RTX Corporation Patente

An aircraft propulsion system (20) includes a core engine (24) that includes a core flow path (C) through a main compressor (26) where an inlet airflow (58) is compressed and communicated to a combustor (28) to generate an exhaust gas flow (56) that is expanded through a main turbine section (30) to generate mechanical power to drive the main compressor (26) and a propulsive fan (22), a cryogenic fuel system (36) that includes a cryogenic fuel storage tank (38), a fuel flow path (62) for routing a cryogenic fuel flow (42) to the combustor (28) of the core engine (24), and a thermoelectric generator (46) that is in thermal communication with the cryogenic fuel flow (42) and a heat source (68) to generate a temperature differential (72) across the thermoelectric generator (46) utilized to generate electric power (48).

An assembly is provided for an aircraft powerplant. This assembly includes a gearbox (98), an engine accessory, an engine core (46), a motor accessory and an electric motor (100). The gearbox (98) includes an engine power transfer system (110, 111) and a motor power transfer system (112) independent of the engine power transfer system (110, 111). The engine accessory includes an engine accessory rotor. The engine core (46) includes a flowpath (82), a compressor section (41), a combustor section (42), a turbine section (43) and an engine rotating assembly. The flowpath (82) extends through the compressor section (41), the combustor section (42) and the turbine section (43). The engine rotating assembly includes a turbine rotor in the turbine section (43). The engine rotating assembly is operatively coupled to the engine accessory rotor through the engine power transfer system (110, 111). The motor accessory includes a motor accessory rotor (142). The electric motor (100) includes an electric motor rotor (140) operatively coupled to the motor accessory rotor (142) through the motor power transfer system (112).

A gas turbine engine (22) for an aircraft propulsion system includes a turbine section (32), an exhaust section (34), an engine case (60), and a thermophotovoltaic (TPV) assembly (70). The turbine section (32) and the exhaust section (34) form a core flow path (66) along an axis (44) of the gas turbine engine (22). The engine case (60) is disposed within the exhaust section (34). The engine case (60) includes an inner liner (72) and an outer liner (74). Each of the inner liner (72) and the outer liner (74) extend circumferentially about the axis (44). The inner liner (72) forms the core flow path (66) through the exhaust section (34). The outer liner (74) is disposed radially outward of the inner liner (72) and forms an annular cavity (84) between the inner liner (72) and the outer liner (74). The TPV assembly (70) includes a plurality of TPV cells (86). The plurality of TPV cells (86) are disposed within the annular cavity (84).

A ceramic matrix composite (CMC) blade outer air seal (BOAS) assembly includes a CMC BOAS (200) including a base (210) and a pair of rails (220) having a substantially π-shaped cross-section. The pair of rails include a first pair of aligned holes (235) at a first end and a second pair of aligned holes at a second end. A first pair of parallel metallic pins (230) have a first end connected to a first central support pin (260) and a second end connected to a second central support pin, respectively, wherein the first pair of parallel metallic pins are configured to engage the first pair of aligned holes. A second pair of parallel metallic pins also have a first end connected to a first central support pin and a second end connected to a second central support pin, respectively, wherein the second pair of parallel metallic pins are configured to engage the second pair of aligned holes.A dual pin loaded BOAS includes a CMC BOAS base having a first end and a second end. A pair of CMC BOAS rails extend transversely from the CMC BOAS base. The CMC BOAS rails are substantially parallel and extend from the first end to the second end. Each CMC BOAS rail has a first pair of holes adjacent the first end and a second pair of holes adjacent the second end. A first metallic rotatable support engages each of the first pair of holes. A second metallic rotatable support engaging each of the second pair of holes. Each of the first and second rotatable metallic supports includes:• a first central support pin and a second central support pin disposed on a first axis, wherein the first and second central support pins are configured to rotate about the first axis when engaged within respective holes (445) on a support (440);• a pair of parallel pins having a first end connected to the first central support pin and a second end connected to the second central support pin, respectively, wherein the pair of parallel pins are configured to engage a respective pair of holes in each CMC BOAS rail.A method of supporting a CMC BOAS is also provided.

A method of manufacture is provided during which a component (60) is scanned using a computed tomography machine (126) to provide scan data. The component (60) includes a first member (74) and a second member (76). A first member aperture (116) extends through the first member (74) to the second member (76). Aperture data is determined for the first member aperture (116) based on the scan data. A center (136) of the scanned geometry is located. An eigen vector (138) is determined for the scanned geometry at the center (136). An exterior point of intersection (142) between an exterior surface (70) of the second member (76) and an axis (140) is located. The axis (140) is coincident with the center (136) and determined using the eigen vector (138). A second member aperture (118) is formed in the second member (76) according to a formation operation that aligns the second member aperture (118) being formed with the first member aperture (116) in the first member (74).

A gas turbine engine case structure (150) includes the unitary combination of: a case wall having an inner surface and an outer surface; at least one fuel inlet (171); circumferentially-distributed fuel injectors protruding inward from the case wall and having an outlet; a fuel plenum fluidically between the at least one fuel inlet (171) and the fuel injectors and configured so that each inlet (171) of the at least one inlet (171) is coupled to feed multiple of the fuel injectors; and a fuel inlet conduit (260) fluidically between the fuel inlet (171) and the fuel plenum. The inlet conduit (260) has a forward inlet and is held spaced from the case wall at multiple longitudinally-spaced locations.

A method of repairing a stator stage (200) is disclosed herein. The method includes receiving a stator stage (200) including a plurality of stator vanes (406) disposed between an outer diameter (402) and an inner diameter (404), analyzing the stator stage (200) for defects, determining based on the analysis that there is a first defect (412a) on an edge of a first stator vane (406a) of the plurality of stator vanes (406), removing a portion of the first stator vane (406a; 506) including the first defect (412a) to form a first scallop (422a) on the edge of the first stator vane (406a , repairing the first stator vane (406a) including filling the first scallop (422a) to fill the first stator vane (406a) to its original size and shape creating a repaired portion (432a), and performing a blending process to the stator stage (200) including the first stator vane (406a) and the repaired portion (432a) to smooth the plurality of stator vanes (406).

An assembly for an aircraft powerplant (20) includes an engine case (72), a gearbox (102), a conduit (122), a first electric machine (92A), a first controller (94A) and a first electric cable (114A). The gearbox (102) is offset from the engine case (72). The conduit (122) extends longitudinally from a first end (124) of the conduit (122) to a second end (126) of the conduit (122). The conduit (122) is anchored to the engine case (72) at the first end (124) of the conduit (122) by a first conduit mount (136). The conduit (122) is anchored to the gearbox (102) at the second end (126) of the conduit (122) by a second conduit mount (138). The conduit (122) is self-supported longitudinally between the first conduit mount (136) and the second conduit mount (138). The first electric machine (92A) is mounted to the gearbox (102). The first controller (94A) is mounted to the engine case (72). The first electric cable (114A) electrically couples the first electric machine (92A) to the first controller (94A). The first electric cable (114A) extends longitudinally through a bore (134) of the conduit (122).

A method for processing a CMC airfoil includes nesting an airfoil fiber preform (10) in a cavity (20) of a fixture (12) that has first and second tool segments (16/18), closing the fixture (12) by rotating a first tool segment (16) about a hinge (22), the closing causing the tool segments (16) to clamp on a tail portion (14) of the fiber preform (10) and thereby conform the tail portion (14) to the fixture (12). While in the fixture (12), the fiber preform (10) is then partially densified with an interface coating material to form a partially densified fiber preform. While still in the fixture (12), one or more cooling holes are drilled into the trailing edge (10b) of the partially densified fiber preform. After the drilling, the partially densified fiber preform is removed from the fixture (12) and further densified with a ceramic matrix material to form a fully densified CMC airfoil.

A component (60) is scanned with one or more artifacts (132) using a computed tomography machine (126). The component includes a first member (74) and a second member (76) covering the first member. The first member includes a first member aperture (116) extending through the first member to the second member. Scan data is scaled using known dimensional data for the one or more artifacts to provide scaled scan data. Aperture data is determined for the first member aperture based on the scaled scan data. The aperture data is indicative of a scanned geometry of the first member aperture. A first drilling reference is determined based on the aperture data. A second member aperture is formed in the second member according to a formation operation that aligns the second member aperture being formed with the first member aperture in the first member.

An assembly for an aircraft propulsion system includes an open propulsor rotor (46) and an open guide vane structure (48). The open propulsor rotor (46) is configured to rotate about a rotational axis (24). The open guide vane structure (48) is axially next to the open propulsor rotor (46). The open guide vane structure (48) includes a plurality of open guide vanes (108) arranged circumferentially about the rotational axis (24). The open guide vanes (108) include a first open guide vane (108). The first open guide vane (108) includes a first vane leading edge (122), a first vane trailing edge (124) and a fixed first vane chord length (126) measured from the first vane leading edge (122) to the first vane trailing edge (124). The first open guide vane (108) is configured to translate axially along the rotational axis (24).

A method comprises: flowing a potted component in a liquid state over a tip (214) of an airfoil (210), the tip (214) of the airfoil (210) having a coating (220) disposed thereon, the coating (220) comprising a metal plating (221) and a plurality of protrusions (222), each protrusion in the plurality of protrusions (222) extending from the metal plating (221); allowing the potted component to harden to form a hardened potted component; and removing the hardened potted component from the tip (214) of the airfoil (210).

An apparatus includes a blade outer air seal (BOAS) (100; 400), a measurement instrument adapter (200; 300), and a measurement instrument (130; 430). The BOAS (100; 400) includes at least one first channel (202, 204, 206; 302) configured to permit air to flow from an entry point (416A) of the at least one first channel (202, 204, 206; 302) to an exit point (416B) of the at least one first channel (202, 204, 206; 302). The measurement instrument adapter (200; 300) includes at least one second channel (110, 116; 404, 416) configured to permit air to flow from an entry point (110A) of the at least one second channel (110, 116; 404, 416) to an exit point (110B) of the at least one second channel (110, 116; 404, 416). The measurement instrument adapter (200; 300) is disposed within a first receptacle (114; 208; 414) of the measurement instrument adapter (200; 300) configured to receive the measurement instrument (130; 430). The measurement instrument adapter (200; 300) is disposed within a second receptacle (114; 414) of the BOAS (100; 400) configured to receive the measurement instrument adapter (200; 300). The apparatus is configured for installation in a gas-turbine engine.

A method of operation is provided during which hydrogen fuel is produced using a reversible fuel cell system (32) onboard a vehicle while the vehicle is stationary and/or docked. The reversible fuel cell system (32) receives water and input electricity to produce the hydrogen fuel. The hydrogen fuel is stored onboard the vehicle. Output electricity is generated using the reversible fuel cell system (32) while the vehicle is moving. The reversible fuel cell system (32) receives the hydrogen fuel stored onboard the vehicle and air to generate the output electricity.

A method of making a metallic component (114) includes mounting a polymeric scaffold (102) onto a mandrel (108), wherein the mandrel (108) includes one or more chucks (108a) configured to secure the polymeric scaffold (102) onto the mandrel (108) and wherein the mandrel (108) is configured to rotate the polymeric scaffold (102); depositing, with a second additive manufacturing technique using a material applicator (110), a component material (112) onto the polymeric scaffold (102) to form the metallic component (114) on the polymeric scaffold (102), wherein the component (114) has a shape that is the same as the polymeric scaffold (102); removing the polymeric scaffold (102); and the metallic component (114) deposited on the polymeric scaffold (102) from the mandrel (108); and removing the polymeric scaffold (102) from the metallic component (114).

A method of assembling a tube assembly (100) includes the steps of providing a saddle clamp (104) having a first clamp half (109) extending for less than 50 percent of a circumference of a tube (102), and a second clamp half (108) extending for more than 50 percent of the circumference of the tube (102), placing a tube (102) on the first clamp half (109), resting on a static structure (106), and applying a lubricant (122) to a laterally inner surface (117) of the second clamp half (108) and then bringing the second clamp half (108) onto the tube (102), and then driving a fastener (110) to secure the first and second clamp halves (108, 109). A gas turbine engine (20) is also disclosed.

A fan exit guide vane with acoustic treatment includes a leading edge and a trailing edge opposite chordwise from the leading edge; a radially inner attachment region opposite spanwise from a radially outer attachment region; a span dimension extending between the radially inner attachment region and the radially outer attachment region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the fan exit guide vane; an acoustic panel receiver formed within the fan exit guide vane extending at least one of spanwise through the fan exit guide vane between the radially inner attachment region and the radially outer attachment region or chordwise between the leading edge and the trailing edge; and an acoustic panel inserted into the acoustic panel receiver. A fan exit guide vane system with acoustic treatment for a gas turbine engine includes fan exit guide vane with acoustic treatment. A gas turbine engine includes the fan exit guide vane system with acoustic treatment for a gas turbine engine. A process forms the fan exit guide vane with acoustic treatment.

A swirl recovery vane with structurally capable acoustic treatment includes a leading edge and a trailing edge opposite chordwise from the leading edge; an attachment region opposite spanwise from a tip region; a span dimension extending between the attachment region and the tip region; a chord dimension extending between the leading edge and trailing edge; a pressure side opposite a suction side of the swirl recovery vane; a structural feature formed within the swirl recovery vane extending at least one of spanwise through the swirl recovery vane between the attachment region and tip region and/or chordwise between the leading edge and the trailing edge; an acoustic receiver formed within the structural feature extending at least one of spanwise through the swirl recovery vane between the attachment region and the tip region or chordwise between the leading edge and the trailing edge; and an acoustic panel inserted into the acoustic receiver. A gas turbine engine includes the swirl recovery vane with structurally capable acoustic treatment. A process forms the gas turbine engine with the swirl recovery vane with structurally capable acoustic treatment.

A component retaining assembly (65) includes a housing (62) including a circumferential slot (76). A component (64) is mated to the housing (62). The component (64) includes a first face (86) including at least two tabs (78). The at least two tabs (78) extending at outward from the first face (86) at least partially past a portion of the circumferential slot (76). A retaining ring (66) is disposed within the circumferential slot (76) and abuts the first face (86) of the component (64). The at least two tabs (78) overlap a portion of the retaining ring (66). A gas turbine engine (20) and a method are also disclosed.

A manufacturing method is provided. During this method, a preform component (60') for a turbine engine is provided that includes a substrate (74). A meter section (102) of a cooling aperture (64) is formed in the substrate (74'). An external coating (76',78') is applied over the substrate. At least a portion of the substrate (74') and the external coating (76',78') is scanned with an imaging system (156) to provide scan data indicative of an internal structure of the portion of the substrate (74') and the external coating (76',78'). A diffuser section of the cooling aperture is formed in the external coating (76',78') and the substrate (74') based on the scan data.

A gas turbine engine (20) has: a compressor (24); a combustor (56); a turbine (28); a gaspath (C) through the compressor (24), combustor (56) and turbine (28); a bleed flowpath (920) from the gaspath (C); an air-air heat exchanger (200) having a heat donor leg along the bleed flowpath (920) ; and a Ranque-Hilsch vortex tube (220) having an inlet (222) along the bleed flowpath (920) and having a cool air outlet (228) and hot air outlet (226). The cool air outlet (228) is connected to a cold air load.

A swirl recovery vane with acoustic treatment including a leading edge and a trailing edge opposite chordwise from the leading edge; an attachment region opposite spanwise from a tip region; a span dimension extending between the attachment region and the tip region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the swirl recovery vane; an acoustic zone integrally formed within the swirl recovery vane. A gas turbine engine includes the swirl recovery vane with acoustic treatment. A process forms the swirl recovery vane with acoustic treatment.

A gas turbine engine includes a a fan case (108) surrounding a fan rotor, and a core engine (232) including a low pressure compressor. A rigid connection between the fan case (108) and the core engine (232) includes three triangular-frame connecting members (116) rigidly connected to the fan case (108) at a fan case connection point, and to the core engine (232) at a core engine connection point. The triangular-frame connecting members (116) each are defined by two rigid legs (118, 120) which extend between the fan case (108) and to the core engine (232), along directions each have a component extending radially inwardly and a component in opposed circumferential directions to each other. A plurality of non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise and are rigidly mounted to at the fan case (108) and/or the core engine (232).

A swirl recovery vane with acoustic treatment includes a leading edge and a trailing edge opposite chordwise from the leading edge; an attachment region opposite spanwise from a tip region; a span dimension extending between the attachment region and the tip region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the swirl recovery vane; an acoustic panel receiver formed within the swirl recovery vane extending at least one of spanwise through the swirl recovery vane between the attachment region and the tip region or chordwise between the leading edge and the trailing edge; and an acoustic panel inserted into the acoustic panel receiver. A gas turbine engine includes the swirl recovery vane with acoustic treatment. A process forms the swirl recovery vane with acoustic treatment.

A fan exit guide vane with structurally capable acoustic treatment includes a leading edge and a trailing edge opposite chordwise from the leading edge; a radially inner attachment region opposite spanwise from a radially outer attachment region; a span dimension extending between the radially inner attachment region and the radially outer attachment region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the fan exit guide vane; a structural feature and an acoustic receiver formed within the fan exit guide vane extending at least one of spanwise through the fan exit guide vane between the radially inner attachment region and the radially outer attachment region and/or chordwise between the leading edge and the trailing edge; and an acoustic panel inserted into the acoustic receiver. A gas turbine engine includes the fan exit guide vane with structurally capable acoustic treatment. A process forms the fan exit guide vane with structurally capable acoustic treatment.

A heat exchanger plate 44 for provides heat transfer between a first flow 910 along a first flowpath 900 and a second flow 912 along a second flowpath 902. The heat exchanger plate 44 comprised a body 52 having: a first face 62 and a second face 64 opposite the first face 62; a leading edge 54 along the second flowpath 902 and a trailing edge 56 along the second flowpath 902; a proximal edge 60 having at least one inlet port 22 along the first flowpath 900 and at least one outlet port 24 along the first flowpath 900; and at least one passageway along the first flowpath 900. Along a proximal portion 302, the first face 62 and the second face 64 converge at a first angle. Along a distal portion 304, the first face 62 and the second face 64 converge at a second angle less than the first angle.

A gas turbine engine (20) includes a fan (42) delivering air into a bypass duct (13) defined between a nacelle (301) and an inner core housing (15). The inner core housing (15) receives a compressor section (24), a turbine section (28) and a combustor (56). The nacelle (301) has an inner periphery (141) with a forwardmost point, and receives an acoustic structure on the inner periphery (141) adjacent the forwardmost point. The acoustic structure is defined by a three dimensional array (100) of interconnected resonators (102), with the interconnected resonators (102) extending in a radial direction (R), a circumferential direction (C) and an axial direction (A) all defined about a rotational axis (A) of the engine (20), with the interconnected resonators (102) having a larger cross-sectional area central body (104), and six members (106, 108, 110, 112, 143; 198) connecting the central body (104) of the resonators (102) to adjacent resonators (102) at respective central bodies (104). A perforated face sheet (118) is inward of the three dimensional array (100) of interconnected resonators and an anti-icing system.

A fan exit guide vane with acoustic treatment includes a leading edge and a trailing edge opposite chordwise from the leading edge; a radially inner attachment region opposite spanwise from a radially outer attachment region; a span dimension extending between the radially inner attachment region and the radially outer attachment region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the fan exit guide vane; and an acoustic zone integrally formed within the fan exit guide vane.A gas turbine engine includes the fan exit guide vane with acoustic treatment. A process forms the fan exit guide vane with acoustic treatment.

A three-dimensional (3-D) composite structure includes a 3-D lattice structure (32) having a plurality of electrically insulative struts, a matrix phase (38) surrounding the 3-D lattice structure (32), first and second electrically conductive face sheets (34, 35) positioned on two faces of the 3-D lattice structure (32), and a plurality of electrically insulative containment sheets positioned on all faces of the 3-D lattice structure (32) that do not include the first and second face sheets (34, 35). The matrix phase (38) includes an electrorheological material. The first and second face sheets (34, 35) are positioned such that an electric potential applied between the first and second face sheets (34, 35) creates an electric field in the matrix phase (38) that causes a desired reversible alteration to the viscosity of the matrix phase (38). The first and second face sheets (34, 35) and the plurality of containment sheets are collectively configured to contain the matrix phase (38) within the 3-D lattice structure (32).

An assembly for an aircraft propulsion system (20) includes a propulsor rotor (42) and a guide vane structure (88). The guide vane structure (88) includes a plurality of guide vanes (90). The guide vane structure (88) is disposed axially next to the propulsor rotor (42). The guide vanes (90) include a first guide vane. The first guide vane includes a camber line (98), a leading edge (100), a trailing edge (102), an inner end (94), an outer end (96), at least one leading edge section (106) and a trailing edge section (108). The first guide vane extends longitudinally along the camber line (98) from the leading edge (100) to the trailing edge (102). The leading edge section (106) forms a respective portion of the leading edge (100) and is arranged radially along the trailing edge section (108) between the inner end (94) and the outer end (96). The trailing edge section (108) forms the trailing edge (102). The leading edge section (106) is configured to translate longitudinally along the camber line (98).

An assembly for an aircraft includes a first propulsion system (24B) and a second propulsion system (24A). A first open propulsor rotor (60) is configured to rotate in a first rotational direction about a first axis (46). A first engine core (102) includes a first rotating structure (116) with a first turbine rotor (107). The first rotating structure (116) is configured to rotate in the first rotational direction about the first axis (46) and drive rotation of the first open propulsor rotor (60) through a first geartrain (58). A second open propulsor rotor is configured to rotate in a second rotational direction about a second axis that is opposite the first rotational direction. A second engine core includes a second rotating structure with a second turbine rotor. A second rotating structure is configured to rotate in the first rotational direction about the second axis and drive rotation of the second open propulsor rotor through a second geartrain.

A space filler (210; 230; 240; 260) for forming a fibrous preform (200) may comprise an additively manufactured ceramic material. The additively manufactured ceramic material may define a plurality of pores (214; 234; 244; 264). A shape of the additively manufactured ceramic material may complement a shape of a void (204) formed by fibrous regions (202) of the fibrous preform (200).

A propulsion system (20) for an aircraft includes an open propulsor rotor (38), a turbine engine (32) and a thermal management system (36). The turbine engine (32) includes an engine core (82) and an engine flowpath (76). The engine core (82) includes a compressor section (71), a combustor section (72) and a turbine section (73). The engine flowpath (76) extends longitudinally though the compressor section (71), the combustor section (72) and the turbine section (73) from an airflow inlet (78) into the engine flowpath (76) to a combustion products exhaust (80) from the engine flowpath (76). The thermal management system (36) is configured to manage heat energy generated by the propulsion system (20) during operation of the turbine engine (32). The thermal management system (36) includes an electric boost compressor (114), a heat exchanger (116) and a system flowpath (110) disposed outside of the engine core (82). The system flowpath (110) extends longitudinally through the electric boost compressor (114) and the heat exchanger (116) from an airflow inlet (122) into the system flowpath (110) to an airflow exhaust (124) from the system flowpath (110).

The incorporation of non-structural ceramic plies to a CMC component (10) or preform is used to provide one or more passages (25) allows for cooling air (45) to pass between segregated cooling cavities. In various embodiments, this may include preforming additional layers of non-structural plies (20) to allow for machining of one or more metered passages (25), or the incorporation of a pre-formed passage or passages (25) into non-structural ply layups. The one or more passages (25) may be machined and/or preformed and may include cylindrical, racetrack, slot, conical or similar geometries. The one or more passage (25) may be used to pass cooling air (45) between cavities separated by sealing features (W-seal, feather seal, etc.). Bypassing discrete sealing features is beneficial to control and/or meter cooling air flow (45) between high-pressure and lower-pressure cavities.

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An abradable coating/thermal barrier coating (235) suitable for use with jet engine CMC components is described which comprises a material selected from hafnon, mixtures of hafnon and zircon, and rare earth disilicates (RE₂Si₂O₇), wherein RE is Sc, Y, La, Ce, Pr, Nd, Pm, Sm, Eu, Gd, Tb, Dy, Ho, Er, Tm, Yb, or Lu. The coating (235) has a porosity gradient wherein the porosity decreases in a radial direction. The porosity gradient provides a progressively increasing wear resistance to slow the rate of rub interaction. The porosity gradient also reduces the thermal gradient through the coating (235) thereby reducing the formation of thermal stresses within the coating (235) and any underlying CMC component (200).

A method of forming a multilayer environmental barrier coating, comprising: providing a ceramic matrix composite; depositing a bond coat layer on the ceramic matrix composite; depositing a top coat layer on the bond coat layer; depositing an over coat layer on the top coat layer, wherein the over coat layer includes no more than approximately 50 percent by volume of porosity.

A plastic encased mandrel includes an open cell carbonaceous foam that is shaped to correspond to a core of a ceramic matrix composite (CMC) part. A polyvinyl butyral (PVB) plastic is injection molded into the carbonaceous foam, before or after shaping, to encase the carbonaceous foam core and form an encased mandrel. The open cell carbonaceous foam may be silicon carbide (SiC) foam or reticulated vitreous carbon (RVC) foam. The foam core may be incorporated into the CMC part when SiC foam is used for better distribution of thermal and structural loads or the foam core may be removed from the CMC part when RVC foam is used, such as when a CMC vane incorporates a metal spar.

A propulsion system (20) for an aircraft includes a propulsor rotor (38), a powerplant (24), a nacelle (36) and a lockout system (100). The powerplant (24) is coupled to and configured to drive rotation of the propulsor rotor (38). An electrical system (86) for the powerplant (24) is configured to receive electrical power from an electrical power source (90). The nacelle (36) includes an access structure (92A) configured to move between a first position and a second position. The access structure (92A) at least partially covers the electrical system (86) when in the first position. The access structure (92A) is arranged in the first position for propulsion system operation. The access structure (92A) is operable to be arranged in the second position for propulsion system maintenance and/or propulsion system inspection. The lockout system (100) is configured to automatically trigger an operation to cut off the electrical power to the electrical system (86) from the electrical power source (90) when the access structure (92A) moves away from the first position.

The disclosure describes a ceramic matrix composite (CMC) component having a cooling cavity (130). The CMC component has a base (105) having a bottom inner surface (120) and a top outer surface (110), in which the base (105) is made from a plurality of ceramic fiber plies within a matrix. A cooling cavity (130) is provided within the base (105) in which the cooling cavity (130; 235) has an opening in the top outer surface (110) of the base (105). A cover plate (140) made from one or more ceramic fiber plies within a matrix is provided to close the opening of the cooling cavity (130). The cover plate (140) is bonded to the base (105) by simultaneous densification of the base (105) and cooling cavity (130).