inner and outer axially-extending side walls of composite material and an end wall of metal material, provision is made to hold the end wall in position between the side walls of the annular combustion chamber by means of a plurality of flexible metal tongues fixed firstly to said side walls by fixing means and secondly to said end wall by brazing or welding, said end wall further comprising means for providing sealing between said end wall and said side walls.

Abstract: The liner supporting structure for the annular combustor of the present invention has a flexible connection member of connecting a structural body constituting a part of the combustor to a liner, the connection member having a first end positioned on a side of the structural body and a second end positioned on a side of the liner, the second end being disposed on an outer surface of an end part of the liner positioned on the side of the structural body; a first fixing means of fixing the first end of the connection member to the structural body; and a second fixing means of fixing the second end of the connection member to the end part of the liner.

Abstract: The invention relates to a heat-shield brick (1), in particular for lining a combustion chamber wall, comprising a hot side (3) that can be exposed to a hot medium, a wall side (5) that lies opposite said hot side (3) and a peripheral side (7) that lies adjacent to the hot side (3) and the wall side (5) and that has a peripheral lateral face (9). A tensioning element (11), pre-stressed in the peripheral direction is provided on the peripheral side (7), whereby a compressive stress is generated perpendicularly to the peripheral lateral face. The invention also relates to a combustion chamber comprising a combustion chamber lining, which has heat-shield bricks of this type and to a gas turbine comprising a combustion chamber.

Abstract: A component formed at least in part by a CMC material and equipped with an integrally-formed surface feature, such as an airflow enhancement feature in the form of a turbulator or flow guide. The CMC material comprises multiple sets of tows woven together to form a preform that is infiltrated with a matrix material. The surface feature is integrally defined at a surface of the cooling passage by an insert member disposed between adjacent tows of at least one of the tow sets. The insert member has a cross-sectional size larger than the adjacent tows, forming a protrusion in the preform that defines the surface feature in the infiltrated, consolidated and cured CMC material.

Abstract: A ceramic member support structure for a gas turbine supports ceramic members so that the ceramic members may not be easily damaged and damage to one of the ceramic members may not affect the rest of the ceramic members easily. Ceramic members (26, 27) of a gas turbine with which a combustion gas comes into contact are supported on housings (14, 14A) by metallic support members (28, 29, 41). The ceramic members (26, 27) are connected to the support members (29, 41) by elastic members (32, 42) so as to be movable relative to the support members (29, 41). The difference in thermal expansion between the ceramic members (26, 27) and the metallic support members (28, 29, 41) is absorbed by the elastic members (32, 42) and hence the ceramic members (26, 27) are not damaged easily. Since the ceramic members (26, 27) are supported by the metallic support members (28, 29, 41), damage to one of the ceramic members does not influence easily on the other ceramic member.

Abstract: A combustor liner is provided on its backside cooling surface with a braze alloy coating and cooling enhancement material, preferably metallic particles to enhance the heat transfer between the liner and the cooling medium. The surface area of the backside coated area is increased substantially by the coating and particles in relation to the uncoated surface areas. Consequently, the life of the liner is extended.

Abstract: A combustor having liners made from ceramic matrix composite materials (CMC's) that are capable of withstanding higher temperatures than metallic liners. The ceramic matrix composite liners are used in conjunction with mating components that are manufactured from superalloy materials. To permit the use of a combustor having liners made from CMC materials in conjunction with metallic materials used for the mating forward cowls, and aft seals with attached seal retainer over the broad range of temperatures of a combustor, the combustor is designed to allow for the differential thermal expansion of the differing materials at their interfaces in a manner that does not introduce stresses into the liner as a result of thermal expansion and also balances the flow of cooling air as a result of the thermal expansion.

Abstract: A turbine combustion system comprises at least one turbine combustion component that is provided with a protective coating on the outer surface otherwise known as the non-flame surface of at least one turbine combustion component. The at least one turbine combustion component comprises an inner surface that defines a hot flame path area and an outer surface. The coating on the outer surface of the combustion components provides protection for the components from at least one of oxidation, nitridation, and hot corrosion.

Abstract: A combustor liner has a stepped combustor liner surface and an overhang portion forming an air cooling slot. A contoured rear facing edge of the overhang portion reduces turbulence of combustion gas flow and reduces a combustor liner surface area exposed to combustion gases. A thermal barrier coating is also applied to the contoured rear facing edge, reducing heat flow into the overhang portion and hence reducing the operating temperature of the combustor liner. Thus, the amount of cooling air is reduced, which can reduce exhaust emissions increase engine performance and extend a working life of the combustor assembly.

Abstract: A porous, low-conductivity material formed of metal or ceramic fibers provides the burner face of a gaseous fuel combustor for gas turbines capable of minimizing nitrogen oxides (NOX) emissions in the combustion product gases. A preferred burner face, when fired at atmospheric pressure, yields radiant surface combustion with interspersed areas of blue flame combustion. A rigid but porous mat of sintered metal fibers with interspersed bands of perforations is illustrative of a preferred burner face that can be fired at pressures exceeding 3 atmospheres at the rate of at least about 500,000 BTU/hr/sf/atm. By controlling the excess air admixed with the fuel in the range of about 40% to 150% to maintain an adiabatic flame temperature in the range of about 2600° F. to 3300° F., the NOX emissions are suppressed to 5 ppm and even below 2 ppm. At all times, carbon monoxide and unburned hydrocarbons emissions do not exceed 10 ppm, combined.

Abstract: A wall element (29) for a wall structure (21) of a gas turbine engine combustor (15). The wall element (29) comprises a main member (36) with an upstream edge region (30) and a downstream edge region (31). A plurality of heat removal members (38) are provided on the main member (36). The downstream edge (35) of the wall element and/or the downstream facing surface of the heat removal members closest to the downstream edge (35) are provided with a thermally resistant coating.

Abstract: A combustor liner has a stepped combustor liner surface and an overhang portion forming an air cooling slot. A contoured rear facing edge of the overhang portion reduces turbulence of combustion gas flow and reduces a combustor liner surface area exposed to combustion gases. A thermal barrier coating is also applied to the contoured rear facing edge, reducing heat flow into the overhang portion and hence reducing the operating temperature of the combustor liner. Thus, the amount of cooling air is reduced, which can reduce exhaust emissions, increase engine performance and extend a working life of the combustor assembly.

Abstract: The invention relates to a ceramic lining for combustion chambers (8) consisting of several juxtaposed elements (1) that are attached by means of a retention bolt (3) to the inside of a thermally highly stressed metallic support wall (4), whereby at least one insulation body (5) is arranged between the metallic support wall (4) and the ceramic elements (1). It is characterized in that the ceramic elements (1) essentially have the shape of a straight, regular pyramid whose base (2) has n corners, preferably three corners, and faces the combustion chamber (8), and into whose tip the retention bolt (3) is integrated.

Abstract: A hybrid combustor for a gas turbine engine includes a plurality of circularly arrayed ceramic can combustors whose outlets communicate with the inlet of an annular, metal combustor. The combustion process is continuous through the plurality of can combustors and into the single annular combustor. Preferably only fuel-rich combustion occurs within each of the can combustors, and fuel-lean combustion continues within the single annular combustor.

Abstract: A separator (10) for a two-head combustion chamber which comprises a plurty of segments (11) that are arrayed circumferentially and spaced apart a distance (x) from one another and that are affixed to appropriately shaped sheetmetal (12, 13) of the combustion chamber. Each segment (11) is shaped as an elongated, hollow body with an upper wall (17) and a lower wall (18). The segments are cooled by air from an upstream portion of an end of the combustion chamber entering the segments (11) and exhausting by a plurality of orifices (34) in the segment walls (17, 18). The upper and lower segment walls (17) and (18) are respectively heat-protected by two sets of tiles (40, 41), each of the tiles straddling two adjacent segments, being spaced a distance from the walls (17, 18) to subtend an enclosed space (46, 47) therebetween, and having at least one tile orifice (48, 49, 50) which allows the cooling air from the segments (11) to exhaust into the heads.

Abstract: A heat shield configuration, particularly for protecting supporting structures and structural parts of gas turbine plants against a hot fluid. The heat shield has an inner lining which consists of a heat-resistant material and which is composed of plate-shaped heat-shield elements resistant to high temperatures and are disposed next to one another with gaps in-between. The heat-shield elements cover the entire area of the structural part. The heat shield is anchored so as to be thermally movable to the supporting structures and parts by bolts. The heat-shield element is formed of an erosion-proof and corrosion-proof material. An insulating brick formed of a refractory ceramic is disposed in each case between each heat-shield element and the supporting structure or parts.

Abstract: A method of preventing water erosion and hot corrosion in a combustor of a gas turbine engine, wherein water is injected into said combustor for NOx abatement, which involves the step of applying a dense vertically cracked thermal barrier coating to certain components thereof The dense vertically cracked thermal barrier coating has a porosity of less than approximately 8% and a tensile strength in the range of approximately 4-7 ksi. The dense vertically cracked thermal barrier coating is applied to such combustor components so as to produce a segmented ceramic structure having macrocracks formed therein which are oriented substantially perpendicular to an interface of the combustor component and the segmented ceramic structure.

Abstract: A combustor liner includes an inner layer for facing combustion gases, and an opposite outer layer for facing a cooling fluid. The outer layer has a greater coefficient of thermal conductivity than the inner layer for reducing temperature gradients in the liner. In a preferred embodiment, the outer layer significantly reduces temperature gradients in the liner which are caused by the varying cooling ability of impingement cooling air jets for more uniformly cooling the combustor liner and reducing the maximum temperature thereof.

Abstract: A metal or ceramic matrix composite part and corresponding method are provided exhibiting desired heat transfer characteristics. The part has a metal or ceramic matrix composite substrate and a multilayer dielectric coating. The coating has high reflectivity at wave lengths corresponding to radiation wavelengths of various combustion gases and has low reflectance at radiation wavelengths corresponding to the substrate. The multilayer coating allows the heat generated external of the part at wavelengths corresponding to combustion gases to be reflected from the part while permitting radiation wavelengths associated with the substrate to pass through the coating. The parts are useful for use in combustive gas atmospheres.

Abstract: The present invention relates generally to a low emission can-annular combustion system for an industrial gas turbine engine to satisfy increasingly stringent environmental requirements. The combustion system of the present invention employs a dual mode combustion technique to meet engine operability requirements and high power emission targets without the use of combustor diluent injection or post combustor exhaust treatment. A lean premix combustion mode is utilized to minimize primary zone combustion temperatures and limit the oxide of nitrogen production during high power engine operation. A pilot-starting auxiliary fueling system is utilized to augment the main premix fueling system. The lean premix combustion mode is enabled by a lean premix dome having a fixed axial swirler with radial fuel pathways connecting to a circumferential main fuel manifold for distributing the fuel more uniformly across the flow path.

Abstract: A combustor liner includes an inner layer for facing combustion gases, and an opposite outer layer for facing a cooling fluid. The outer layer has a greater coefficient of thermal conductivity than the inner layer for reducing temperature gradients in the liner. In a preferred embodiment, the outer layer significantly reduces temperature gradients in the liner which are caused by the varying cooling ability of impingement cooling air jets for more uniformly cooling the combustor liner and reducing the maximum temperature thereof.

Abstract: Combustion plants and gas turbines have inner regions through which hot gases flow. A thermal insulation for the inner regions includes a fiber composite having a surface. A protection against destruction is disposed at the surface.

Abstract: A combustor for a gas turbine engine having a porous outer metallic shell and a thin-walled, nonporous ceramic liner whose backside is impingement cooled, with only primary and secondary openings in the ceramic liner for delivering pressurized primary air and dilution air to the combustion zone.

Abstract: A multipiece combustor has a portion thereof being made of a plurality of ceramic segments. Each of the plurality of ceramic segments have an outer surface and an inner surface. Each of the plurality of ceramic segments have a generally cylindrical configuration and including a plurality of joints. The joints define joint portions, a first portion defining a surface being skewed to the outer surface and the inner surface. The joint portions have a second portion defining a surface being skewed to the outer surface and the inner surface. The joint portions further include a shoulder formed intermediate the first portion and the second portion. The joints provide a sealing interlocking joint between corresponding ones of the plurality of ceramic segments.