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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24 Nov 2015TL;DR: In this paper, a turbine shroud assembly or blade track assembly adapted to extend around a turbine wheel assembly is disclosed, which includes a carrier and a blade track coupled to the carrier.
Abstract: A turbine shroud assembly or blade track assembly adapted to extend around a turbine wheel assembly is disclosed. The turbine shroud assembly includes a carrier and a blade track coupled to the carrier. The blade track is movable between a radially-inward position having a first inner diameter and a radially-outward position having a second inner diameter larger than the first inner diameter.
22 citations
Patent•
18 Jun 2013TL;DR: In this article, a radial position control assembly for a gas turbine engine stage includes a case structure with an annular recess, and a support ring is received in the recess to maintain the supported structure relative to the sealing structure at a clearance during thermal transients.
Abstract: A radial position control assembly for a gas turbine engine stage includes a case structure. A supported structure is operatively supported by the case structure. The supported structure includes a hook providing an annular recess. A support ring is received in the recess. The supported structure and the support ring have different coefficients of thermal expansion. A sealing structure is adjacent to the supported structure. The support ring maintains the supported structure relative to the sealing structure at a clearance during thermal transients based upon a circumferential gap between adjacent supported structure and based upon a radial gap between the support ring and the supported structure.
20 citations
Patent•
06 Jan 2009TL;DR: In this article, a turbine blade platform may be disposed between two turbine blades and the platform may include a first exterior side configured to interface with a first turbine blade and a second exterior side disposed generally opposite the first turbine side.
Abstract: In one embodiment, a turbine blade platform may be disposed between two turbine blades. The platform may include a first exterior side configured to interface with a first turbine blade. The platform also may include a second exterior side disposed generally opposite the first exterior side and configured to interface with a second turbine blade.
19 citations
Patent•
12 Jan 2006TL;DR: In this article, a flaired tubular geometry is used to fabricate a turbine shroud ring segment for a gas turbine engine, which is then attached to a surrounding support structure.
Abstract: Fabricating a refractory component for a gas turbine engine, such as a turbine shroud ring segment, by arranging refractory fiber tows ( 24 ) in a flaired tubular geometry ( 20 ) comprising a stem portion ( 21 ) and a funnel-shaped portion ( 22 ); impregnating the refractory fibers ( 24 ) with a ceramic matrix to form a flaired tube ( 20 ) of ceramic composite matrix material; at least partially filling the funnel-shaped portion ( 22 ) with a ceramic core ( 30 ) extending beyond the end of the funnel-shaped portion to provide a working gas containment surface ( 31 ); curing the flaired tube ( 20 ) and the ceramic core ( 30 ) together; cutting the funnel-shaped portion ( 22 ) to provide rectangular edges ( 27 ); and providing an attachment mechanism ( 34, 36, 38, 40 ) on the stem portion ( 21 ) for attaching the component to a surrounding support structure. Additional tows ( 24 ) may be introduced at intermediate stages to maintain a desired fabric density.
19 citations
Patent•
29 Apr 2010TL;DR: In this article, a gusset (40A-G) between two CMC walls (26, 28) has fibers oriented diagonally to oppose in tension a wall-spreading moment of the walls about the intersection.
Abstract: A gusset (40A-G) between two CMC walls (26, 28) has fibers (23) oriented to provide anisotropic strengthening of the wall intersection (34). The fibers (23) may be oriented diagonally to oppose in tension a wall-spreading moment of the walls (26, 28) about the intersection (34). Interlocking features (46, 48, 52, 56, 58) may be provided on the gusset to improve load sharing between the gusset and the walls. The gusset may have one or more diagonal edges (50, 51) that contact matching edges of a slot (42, 42D, 43D) to oppose wall-spreading (M1) and wall-closing (M2) bending of the walls (26, 28). The gusset may be installed in the slot after preparing the gusset and the walls to different temperatures. Then the assembly may be final-fired to produce differential shrinkage that causes compression of the gusset or the wall intersection.
18 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