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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.
Citations
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Patent
24 Mar 2020
TL;DR: In this paper, a turbine case made of a metal material is attached to a turbine stator vane and the turbine case is hooked to a hanger made of metal material with a front portion and a rear portion of the outer band portion engaged with the front and rear hook portions of the hanger.
Abstract: A turbine is provided which is simple in structure and which allows a gas now passage to be formed using a CMC over a wide range while suppressing thermal stress on turbine stator vanes, thereby achieving further improved jet engine performance and reduced fuel consumption. The turbine includes a turbine stator vane formed of a CMC and including an outer band portion and an inner band portion extending continuously from an airfoil portion is hooked to a hanger made of a metal material with a front portion and a rear portion of the outer band portion engaged with a front hook portion and a rear hook portion of the hanger, respectively, and the hanger in turn is attached to a turbine case made of a metal material.
Patent
21 Jan 2021
TL;DR: A blade outer air seal assembly includes a support structure as discussed by the authors, which consists of a plurality of segments that extend circumferentially about an axis and mounted in the support structure via a carrier.
Abstract: A blade outer air seal assembly includes a support structure. A blade outer air seal has a plurality of segments that extend circumferentially about an axis and mounted in the support structure via a carrier. At least one of the plurality of segments has a base portion that extends between a first circumferential side and a second circumferential side and from a first axial side to a second axial side. A first hook extends from the base portion near the first axial side and faces towards the second axial side. A second hook extends from the base portion near the second axial side and faces towards the first axial side. A slot is in the second hook configured to receive a pin.
Patent
27 Sep 2018
TL;DR: In this paper, an assembly for a gas turbine engine includes a carrier and a supported component comprising ceramic matrix composite materials, including an annular runner and a plurality of inserts that extend radially outward away from the runner.
Abstract: An assembly for a gas turbine engine includes a carrier and a supported component comprising ceramic matrix composite materials. The supported component includes an annular runner and a plurality of inserts that extend radially outward away from the annular runner. The inserts extend into the carrier to couple the supported component with the carrier.
Patent
24 May 2012
TL;DR: In this article, a gas turbine engine with a ceramic matrix composite (CMC) static structure (60) and a rotor module (62) with a multiple of CMC airfoils (66), a radial growth of said rotor module was matched with said CMC static structure, a corresponding method of tip clearance control was also provided.
Abstract: A gas turbine engine (20) includes a ceramic matrix composite (CMC) static structure (60) and a rotor module (62) with a multiple of CMC airfoils (66), a radial growth of said rotor module (62) matched with said CMC static structure (60). A corresponding method of tip clearance control is also provided.
Patent
23 Jan 2020
TL;DR: In this paper, a gas turbine engine includes a compressor section and a turbine section, and an attachment block is supported on structure within the engine, which mounts the blade outer air seal and communicates with cooling holes through a radially inner face of the attachment block.
Abstract: A gas turbine engine includes a compressor section and a turbine section. The turbine section includes at least one turbine rotor having a radially extending turbine blade. The turbine section is rotatable about an axis of rotation. A blade outer air seal is positioned radially outwardly of a radially outer tip of the at least one turbine blade. The blade outer air seal has axially spaced forward and aft portions and a central web between the axially spaced portions. An attachment block is supported on structure within the engine. The attachment block mounts the blade outer air seal. A passage extends into a central chamber within the attachment block, and communicates with cooling holes through a radially inner face of the attachment block to direct cooling air at the central web of the blade outer air seal. A blade outer air seal is also disclosed.
References
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Patent
11 Feb 1999
TL;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 1991
TL;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