RTX Corporation Patente

A method includes obtaining, from a first plurality of sources, production data relating to a plurality of stages of production of a part of a machine. The production data includes at least one of test data, simulation data, validation data, verification data, and reliability data for the part. The method also includes obtaining, from a plurality of second plurality of sources that differs from the first plurality of sources, manufacturing metadata that is separate from the production data. The manufacturing metadata data including data that describes machinery used to manufacture the part. The method also includes storing the production data and the manufacturing metadata in one or more blockchains. A system is also disclosed.

A method of manufacture is provided during which a first material layer (62A) is disposed with a second material layer (64) to provide a multi-layered preform (60). The first material layer (62A) lengthwise and widthwise overlaps the second material layer (64). The first material layer (62A) is configured from or otherwise includes a titanium material. The second material layer (64) is configured from or otherwise includes an aluminum material. The multi-layered preform (60) is clamped between a first die (66) and a second die (68). The first die (66) and the second die (68) each lengthwise and widthwise overlap the multi-layered preform (60). The multi-layered preform (60) clamped between the first die (66) and the second die (68) is sintered and bonded to provide a bipolar plate (36; 38) for a fuel cell (28). The titanium material in the first material layer (62A) of the bipolar plate (36; 38) is bonded to the aluminum material in the second material layer (64) of the bipolar plate (36; 38) during the sintering of the multi-layered preform (60).

An additive manufacturing apparatus (100) includes a chamber (110), a controller (140) including processing circuitry and a memory (144), the controller (140) operable to determine an expected life of an additive manufactured part (130), by receiving an initial part design and initial additive manufacturing parameters and 1) utilizing a machine learning model trained on historic images to estimate grain structure based upon the initial part design and initial additive manufacturing parameters, 2) manufacturing a part (130) by additive manufacturing based upon the initial part design and initial additive manufacturing parameters, and monitoring for the location of likely defects during the manufacturing based upon sensed information during the manufacturing, 3) inferring potential defects' associated shapes, sizes, and morphologies, combining 2) and 3) to reach a defect map, superimposing the defect map on the grain structure and determining an expected life of the part (130) based upon the superimposed defect map and grain structure. A method is also disclosed.

An additive manufacturing machine (100) forms a part (130) utilizing additive manufacturing and with a plurality of layers. A controller (140) includes processing circuitry and a memory (144), and is operable for receiving design information about a proposed part (130) to be manufactured utilizing additive manufacturing, and also receives proposed process parameters for the additive manufacturing machine (100). The controller (140) is operable to predict a temperature at each of a plurality of layers that will be found in the proposed part (130), and identify a temperature history for a plurality of the layers including identifying the time to solidify for at least one of the plurality of layers, and determine a derivative of a change in temperature of the at least one of the plurality of layers over time and predicting grain structure utilizing the derivative. The controller (140) is operable to determine an expected life based upon the predicted grain structure. A method is also disclosed.

A ceramic matrix composite (CMC) component (200) for a turbine engine includes a base (210) having a radial outer surface (212) and a radial inner surface inward surface (215). The base includes a plurality of ceramic fiber plies and a ceramic matrix. A forward flange structure (220) and an aft flange structure (230) each extend from the radial outer surface of the base. A splash plate (270) is positioned between the forward flange structure and the aft flange structure to direct cooling air flowing from a forward end of the base to an aft end of the base towards the radial outer surface of the base. The splash plate has a first end proximate the forward flange structure, a second end proximate the aft flange structure, and an inner surface facing the radial outer surface of the base and including a curved surface region (275) that curves towards the outer radial surface of the base. The CMC component is a blade outer air seal (BOAS). A BOAS assembly includes a plurality of the BOAS segments arranged to form an annular shaped structure. A method of assembling the CMC component includes: providing the CMC component and providing the splash plater plate. The turbine engine includes a fan section (22), a compressor section (24), a combustion chamber (56), and a turbine section (28). The turbine section includes at least one rotor and one or more turbine blade(s) extending radially outwardly from the at least one rotor; the blade outer air seal assembly positioned between the one or more turbine blade(s); and an outer casing (36).

An under-root fan blade (20) spacer (30) system (10) for a gas turbine engine includes a fan hub (12) having a fan hub receiver (14) with a hub receiver floor (28) and an opening configured to receive a fan blade root (18). The fan blade root (18) includes a base end (26). A gap (24) is formed between the base end (26) of the fan blade root (18) and the hub receiver floor (28). The spacer (30) disposable within the gap (24) includes a shape-memory-alloy material that allow for deformation from an original shape (32) to a deformed shape (34) and rebound back into the original shape (32). An under-root fan blade ( 20) spacer (30) is provided. A process of forming the under-root fan blade (20) spacer (30) system (10) for a gas turbine engine includes: forming the fan hub (12); and disposing the at least one spacer (30) within the gap (24).

A ceramic matrix composite (CMC) component (200) for a turbine engine (20) includes at least one cooling cavity (260) having a cooling cavity opening (262) and defined by side walls (264) and a cavity bottom wall (266), and a cover plate (270) covering the cooling cavity opening (262). The cover plate has one or more cooling air inlets (275, 285) to permit cooling air to flow from a region above the cover plate into the cooling cavity. The cover plate further has one or more walls (280) extending downward from an inner surface of the cover plate to the cavity bottom wall thereby dividing the cooling cavity into a plurality of cooling channels or cooling subcavities (290, 292, 295), and/or creating a circuitous cooling pathway within the cooling cavity (260).The CMC component is a blade outer air seal (BOAS) segment. A BOAS assembly comprising a plurality of the BOAS segments arranged to form an annular shaped structure.A method of controlling cooling air flow within a CMC component for a turbine engine includes providing the CMC component; and covering the cooling cavity opening of the cooling cavity with the cover plate.The turbine engine includes a fan section (22), a compressor section (24), a combustion chamber (26), and a turbine section (28), the turbine section (28) including at least one rotor and one or more turbine blade(s) extending radially outwardly from the at least one rotor; and the blade outer air seal assembly positioned between the one or more turbine blade(s) and an outer casing to the engine.

A multi-mode blade tip timing sensor (100; 200) includes a housing (201), a sensing coil (205) provided within the housing (201) and a powered exciter coil (207). The multi-mode blade tip timing sensor (100; 200) can operate in a first, sensing mode, and a second, de-icing mode. When operating in the first, sensing mode, a first current is provided to the powered exciter coil (207) to create a magnetic field such that a movement of one or more moving metallic fan blades (109) through the magnetic field creates a signal in the sensing coil (205). When operating in the second, de-icing mode, a second current is provided to the powered exciter coil (207) to induce an eddy current in a tip of the one or more moving metallic turbine fan blades (109), wherein the eddy current raises a temperature of the one or more moving metallic turbine fan blades (109) above an icing temperature.

A ceramic matrix composite (CMC) component (400) for a turbine engine (20) includes a base (410) having a radial outer surface (412) and a radial inner surface (415), a cooling cavity (460) within the base that extends from the radial outer surface (412) of the base into an interior region of the base, the cooling cavity having a cooling cavity opening (462) at the outer radial surface of the base and being defined by cavity side walls ( 464) and a cavity bottom wall (466), and a cover plate (200) covering the cooling cavity opening of the cooling cavity. The cover plate comprises an edge region (210) around the perimeter of the cover plate which is in contact with the base and a central recessed region (220) which extends into the cooling cavity. The central recessed region is defined by side walls (230) and a bottom wall (240) wherein each side wall of the central recessed region is spaced from a corresponding cavity side wall (464) and the bottom wall of the central recessed region is spaced from the cavity bottom wall. The central recessed region has a plurality of cooling air inlets (250) to permit cooling air to flow from a region above the cover plate into the cooling cavity. The CMC component is a blade outer air seal (BOAS) segment. A BOAS assembly comprising a plurality of the BOAS segments arranged to form an annular shaped structure. A method of controlling cooling air flow within a CMC component for a turbine engine includes providing the CMC component; and covering the cooling cavity opening of the cooling cavity with the cover plate. The turbine engine includes a fan section (22), a compressor section (24), a combustion chamber (26), and a turbine section (28), the turbine section (28) including at least one rotor and one or more turbine blade(s) extending radially outwardly from the at least one rotor; and the blade outer air seal assembly positioned between the one or more turbine blade(s) and an outer casing to the engine.

A ceramic matrix composite (CMC) component (200) of a gas turbine engine forming a portion of a substantially ring-shaped gas turbine engine stage includes an inner surface (210) configured for exposure to a hot gas path of the gas turbine engine and an outer surface (220) on a non-gas path side. An inner heat exchange passage (212) is disposed adjacent to the inner surface (210) of the CMC component at a first radial distance relative to an axis of the gas turbine engine. An outer heat exchange passage (222) is disposed adjacent to the outer surface (220) at a second radial distance relative to an axis of the gas turbine engine greater than the first radial distance. A method of reducing thermal stress in the ceramic matrix composite (CMC) component includes feeding a cold air (214) flow to the inner heat exchange passage (212); and feeding a hot air (224) flow to the outer heat exchange passage (222).

A vane arm (22) includes a first end (70) and a second end (72); a shank (74) extending between the first end and the second end and having a first face (78) and a second face (76); a first hole (110) in the shank proximate the first end; a second hole (114) in the shank proximate the second end; and a transverse groove (80) in the first face intersecting the second hole. The transverse groove includes a base (83); a first open end (81) and a second open end (82); and a first side face (84) and a second side face (85). An asymmetry between the transverse groove first side face and second side face provides for uniquely angularly registering the vane arm to a shaft. At least one of the transverse groove first side face and second side face has a first facet (107) at a first angle of 0° to 20° over at least 20% of a depth (DG) of the transverse groove; and a second facet (104) at a second angle of 20° to 45° over at least 50% of the depth of the transverse groove. A vane assembly (20), a vane stage, as well as a method for using the vane arm are also provided.

A vane arm (22) includes a first end (70) and a second end (72); a shank (74) extending between the first end and the second end and having a first face (78) and a second face (76); a first hole (110) in the shank proximate the first end; a second hole (114) in the shank proximate the second end; and a transverse groove (80) in the first face intersecting the second hole. The transverse groove includes a base (83); a first open end (81) and a second open end (82); and a first side face (84) and a second side face (85). The transverse groove first side face and second side face converge toward each other from the first open end to the second open end. A vane assembly (20), a vane stage, as well as a method for using the vane arm are also provided.

A process of fabricating at least one channel in a ceramic-containing material (110) comprising providing at least one ceramic-containing material (110) having a first major surface (120), a second major surface (130) opposite the first major surface (120), at least one first side surface (140a,b), and at least one second side surface (150a,b) adjacent the first side surface (140a,b); removing at least a portion of material from the first major surface (120) or the second major surface (130) of the ceramic-containing material (110) and forming one or more structural features (310a-e); providing at least one metal cover plate (410); disposing the metal cover plate (410) adjacent the ceramic-containing material (110) and covering the structural features (310a-e); and, securing together the ceramic-containing material (110; 610) and the metal cover plate (410).

An aircraft propulsion system (20) assembly includes a plurality of airfoils (110) and an actuation system (106). Each of the airfoils (110) projects radially outward from a respective airfoil base (108) to a respective airfoil tip (118). The airfoils (110) include a first airfoil (110) and a second airfoil (110). The actuation system (106) is configured to pivot the first airfoil (110) about a first airfoil pivot axis (128) from a first airfoil first pitch position (129A), through a first airfoil feather pitch position (129C), to a first airfoil second pitch position (129E). The actuation system (106) is configured to pivot the second airfoil (110) about a second airfoil pivot axis (128) from a second airfoil first pitch position (129A), through a second airfoil feather pitch position (129C), to a second airfoil second pitch position (129E). The actuation system (106) is configured to asynchronously schedule pivoting the first airfoil (110) through the first airfoil feather pitch position (129C) with pivoting the second airfoil (110) through the second airfoil feather pitch position (129C).

A system is provided for an aircraft (20). This aircraft system includes an aircraft propulsion system (24) and a pylon structure (144). The aircraft propulsion system (24) includes a housing structure (90), an open propulsor rotor (66) and a turbine engine (64). The turbine engine (64) is configured to drive rotation of the open propulsor rotor (66). The turbine engine (64) includes an engine centerline and an engine core (114) that is housed within the housing structure (90). The pylon structure (144) is mounted to the aircraft propulsion system (24) and projects longitudinally along a pylon centerline (146) away from the aircraft propulsion system (24). The pylon centerline (146) is circumferentially offset from a reference point (158) about the engine centerline (54) by a first offset angle (160). The reference point (158) is disposed at a top-dead-center location (152) of the housing structure (90). The first offset angle (160) is a non-zero acute angle.

A method of reducing depletion of elements from an airfoil (140) includes disposing a bond coating on the airfoil (140), where the airfoil (140) comprises cooling channels (150). The airfoil (140) includes a material that has an initial composition. A cooling hole (150a) is machined in the airfoil (140) to contact the cooling channels (150) such that a fluid travelling in the cooling channel (150) may be discharged via the cooling hole (150a). The machining of the cooling hole (150a) results in a formation of a depleted region around the cooling hole (150a). The depleted region is depleted of a portion of the initial composition of the airfoil (150). A top coat is disposed on the bond coat. The airfoil (150) is subjected to a heat treatment at a temperature effective to promote diffusion of elements from a non-depleted region to the depleted region around the cooling hole.

A cover plate retention system of a turbine engine (20) includes a gaspath component (300) having a hot side configured for exposure to a hot gaspath of the turbine engine and an opposing non-gaspath side (305). The cover plate retention system includes a cavity (310). The cavity has an opening disposed on the non-gaspath side, and is collectively defined by a pair of parallel side walls (312), a first end wall, and a second end wall (315). A top rail (330) is disposed at a top of the first end wall and extends therefrom along a major portion of the top of each of the parallel side walls (312). A bottom rail (332) is disposed a fixed or specified distance below the top rail (330) on the first end wall and extends therefrom along at least the major portion of each of the parallel side walls (312) to form a three-sided cover plate retention slot.A cover plate retention method for the turbine engine includes forming the cavity in the gaspath component; forming the top rail; and forming the bottom rail.

An aircraft propulsion system (24) is mounted to an aircraft airframe (22). An open propulsor rotor (66) of the propulsion system (24A) is configured to rotate about a rotational axis (54) in a rotational direction within a plane of rotation. The propulsor rotor (66) includes a propulsor blade (74) with a blade tip (80) and a blade span (152). The propulsor blade (74) is configured to rotate about the rotational axis (54) through a blade reference position (148A, 148B) in a reference direction towards a reference point (146A, 146B) on a fuselage (34). The propulsor blade (74) is oriented vertically at the blade reference position (148A, 148B). The blade tip (80) is vertically offset from the reference point (146A, 146B) by an offset distance (150A, 150B) when the propulsor blade (74) is at the blade reference position (148A, 148B). The offset distance (150A, 150B) is equal to or less than fifty percent of the blade span (152). The reference point (146A, 146B) is coincident with the plane of rotation at a vertical periphery of the fuselage (34).

An apparatus is provided for an aircraft propulsion system (20). This apparatus includes an inner housing structure (72) extending axially along and circumferentially about an axis (22). A radial outer side of the inner housing structure (72) forms an inner peripheral boundary (84) of a flowpath (70). The inner housing structure (72) includes an inner case (86), an inner platform (88) of a guide vane structure (80), a flowpath wall (90) and a bulkhead (92). The flowpath wall (90) is axially adjacent the inner platform (88) along the inner peripheral boundary (84) of the flowpath (70). The flowpath wall (90) is spaced radially outboard from the inner case (86). The flowpath wall (90) is configured with a flowpath wall acoustic treatment (136C) extending along the inner peripheral boundary (84) of the flowpath (70). The bulkhead (92) extends radially between and is connected to the inner case (86) and the flowpath wall (90).

An assembly for an aircraft propulsion system (20) includes a first open propulsor rotor (40), a turbine engine (32), a pylon structure (210), a support structure (112), a second open propulsor rotor (42) and a drive system. The turbine engine (32) includes a flowpath, a compressor section (71), a combustor section (72), a turbine section (73) and a rotating structure (94). The rotating structure includes a turbine rotor (84) in the turbine section (73). The rotating structure is configured to drive rotation of the first open propulsor rotor (40). The pylon structure (210) is mounted to the turbine engine (32). The support structure (112) is arranged with the pylon structure (210) independent of the turbine engine (32). The second open propulsor rotor (42) is rotationally supported by the support structure (112). The drive system is configured to drive rotation of the second open propulsor rotor (42).

A vane arm (22) includes a first end (70) and a second end (72); a shank (74) extending between the first end and the second end and having a first face (76) and a second face (78); a first hole (110) in the shank proximate the first end; a second hole (114) in the shank proximate the second end; a first tab (80) projecting laterally from the shank proximate the second end and having a bend (84) and a distal portion (90) distally of the bend; and a second tab (82) projecting laterally from the shank proximate the second end laterally opposite the first tab and having a bend (86), a distal portion (92) distally of the bend and partially overlapping the first tab distal portion. In one vane arm, a third hole (120) in the second tab distal portion overlapping a distal end edge of the first tab and having a proximal edge (122) and a distal edge (124). At least along a region across an axis (115) of the second hole (114), the third hole distal edge and proximal edge diverge from each other in a direction toward the first end. The first tab distal end edge and the third hole proximal edge diverge from each other in a direction toward the first end. In another vane arm, a hole (120) in the second tab distal portion overlapping a distal end edge of the first tab and having a proximal edge (122). At least along a region across an axis (115) of the second hole (114), the first tab distal end edge and the proximal edge of the hole provide means for uniquely angularly registering the vane arm to a shaft (42). Vane assemblies (20), a vane stage, as well as a method for using the vane arm are also provided.

An aircraft propulsion system (20) includes a core engine (25), a propulsive fan, and a bottoming cycle (62) that includes a working fluid mixture (76) within a closed circuit (74) that is heated and expanded through a bottom turbine (78) to generate shaft power (66). The working fluid mixture (76) includes a proportion of at least two fluids (104,106) having different properties that is adjustable to change a capability of the working fluid mixture (76) at different bottoming cycle operating points. The bottoming cycle (62) includes a reactor assembly (96) configured to generate a reaction with one of the at least two fluids (104,106) for adjusting the proportion of the at least two fluids (104,106) of the working fluid mixture (76).

A system for monitoring operation of a gas turbine engine comprises a plurality of expert networks each configured to generate a separate gas path parameter responsive to a separate actuator position. A router network configured to generate a weighting vector responsive to at least one ambient condition parameter. The weighting vector includes a plurality of weighting values each associated with one of the plurality of expert networks. Summing circuitry configured to apply the plurality of weighting values to each of the associated separate gas path parameters from the plurality of expert networks and sum each of the plurality of weighted separate gas path parameters from the plurality of expert networks to a weighted sum value.

A static impact test fixture (100) includes a load plate (104), a broach (102), and an under-root padding (106). The load plate (104) is configured to support weight of a fan blade (122) when a blade root (120) of the fan blade (122) is placed onto the load plate (104). The broach (102) includes a notch (118), a flange portion (115), and a throat opening (114). The notch (118) is configured to fit over the load plate (104) and to allow the broach (102) to slide along a first axis relative to the load plate (104). The flange portion (115) extends in a first direction over the notch (118) and is configured to overlap the load plate (104). The throat opening (114) is configured for a blade neck (124) to pass through. The under-root padding (106) is composed of a compliant material having a specified thickness and compliance that enable the under-root padding (106) to deform to an initial deformation caused by a clamping load between the load plate (104) and the broach (102).

A process of fabricating at least one channel (310a,b,c,d,e) in a ceramic-containing material (110) comprising providing at least one ceramic-containing material (110) having a first major surface (120), a second major surface (130) opposite the first major surface (120), at least one first side surface (140a,b), and at least one second side surface (150a,b) adjacent the at least one first side surface (140a,b); removing at least a portion of material from at least one of the first major surface (120) or the second major surface (130) of the at least one ceramic-containing material (110) and form one or more channels (310a,b,c,d,e) comprising at least one heat augmentation feature (315a,b,c); providing at least one metal cover plate (410); disposing the at least one metal cover plate (410) adjacent the at least one ceramic-containing material (110) and covering the one or more channels (310a,b,c,d,e) and the at least one heat augmentation feature (315a,b,c); and securing together the at least one ceramic-containing material (110) and the at least one metal cover plate (410).

An apparatus for detecting proximity within a gas turbine engine comprising a proximity probe (202) configured to monitor proximity between engine components within areas of the gas turbine engine. An articulated housing (108) contains the proximity probe. The articulated housing (108) enables the proximity probe to monitor the proximity between the engine components in a 360° range about the apparatus.

A feather seal arrangement for a ceramic matrix composite (CMC) blade outer air seal (BOAS) of a turbine engine includes a CMC BOAS segment (300) having a specific geometry and a first feather seal (350). The CMC BOAS segment includes a shoe (310) having a flowpath side (312), an opposing non-flowpath side (314), and first and second matefaces (315) extending axially between the flowpath side and the non-flowpath side from an upstream leading edge (317) to a downstream trailing edge (319). One or more support flanges (320) extend substantially transverse from the non-flowpath side. A geometry of the CMC BOAS segment allows a substantially flat first feather seal surface to be disposed on the non-flowpath side adjacent and substantially transverse to the first mateface. A first feather seal (350) is disposed atop the non-flowpath side on the first feather seal surface and extends to an adjacent feather seal surface of an adjacent CMC BOAS segment.

A probe assembly (100; 300) includes a cup (102; 302) including at least one cooling outlet (104; 304), and a cap (106; 306) including an inlet, the cap (106; 306) affixed to the cup (102; 302). The probe assembly (100; 300) also includes hypo tubing (108; 308) affixed to the inlet, and a probe (110; 310) disposed within an interior volume of the cup (102; 302) and the cap (106; 306), the probe (110; 310) including a lead (112; 312) at least partially disposed within the hypo tubing (108; 308).

A method for machining a thin-walled part includes selecting a tool for machining a part; obtaining tool-based dynamics including tool-based frequency response; developing a preliminary tool path using the tool-based dynamics; simulating a machining process of the part with the preliminary tool path to produce an in-process model; developing a cut stock frequency response from the in-process model; determining stability lobes for machining the part with the tool, the stability lobes being determined from the tool-based frequency response and the cut stock frequency response; choosing operating parameters for a stable machining step from the stability lobes; and machining the part with the tool at the operating parameters.

A cooling flow arrangement (200) for a ceramic matrix composite (CMC) component (210) of a turbine engine is provided, wherein the CMC component includes a hot side (230) configured for exposure to a hot gas path (232) of the turbine engine and an opposing cold side (240). At least one cavity (220) is provided in the CMC component, the at least one cavity having an entrance disposed on the cold side that is configured for receiving a cooling flow. A seal (260) is disposed adjacent the at least one cavity and the cold side of the CMC component and is configured to provide a seal between the CMC component and an upstream turbine component. At least one hole (262) extends through the seal and is positioned to provide the cooling flow to the at least one cavity of the CMC component.A method (300) of providing a cooling flow to the CMC component (210) includes: providing (310) the at least one cavity in the CMC component ; disposing (320) the seal adjacent the at least one cavity and the cold side of the CMC component to provide a seal between the CMC component and an upstream turbine component; and (330) passing the cooling flow from a source (250) through the at least one hole.

A ceramic matrix composite (CMC) component (200) includes a base (210) having a radial outer surface (212) and a radial inner surface inward surface (315). The base includes a plurality of ceramic fiber plies and a ceramic matrix. The radial outer surface has a convex shape, at least one cooling cavity (260a, 260b) within the base that extends from the radial outer surface of the base into an interior region of the base. The at least one cooling cavity has a cavity opening at the radial outer surface of the base. The at least one cooling cavity is defined by cavity side walls (264a, 264b), a cavity bottom wall (266a, 266b), and a cover plate (270a; 270b) positioned within the at least one cooling cavity to cover the cavity opening. The cover plate is angled to vary a depth of the cooling cavity between the cover plate and the cavity bottom wall so that the depth varies from a cavity region of greater depth to a cavity region of lesser depth. A method of controlling cooling air flow within a CMC component includes providing the CMC component. A turbine engine (20) includes a fan section (22), a compressor section (24), a combustion chamber (26), and a turbine section (28). The turbine section includes at least one rotor and one or more turbine blade(s) extending radially outwardly from the at least one rotor, and a blade outer air seal assembly positioned between the one or more turbine blade(s) and an outer casing of the engine. The blade outer air seal assembly is formed of a plurality blade outer air seal segments, wherein each blade outer air seal segment is the CMC component.

An aircraft propulsion system (20), a propulsor section (22) and a core engine section (24) configured to generate an exhaust gas flow (58) that is expanded through a turbine (34) to drive a shaft (36) for driving the fan (26). A propulsor mount assembly (64) is attached to the propulsor section (22) and supports the propulsor section (22) on an aircraft structure and isolates loads generated in the propulsor section (22) from the core engine section (24).

A method of repairing a ceramic vane includes providing a ceramic vane (49) that includes a platform (64) and an airfoil section (66) that extends from the platform. The platform includes a mounting surface (78) for supporting the ceramic vane and the airfoil section includes a damaged region (70). An over-wrap fiber ply (72) is wrapped around the airfoil section and covers the damaged region. The over-wrap fiber ply is then densified with a ceramic matrix material to form a ceramic matrix composite over-wrap repair layer (74). A build-up layer (76) is applied to the at least one mounting surface. A desired stagger angle is provided to the ceramic vane by adjusting a thickness and contour of the at least one build-up layer.

A ceramic matrix composite (CMC) component (200) for a turbine engine (20) includes a base (210) having a radial outer surface ( 212) and a radial inner surface inward surface (215). The base includes a plurality of ceramic fiber plies and a ceramic matrix. A cooling cavity (260) extends from the outer radial surface of the base into an interior region of the base. The cooling cavity has a cavity opening (262) at the outer radial surface of the base. The cooling cavity is defined by cavity side walls (264), a cavity bottom wall (266), and an impingement plate (270) or a cover plate covering the cavity opening. At least one spring member (280) includes a contact region (310), a first curved region (320a) and a first end section (330a). The first curved region is connected to both the contact region and the first end section. The first end section is in contact with a restraining member (240), and the contact region is in contact with the impingement plate or cover plate and acts to hold the impingement plate or cover plate in position to cover the cavity opening. The CMC component is a blade outer air seal (BOAS) segment. A BOAS assembly includes a plurality of the BOAS segments arranged to form an annular shaped structure. A method of assembling the CMC component includes: providing the CMC component; providing the impingement plate or the cover plate; retaining the impingement plate or cover plate in a position to cover the cavity opening using the at least one spring member. The turbine engine includes a fan section (22), a compressor section (24), a combustion chamber (56), and a turbine section (28). The turbine section includes at least one rotor and one or more turbine blade(s) extending radially outwardly from the at least one rotor; the blade outer air seal assembly positioned between the one or more turbine blade(s); and an outer casing (36).

An aerospace piston seal ring including a body shaped as an annular ring, the body defines an axis centered within the body; a sealing surface extending radially relative to the axis between an inner diameter and an outer diameter; and a surface feature formed along the sealing surface, the surface feature configured as an interruption in the sealing surface, wherein the surface feature is configured to contain a lubricant.

A process for restoring a ceramic containing material comprising the steps of cleaning at least a portion of a location (30) on a surface of a ceramic containing material to form a cleaned location (50), the ceramic containing material comprising at least one of the following: a coated preform, a partially densified ceramic matrix composite, a ceramic matrix composite component, and combinations thereof; masking the surface, except for the cleaned location, of the ceramic matrix composite with a maskant (60) to form an aperture (70) exposing the cleaned location; shaping the aperture to form a shaped, exposed cleaned location; depositing a material onto the shaped exposed location to form a location having a deposited material (80); densifying the deposited material to form a densified location; and, reshaping the densified location to form a reshaped densified location.

A method for detecting and removing at least one atmospheric pollutant from a surface of an aircraft engine component (108) includes detecting the at least one atmospheric pollutant on the surface of the aircraft engine component (108) using a test probe (102) and treating the at least one aircraft engine component (108) to generate a layer of at least one of a soot or other carbonaceous substance on the at least one aircraft engine component (108) to at least partially remove the at least one atmospheric pollutant from the surface of the aircraft engine component (108).

A process for fabricating a melt-infiltrated ceramic matrix composite comprises fabricating a ceramic fiber preform comprising at least one fiber, fiber tow, or both; depositing at least one structural support material coating on the fiber, fiber tow, or both to form a partially densified ceramic matrix composite comprising at least one pore, porous network or both; depositing at least one fiber protection material coating on the fiber, fiber tow, or both; infiltrating a slurry composition comprising at least one sacrificial particle into the pore, porous network, or both; depositing at least one reaction control material coating on the fiber, the fiber tow, or the fiber, fiber tow and sacrificial particle of the partially densified ceramic matrix composite; melt-infiltrating at least one molten infiltrant into the partially densified ceramic matrix composite to form a melt-infiltrated ceramic matrix composite.

An attachment pin (240) includes an elongated body with a longitudinal axis extending from a first end (280) to a second end (282), a cooling passageway (290) extending axially from the first end along the longitudinal axis, and one or more cooling air outlets (295) extend radially from the cooling air passageway to provide for discharge of fluid from the cooling passageway. A ceramic matrix composite (CMC) component includes a base (210) having a radial outer surface (212) and a radial inner surface inward surface (215). The base includes a plurality of ceramic fiber plies and a ceramic matrix, a forward flange structure (220) and an aft flange structure (230) each extending from the radial outer surface of the base, the forward flange structure having a first opening (222) and a second opening (225), and the aft flange structure has a first opening (232) and a second opening (235). The CMC component includes a first (240) and a second (250) of the attachment pin for attaching the CMC component to a support structure, the first attachment pin passing through the first opening of the forward flange structure and the first opening of the aft flange structure, and the second attachment pin passing through the second opening of the forward flange structure and the second opening of the second flange structure. A BOAS assembly comprising a plurality of BOAS segments being the CMC component arranged to form an annular shaped structure. A method of cooling a CMC component includes providing the CMC component and the first and the second attachment pins. A turbine engine includes a fan section (22), a compressor section (24), a combustion chamber (56), a turbine section (28), the turbine section including at least one rotor and one or more turbine blade extending radially outwardly from the at least one rotor, and the blade outer air seal assembly being positioned between the one or more turbine blade(s) and an outer casing to the engine.

An apparatus is provided for an aircraft that includes an open rotor propulsion system (20). The open rotor propulsion system (20) includes an open propulsor rotor (34), an open guide vane structure (36) and a turbine engine (32). The open guide vane structure is axially next to the open propulsor rotor. The turbine engine is configured to drive rotation of the open propulsor rotor about an axis (24). An exterior surface (134) of a component (132) of the open rotor propulsion system is exposed to and borders an environment (22) external to the open rotor propulsion system. The component is configured with an acoustic treatment (116) extending axially and circumferentially along the exterior surface (134).