RTX Corporation Patente

A process for melt-infiltrating a partially densified ceramic matrix composite, comprising the steps of: disposing in a chamber of a thermal processing apparatus a partially densified ceramic matrix composite comprising a ceramic matrix including at least silicon carbide and at least one oxide phase precursor; introducing at least one gas into the thermal processing apparatus; elevating a temperature of the chamber to cause a reaction between the silicon carbide and the least one gas; achieving a chemical equilibrium of the gases within the chamber; identifying a gas regime for minimizing the liberation of SiO and SiO<2>; maintaining a temperature and CO<2>:CO ratio of the identified gas regime within the chamber; and, melt-infiltrating the partially densified ceramic matrix composite with the at least one oxide phase precursor to form a molten oxide ceramic matrix composite comprising a matrix including a silicon carbide phase and an oxide phase.

A turbine engine (24) assembly includes an airfoil (102), a first baffle (104) and a second baffle (106). The first baffle (104) is disposed in a cavity (128) in the airfoil (102). The first baffle (104) extends spanwise from a first baffle base end (146) to a first baffle tip end (148) disposed at an airfoil tip end (110). The second baffle (106) is disposed in the cavity (128) longitudinally between and next to the first baffle (104) and an airfoil interior wall (121). The second baffle (106) extends spanwise from a second baffle base end (166) to a second baffle tip end (168) disposed at the airfoil tip end (110).

A turbine engine assembly (86) includes a vane structure (88) and an inner structure (182). The vane structure (88) includes a plurality of structure segments (106) arranged circumferentially about an axis (32). Each of the structure segments (106) includes an inner platform segment (114), an outer platform segment (162) and an airfoil (177). The inner platform segment (114) includes a base (128) and a tab (130). The tab (130) is connected to and projects laterally out from the base (128). The tab (130) is radially next to and extends laterally along the base (128) of the inner platform segment (114) of a respective circumferentially neighboring one of the structure segments (106). The airfoil (177) extends radially across a flowpath (78) and is connected to the inner platform segment (114) and the outer platform segment (162). The inner structure (182) extends circumferentially around the axis (32). The inner structure (182) is configured with a plurality of pins (198). Each of the pins (198) projecting axially into an aperture (146) in the base (128) of a respective one of the structure segments (106).

An assembly is provided for a turbine engine (20). This assembly includes a carrier (80), a control ring (82) and a bladed outer air seal (78) mounted to the carrier (80). The carrier (80) extends circumferentially around an axis (22). The carrier (80) includes an annular cavity (114), a first section (106) and a second section (108). The annular cavity (114) is formed within the carrier (80) by the first section (106) and the second section (108). The first section (106) is welded and/or brazed to the second section (108) at a sealed full-hoop bond joint (148; 150). The first section (106) is further connected to the second section (108) at a second sealed full-hoop joint (148; 150). The control ring (82) extends circumferentially around the axis (22) within the annular cavity (114). A method of manufacture is provided for an assembly for a turbine engine (20).

A seal assembly (60A, 60B; 160B) includes a carbon seal (64A, 64B; 164B) that has a sealing surface (66B, 66B; 166B). A seal seat (62A, 62B; 162B) has a sealing surface (66B, 68B; 168B) and is positioned for rotation relative to the carbon seal (64A, 64B; 164B). A diamond-like carbon coating (80) at least partially forms the sealing surface (68B, 68B; 168B) on the seal seat (62A, 62B; 162B).

A cryogenic fueled aircraft propulsion system includes a core engine (22) where fuel is mixed with compressed air and ignited to generate a high energy exhaust gas flow, a fuel system (24) that is configured to supply the fuel to the core engine (22), a cryogenic fuel tank (26) for storing fuel cryogenically in a liquid phase, a gas collection system (28) where boil off gas (36) from the fuel system (24) is collected, a fuel cell system (34) where a flow of collected boil off gas (36) is utilized to generate electric power (52), and an electrical coupling (48) where electric power (52) generated by the fuel cell is communicated to location outside of the aircraft (20).

A preform (200) for a ceramic matrix composite (CMC) component having a flow path surface, such as a blade outer air seal (BOAS), a blade, a vane, or a combustor panel, includes a base preform (200) includes a polyvinyl butyral (PVB) tackifier for dimensional stability and has a corresponding flow path surface. One or more machined channels (210) are formed on the corresponding flow path surface of the base preform (200), and may include additional features such as ribs, trip strips, or pin fins. One or more overwrap plies are then disposed on the base preform (200) and cover the one or more machined channels (210) to form a cooling feature. The overwrapped base preform may then be densified to provide a CMC component with a cooling feature on the flow path surface.

An aircraft powerplant (20) assembly includes a gearbox (102), a first electric machine (90A), a first electric cable (112A), a second electric machine (90B) and a second electric cable (112B). The first electric machine (90A) is mounted to the gearbox (102) at a first side (135A) of the gearbox (102). A first machine terminal (132A) is disposed at a first side (130A) of the first electric machine (90A). The first electric cable (112A) is electrically coupled to the first machine stator (96A) through the first machine terminal (132A). The second electric machine (90B) is mounted to the gearbox (102) at the first side (135) of the gearbox (102). A second machine terminal (132B) is disposed at a second side (130B) of the second electric machine (90B) that faces the first side (130A) of the first electric machine (90A). The second electric cable (112B) is electrically coupled to the second machine stator (96B) through the second machine terminal (132B).

A method for fabricating an interface coating on a substrate, comprising (i) providing a substrate; (ii) depositing on the substrate at least one layer of boron nitride; and (iii) depositing on the boron nitride layer a Si<3>N<4>-BN multilayer coating to form a coated substrate, the Si<3>N<4>-BN multilayer coating comprises a pattern of alternating layers having at least one boron nitride layer and at least one silicon nitride layer according to the following steps: (iiia) depositing at a first deposition temperature a layer of silicon nitride at a first thickness; (iiib) depositing at a second deposition temperature a layer of boron nitride at a second thickness; and repeating steps (iiia) and (iiib) until forming the Si<3>N<4>-BN multilayer coating. The coated substrate is a ceramic fibrous preform.

An attachment assembly (100) for a ceramic matrix composite (CMC) vane (140) includes an inner diameter ring (110) having a strut recess (118) and a first flange (115) on a downstream side configured for engaging an inner platform (142) of the CMC vane (140), an outer diameter ring (120) having a support structure (125) disposed to form a gap (150) relative to a portion (148) of an outer platform (146) of the CMC vane (140) during a non-operating engine condition, and a strut (130) cantilevered from an inner side of the outer diameter ring (120) and disposed to pass through a hollow portion (148) within an airfoil (145) of the CMC vane (140) to engage the strut recess (118) in the inner diameter ring (110) to support the inner diameter ring (110) and the CMC vane (140). The inner and outer diameter rings and the strut (130) may be metallic and carry the loads rather than the CMC carrying the loads.

A system for indicating rotor to stator clash including a rotor including a forward section and an aft section opposite the forward section; a stator located near the rotor; and a clash indication feature attached to the rotor, the clash indication feature configured to contact the stator responsive to a predetermined deflection dimension of the stator.

An aircraft powerplant (20) assembly includes a gearbox (88) and a power coupler (90). The gearbox (88) includes a first engine power transfer apparatus (120), a second engine power transfer apparatus (122) and an accessory power transfer apparatus (124). The power coupler (90) is mounted to the gearbox (88). During a first operating mode, the power coupler (90) is configured to operatively couple the first engine power transfer apparatus (120) to the accessory power transfer apparatus (124) and operatively decouple the second engine power transfer apparatus (122) from the accessory power transfer apparatus (124). During a second operating mode, the power coupler (90) is configured to operatively couple the second engine power transfer apparatus (122) to the accessory power transfer apparatus (124) and operatively decouple the first engine power transfer apparatus (120) from the accessory power transfer apparatus (122).

A combustor liner (100; 100-1, 100-2; 200-1; 300-1) for a combustor (56) is disposed about an axis (A) in a gas turbine engine (20) having an engine axis (A). The combustor liner (100...300-1) includes a first wall (106; 206; 306) comprising a first opening (110; 210; 310) therethrough, the first opening (110; 210; 310) configured to receive a cooling fluid, a second wall (108; 208; 308) spaced from the first wall (106; 206; 306), the second wall (108; 208; 308) configured to define a portion of a combustion chamber (104) and comprising a second opening (112; 212; 312), the second opening (112; 212; 312) configured to discharge the cooling fluid, a lattice structure (134; 234; 334) disposed between the first wall (106; 206; 306) and the second wall (108; 208; 308), a flow guide (116; 216; 316) defining a portion of the first opening (110; 210; 310) and connecting the first wall (106; 206; 306) to the second wall (108; 208; 308).

An aircraft propulsion system (20) assembly includes an open propulsor rotor (34) and a housing structure (54). The housing structure (54) includes an engine case (106), a nacelle structure (108) and a receptacle (126). The nacelle structure (108) is disposed radially outboard of and partially covers an engine case (106). The nacelle structure (108) forms a first axial section of an exterior flow boundary (56) for the aircraft propulsion system (20) that borders an environment (22) external to the aircraft propulsion system (20). The receptacle (126) is disposed radially outboard of the engine case (106) and axially between the open propulsor rotor (34) and the nacelle structure (108). The receptacle (126) is configured to receive a structure base (50) of an open guide vane structure (36).

An aircraft propulsion system includes aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander (800). The main engine core includes a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the turbo expander (800), prior to being injected into the burner for combustion. The turbo expander (800) includes a rotor (804) separated into a first expander portion (812) and a second expander portion (814) arranged about an output shaft (806) and the output shaft (806) is operably connected to a generator configured to generate electrical power.

A method is described for introducing dormant cooling holes (120) into CMC components that have a base (200) made of a first group of plies (110) and a second group of plies (210). Before or after densification of a CMC preform, dormant cooling holes (120) are provided that pass through the first group of plies (110) but are covered by the second group of plies (210). The dormant cooling holes (120) become open upon erosion of the second plies (210). The opening of the dormant cooling holes (120) allows air to flow through the cooling holes (120) to the area of erosion to slow the erosion process.

A rotor blade (58) for a gas turbine engine (20) is provided that includes an airfoil (68), an attachment section (64), and a neck section (66). The airfoil (68) includes a plurality of internal cooling air passages (90). The attachment section (64) has a base surface (70). A forward center cooling air passage (190AF) is open at the base surface (70) and extends through the attachment and neck sections (64). An aft center cooling air passage (190AA) is open at the base surface (70) and extends through the attachment and neck sections (64). A wall to wall, W2W, cooling air passage (190B) is open at the base surface (70) and extends through the attachment and neck sections (64). The W2W cooling air passage (190B) is disposed between the forward and aft center cooling air passages (190AF). In a transition in a direction from a cross-sectional plane disposed in the attachment section (64) to the airfoil root end (84), a side segment (194, 196) progressively extends outwardly from the W2W cooling air passage.

A system for providing powered device operation on a rotating mechanism comprises static hardware (104) configured to remain in a fixed position and rotating hardware (102) configured to rotate about a central axis. Power generation circuitry associated with the static hardware (104) and the rotating hardware (102) is configured to generate a DC power signal on the rotating hardware (102) and output the DC power signal. A powered device (202) located on the rotating hardware (102) is configured to receive the DC power signal and perform a powered operation on the rotating hardware (102).

A combustor wall (76) includes a panel (88), a boss (86), a wall aperture (90) and multiple cooling apertures (118A, 118B). The panel (88) extends axially along and circumferentially about an axial centerline (22). The panel (88) extends radially between a first panel surface (128) and a second panel surface (130). The first panel surface (128) forms a peripheral boundary of a combustion chamber (58). The boss (86) projects out from the second panel surface (130). The boss (86) extends circumferentially around and forms an outer peripheral boundary of the wall aperture (90). The wall aperture (90) extends along a wall aperture centerline (148) through the combustor wall (76) to the first panel surface (128). The cooling apertures (118A, 118B) are arranged circumferentially about the wall aperture (90). Each of the cooling apertures (118A, 118B) extends along a cooling aperture centerline (180, 192) through the panel (88) and/or the boss (86) to the wall aperture (90). The cooling aperture centerline (180) of a first of the cooling apertures (118A) is angularly offset from the wall aperture centerline (148) by a first acute angle (190).

A system (60) for establishing components of a gas turbine engine (20) may include a constraints engine (62). The constraints engine (62) may be operable to determine a set of constraints (81) based on one or more input parameters (80) and selected values (83) for a group of interrelated design parameters (70) that may be associated with respective elements of an array. The design parameters (70) may correspond to respective components of a gas turbine engine (20). The constraints engine (62) may be operable to communicate the constraints (81) to a design tool (64). The design tool (64) may be operable to select values for the elements of the array from respective ranges of selectable values (82) within a design space (68) of the respective components based on a design model (66) associated with the design parameters (70). A method for establishing components of a system (60) is also disclosed.

A high cycle fatigue test rig including a base; a first preload element attached to the base; a second preload element attached to the base; a test specimen attached to the base; the test specimen having a first end and a second end opposite the first end; the test specimen located in between the first preload element and the second preload element; the first preload element and the second preload element attached to the test specimen proximate the second end; a first clamp in operative communication with the first preload element; a second clamp in operative communication with the second preload element; an end cap in operative communication with the second end of the test specimen, the first preload element and the second preload element; and a dynamic force generator in operative communication with the end cap.

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 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.

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.

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 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).

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 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 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 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).

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).

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).

A fuel delivery system for a combustor (300) of a gas turbine engine (20) has a primary fuel tank (518) configured to store liquid hydrogen (510) and a secondary fuel tank (558) configured to store fuel (550), such as Jet A or SAF. The fuel (510,550) being provided and mixed in the jet engine to alleviate the production of contrails. In addition, the primary fuel tank (518) may be a cryogenic tank (518) to enable the storage of liquid hydrogen, or liquid natural gas. As a result, the primary fuel (510) may be utilized to cool the secondary fuel (550) and/or the gas turbine engine (20).

An aircraft propulsion system (20) includes a gas turbine engine (22), a nacelle (24), and a controller (78). The gas turbine engine (22) includes a rotational assembly rotatable about an axis (44). The rotational assembly includes a bladed fan rotor (58). The bladed fan rotor (58) includes a plurality of fan blades (62). The nacelle (24) houses the gas turbine engine (22). The nacelle (24) includes an air inlet (68). The air inlet (68) forms an inlet flow region (76). The inlet flow region (76) includes a reference section (98) at and upstream of the plurality of fan blades (62). The reference section (98) includes a cross-sectional flow area of the inlet flow region (76). The controller (78) is configured to control the gas turbine engine (22) to operate at or above a target engine speed corresponding to corrected air mass flow rate per unit area through the reference section (98) which is greater than a vortex breakdown threshold for the reference section (98).

A gas turbine engine (22) includes a bladed rotor (66), a case (68), and an outer air seal (OAS) assembly (70). The bladed rotor (66) includes a plurality of rotor blades (80) each extending to a blade tip (82). The case (68) circumscribes the bladed rotor (66). The case (68) forms an OAS cavity (90) at the rotor stage (76). The OAS assembly (70) includes a plurality of OAS segments (86) and an air seal (88). The plurality of OAS segments (86) are disposed within the OAS cavity (90). Each of the OAS segments (86) is radially moveable within the OAS cavity (90) between and to an inboard radial position and an outboard radial position. The air seal (88) is positioned radially between and contacting the case (68) and the seal body (122). The air seal (88) biases each of the plurality of OAS segments (86) towards the inboard radial position.

During a manufacture method, a component (60) is scanned using a CT machine (126). The component (60) includes a first member (74) and a second member (76). The first member (74) includes a first member aperture (116) extending through the first member (74) to the second member (76). Aperture data is determined for the first member aperture (116) based on scan data. A centerline vector (132) is determined for the first member aperture (116). The determining of the centerline vector (132) includes fitting a geometric primitive (134) to the aperture data to describe at least a portion of the scanned outer perimeter geometry of the first member aperture (116). The centerline vector (132) for the first member aperture (116) is determined based on a centerline vector (132) of the geometric primitive (134) and/or a center point of the geometric primitive (134). A point of intersection (136) between the centerline vector (132) and a reference plane (138) is determined. A second member aperture (118) is formed in the second member (76) according to a formation operation.

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 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 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.