Patent•
Ceramic matrix composite component for a gas turbine engine
09 Sep 2002-
TL;DR: In this article, the gap between the CMC member and the support member is kept purposefully small to limit the stress developed in the material when it is deflected against the support by the force of a rubbing blade tip.
Abstract: A ceramic matrix composite (CMC) component for a combustion turbine engine ( 10 ). A blade shroud assembly ( 30 ) may be formed to include a CMC member ( 32 ) supported from a metal support member ( 32 ). The CMC member includes arcuate portions ( 50, 52 ) shaped to surround extending portions ( 46, 48 ) of the support member to insulate the metal support member from hot combustion gas ( 16 ). The use of a low thermal conductivity CMC material allows the metal support member to be in direct contact with the CMC material. The gap ( 42 ) between the CMC member and the support member is kept purposefully small to limit the stress developed in the CMC member when it is deflected against the support member by the force of a rubbing blade tip ( 14 ). Changes in the gap dimension resulting from differential thermal growth may be regulated by selecting an angle (A) of a tapered slot ( 76 ) defined by the arcuate portion.
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Patent•
14 Dec 2015TL;DR: In this paper, a component of a turbine is disclosed, which includes an carrier segment having a rail, an frame segment including a hanger having a flange supported on the rail of the carrier segment, and an inner surface having a section of a track for a turbine blade defined therein.
Abstract: A component of a turbine is disclosed. The component includes an carrier segment having a rail, an frame segment including a hanger having a flange supported on the rail of the carrier segment, and an inner surface having a section of a track for a turbine blade defined therein. The component also includes a retainer segment secured to the carrier segment such that the hanger is secured between the retainer segment and the carrier segment.
8 citations
Patent•
17 Apr 2018TL;DR: In this paper, a gas turbine engine seal assembly is described, where a housing is mounted to the seal such that the impingement face is exposed to define a plenum between the housing and the impeding face, and a plurality of cooling passages define a passage axis that is oriented such that a projection of the passage axis intersects the seal body.
Abstract: A seal assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a seal that has a seal body having a sealing portion that extends from an engagement portion. The sealing portion has a seal face that extends circumferentially between first and second mate faces. The seal body defines an internal cavity that extends circumferentially between the first and second mate faces, and the engagement portion has an impingement face opposite the seal face that defines a plurality of apertures to the internal cavity. A housing is mounted to the seal such that the impingement face is exposed to define a plenum between the housing and the impingement face. The housing defines a plurality of cooling passages, and each of the plurality of cooling passages define a passage axis that is oriented such that a projection of the passage axis intersects the seal body. A method of sealing is also disclosed.
7 citations
Patent•
24 Apr 2008TL;DR: In this article, a CMC anchor is used to join a metal substrate and a ceramic thermal barrier, forming a layer of ceramic insulation locked into the honeycomb, which is then used to enclose the cells.
Abstract: A ceramic matrix composite (CMC) anchor ( 20, 100 ) joining a metal substrate ( 40 ) and a ceramic thermal barrier ( 38 ). The CMC anchor extends into and interlocks with the ceramic barrier, and extends into and interlocks with the metal substrate. The CMC anchor may be a honeycomb ( 20 ) or other extending-into-and-interlocking geometry. A CMC honeycomb may be formed with first ( 22 ) and second ( 24 ) arrays of cells ( 26 ) with open distal ends ( 28 ) on respective opposite sides of a sheet ( 30 ). The cells may have walls ( 32 ) with transverse passages ( 36 ). A metal ( 40 ) may be deposited into the cells and passages on one side of the sheet, forming a metal substrate locked into the honeycomb. A ceramic insulation material ( 38 ) may be deposited into the cells and passages on the opposite side of the sheet, forming a layer of ceramic insulation locked into the honeycomb.
7 citations
Patent•
14 Apr 2016TL;DR: In this article, a variable density coating system (81) on a component (100) is provided, where the external coating (104) has a first density area (90) and a second density area(92) with the first density areas being more dense than the second density areas (92).
Abstract: A variable density coating system (81) on a component (100) is provided. In one embodiment, the variable density coating system (81) comprises: an external coating (104) on the component (100), where the external coating (104) has a first density area (90) and a second density area (92) with the first density area (90) being more dense than the second density area (92). A gas turbine engine (10) is also provided that includes a CMC component (100) with the variable density coating system (81) thereon.
6 citations
Patent•
02 Dec 2016
TL;DR: In this article, a propulsion system includes a first compressor in fluid communication with a fluid source, and a heat exchanger is positioned proximal to the first compressor via the first conduit.
Abstract: A propulsion system includes a first compressor in fluid communication with a fluid source. A first conduit is coupled to the first compressor, and a heat exchanger is in fluid communication with the first compressor via the first conduit. A second conduit is positioned proximal to the heat exchanger. A combustor is in fluid communication with the heat exchanger via the second conduit and is configured to generate a high-temperature gas stream. A third conduit is coupled to the combustor, and a first thrust augmentation device is in fluid communication with the combustor via the third conduit. The heat exchanger is positioned within the gas stream generated by the combustor.
6 citations
References
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Patent•
11 Feb 1999TL;DR: In this article, a ceramic composition for insulating components, made of ceramic matrix composites, of gas turbines is provided, which comprises a plurality of hollow oxide-based spheres of various dimensions, a phosphate binder, and at least one oxide filler powder, whereby the binder partially fills gaps between the spheres and the filler powders.
Abstract: A ceramic composition for insulating components, made of ceramic matrix composites, of gas turbines is provided. The composition comprises a plurality of hollow oxide-based spheres of various dimensions, a phosphate binder, and at least one oxide filler powder, whereby the phosphate binder partially fills gaps between the spheres and the filler powders. The spheres are situated in the phosphate binder and the filler powders such that each sphere is in contact with at least one other sphere and the arrangement of spheres is such that the composition is dimensionally stable and chemically stable at a temperature of approximately 1600 °C. A stationary vane of a gas turbine comprising the composition of the present invention bonded to the outer surface of the vane is provided. A combustor comprising the composition bonded to the inner surface of the combustor is provided. A transition duct comprising the insulating coating bonded to the inner surface of the transition is provided. Because of abradable properties of the composition, a gas turbine blade tip seal comprising the composition also is provided. The composition is bonded to the inside surface of a shroud so that a blade tip carves grooves in the composition so as to create a customized seal for the turbine blade tip.
141 citations
Patent•
15 Nov 1985
TL;DR: A turbine ring has an annular metallic carrier which is mounted within the inside of the turbine casing and within the carrier there is provided a ceramic abradable ring as mentioned in this paper, and a cooling air circuit is provided to regulate the temperature of only the annular carrier such that the latter always exerts a centripetal compression force on the ring under all operational conditions of the gas turbine.
Abstract: A turbine ring has an annular metallic carrier which is mounted within the inside of the turbine casing and within the carrier there is provided a ceramic abradable ring. A cooling air circuit is provided to regulate the temperature of only the annular carrier such that the latter always exerts a centripetal compression force on the ring under all operational conditions of the gas turbine to clamp the abradable ring.
69 citations
Patent•
21 Apr 1987
TL;DR: In this paper, a turbine ring comprises an annular support in two parts and a ring of ceramic sectors for sealing purposes, which are interconnected by an axial groove and a sliding male part.
Abstract: A turbine ring comprises an annular support in two parts and a ring of ceramic sectors for sealing purposes. The two parts of the support are interconnected by an axial groove and a sliding male part. Each part comprises an axial annular groove in which is engaged a respective axial edge portion of each sector. The cooperating internal radial faces respectively of the edge portion and of the groove are circumferential, while the radially outer radial face of each sector comprises at least one flat zone. The corresponding face of each groove may be polygonal.
61 citations
Patent•
26 May 1983
TL;DR: In this article, the ceramic facing material of an outer air seal (30) at the leading edge region (36) is densified by a plasma gun to produce a glazed area (52) which is resistant to erosion.
Abstract: Outer air seal structures of particular suitability for use in gas turbine engines are disclosed. Techniques for improving resistance to erosion while maintaining good abradability are discussed. In one particular structure the ceramic facing material of an outer air seal (30) at the leading edge region (36) is densified by a plasma gun to produce a glazed area (52) which is resistant to erosion.
61 citations
Patent•
14 Jan 1991TL;DR: In this paper, a turbine blade shroud assembly for a gas turbine engine includes a metal substrate ring on the engine, a continous ceramic barrier ring inside the substrate ring and exposed to hot gas in a hot gas flow path of the engine and a wire mesh compliant ring between the barrier and substrate rings.
Abstract: A turbine blade shroud assembly for a gas turbine engine includes a metal substrate ring on the engine, a continous ceramic barrier ring inside the substrate ring and exposed to hot gas in a hot gas flow path of the engine, and a wire mesh compliant ring between the barrier and substrate rings. The temperature of the barrier ring increases faster than the temperature of the substrate ring as the temperature in the hot gas flow path increases. The coefficient of thermal expansion of the substrate ring is less than the coefficient of thermal expansion of the barrier ring so that the barrier ring expands relative to the substrate ring with increasing temperature in the hot gas flow path and development of tensile hoop stress in the ceramic barrier ring is minimized.
60 citations