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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23 Jul 2018TL;DR: In this article, a gas turbine engine includes a compressor section and a turbine section, and the turbine section is rotatable about an axis of rotation, and a blade outer air seal is positioned radially outwardly of a radially outer tip of the at least one turbine blade.
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 chamber within the attachment block, and communicates to circumferential edges of the attachment block to direct cooling air along the central web of the blade outer air seal. A blade outer air seal assembly is also disclosed.
1 citations
Patent•
28 Sep 2012
TL;DR: In this paper, the authors describe a ring with two annular main portions, one continuous on a circumference and in contact with ends of two partitions (42, 43) that are parallel to a main part (40) of a casing.
Abstract: The ring (41) has two parts (47, 51), where the parts include two annular main portions (48, 52), respectively. One of the parts is continuous on a circumference and is in contact with ends of two partitions (42, 43) that are parallel to a main part (40) of a casing. The other part is continuous on the circumference and concentric with the former part. The latter part defines a gas flow passage (2) surrounded by the casing, and is constructed by a material i.e. composite material, different from a material i.e. metal, of the former portion.
1 citations
Patent•
01 Oct 2012TL;DR: In this paper, a turbine system is disclosed, which includes a transition duct (50) having an inlet (52), an outlet (54), and a passage (56) extending between the inlet and the outlet, defining a longitudinal axis (90), a radial axis (94), and tangential axis (92).
Abstract: A turbine system is disclosed. The turbine system (10) includes a transition duct (50) having an inlet (52), an outlet (54), and a passage (56) extending between the inlet (52) and the outlet (54) and defining a longitudinal axis (90), a radial axis (94), and a tangential axis (92). The outlet (54) of the transition duct (50) is offset from the inlet (52) along the longitudinal axis (90) and the tangential axis (92). The turbine system further includes a turbine section (16) connected to the transition duct (50). The turbine section includes a plurality of shroud blocks at least partially defining a hot gas path, a plurality of buckets at least partially disposed in the hot gas path, and a plurality of nozzles at least partially disposed in the hot gas path. At least one of a shroud block, a bucket, or a nozzle includes means for withstanding high temperatures.
1 citations
Patent•
14 Apr 2020TL;DR: In this paper, a support component comprising metallic materials, a supported component comprising ceramic matrix composite materials, and an attachment pin configured to couple the supported component to the support component is presented.
Abstract: An assembly adapted for use in a gas turbine engine. The assembly includes a support component comprising metallic materials, a supported component comprising ceramic matrix composite materials, and an attachment pin configured to couple the supported component to the support component. The attachment pin includes compliant features to distributes loads applied to the supported component.
1 citations
Patent•
15 Feb 2016TL;DR: In this article, a CMC shroud hanger assembly for a gas turbine engine and a method of transferring load from a ceramic matrix composite (CMC) shroud 206 to a cMC shroud Hanger assembly 200 are provided.
Abstract: A shroud 206 and shroud hanger assembly 200 for a gas turbine engine and a method of transferring load from a ceramic matrix composite (CMC) shroud 206 to a CMC shroud hanger assembly 200 are provided. The shroud 206 and shroud hanger assembly 200 includes a shroud hanger assembly 200 formed of a first material having a first coefficient of thermal expansion, the shroud hanger assembly 200 having a forward hanger portion 202 and a separate aftward hanger portion 204 and a shroud 206 formed of a second material having a second coefficient of thermal expansion, the forward hanger portion 202 and the aftward hanger portion 204 configured to couple together to clamp across a forward radially extending wall 208 of the shroud 206.
1 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