RTX Corporation Patente

An aircraft powerplant assembly includes a gearbox (102) and a power coupler (104). The gearbox (102) includes a first power transfer apparatus (128) and a second power transfer apparatus (130). The first power transfer apparatus (128) is configured to couple and transfer first mechanical power through the gearbox (102) between a first engine apparatus (142) and a first engine accessory (106). The second power transfer apparatus (130) is configured to couple and transfer second mechanical power through the gearbox (102) between a second engine apparatus (152) and a second engine accessory (108) during a first mode. The power coupler (104) includes a variable speed transmission (160) and a clutch (162). The variable speed transmission (160) is coupled between the clutch (162) and the second power transfer apparatus (130). The clutch (162) is configured to couple the variable speed transmission (160) to the second engine apparatus (152) during the first mode. The clutch (162) is configured to decouple the variable speed transmission (160) from the second engine apparatus (152) during a second mode.

An apparatus is provided that includes a first electric device (128), a second electric device (130), an electric cable (122) and a cable shield (124). The electric cable (122) electrically couples the first electric device (128) to the second electric device (130). The cable shield (124) includes a first metal and a second metal. The first metal has a first permeability and a first impedance. The second metal has a second permeability and a second impedance. The first permeability is greater than the second permeability. The first impedance is greater than the second impedance. The electric cable (122) projects longitudinally, through a bore (126) of the cable shield (124) and along the first metal and the second metal, between the first electric device (128) and the second electric device (130).

An aircraft powerplant system (22) includes a turbine engine (24) and an accessory system. The turbine engine (24) includes a core flowpath (52), a first rotating structure (72), a compressor section (45), a combustor section (46) and a turbine section (47). The first rotating structure (72) includes a first bladed rotor (65, 66, 67, 68) arranged along the core flowpath (52) in the compressor section (45) or the turbine section (47). The accessory system includes a first electromechanical apparatus (172), an air circuit (136) and an air device (142). The first electromechanical apparatus (172) includes a first electric machine (102) and a first boost compressor (138). The first electric machine (102) includes a first electric machine rotor (106) that is operatively coupled to the first rotating structure (72). The first boost compressor (138) includes a first boost compressor rotor (166) that is operatively coupled to the first electric machine rotor (106) and the first rotating structure (72). The air circuit (136) is configured to direct air received from an air source (144), through the first boost compressor (138), to the air device (142).

A gas turbine includes a propulsor for delivering air as propulsion air, and for delivering air into a core engine. The core engine includes a compressor section, a combustor and a turbine section. A turbine rotor in the turbine section drives the propulsor, and a compressor rotor in the compressor section. There is an epicyclic gear reduction (48) between the turbine rotor and the propulsor. The epicyclic gear reduction (48) includes a sun gear (102) driven by the turbine rotor to engage and rotate intermediate gears (106). The intermediate gears (106) are mounted in a carrier. The intermediate gears (106) also engage a ring gear (100). At least one of the carrier and ring gear (100) drive the propulsor, and the intermediate gears (106) include a rim and gear teeth (109) extending outwardly of the rim. The intermediate gears (106) are mounted on journal bearings (108) which are supported in the carrier. The intermediate gear rims are provided with cooling passages extending with at least a component in a direction parallel to an axis of rotation of the intermediate gears (106).

An aircraft assembly includes an engine core (46), a core case (76), a nacelle structure (78), a core compartment (80), a first electric machine (104A) and a first driveshaft (130A). The engine core (46) includes a first rotating structure (68A) rotatable about a first axis (34) and that includes a first structure rotor (61; 64). The core case (76) is radially outboard of and houses the engine core (46). The nacelle structure (78) is radially outboard of and covers the core case (76). The core compartment (80) is disposed radially between and is formed by the core case (76) and the nacelle structure (78). The first electric machine (104A) is disposed in the core compartment (80). The first electric machine (104A) includes a first machine rotor (108A) and a first machine case (116). The first machine case (116) houses the first machine rotor (108A). The first machine case (116) is connected to and projects out from a section (120) of the core case (76). The first drivetrain (130A) operatively couples the first machine rotor (108A) to the first rotating structure (68A).

An assembly for an aircraft propulsion system (20) includes an open propulsor rotor (34), a housing structure (54), and a turbine engine core (76). The open propulsor rotor (34) includes a plurality of open propulsor blades (42) arranged circumferentially around an axis (24). Each of the plurality of open propulsor blades (42) projects radially out into an environment (22) external to the aircraft propulsion system (20). The housing structure (54) includes an engine case assembly (106) and a nacelle (108). The nacelle (108) forms at least a portion of an exterior flow boundary (56) for the aircraft propulsion system (20). The nacelle (108) forms a guide vane structure receptacle (136) at the exterior flow boundary (56). The nacelle (108) further includes a nacelle interior skeletal structure (140) disposed at the guide vane structure receptacle (136). The turbine engine core (76) is housed within the engine case assembly (106). The turbine engine core (76) is configured to drive rotation of the open propulsor rotor (34) about the axis (24).

An assembly for an aircraft propulsion system (20) includes a housing structure (54), a guide vane structure (36), and a front mount assembly (148). The housing structure (54) includes an engine case assembly (106). The engine case assembly (106) includes a front center body (122). The guide vane structure (36) includes a guide vane structure base (50). The front mount assembly (148) is mounted on the front center body (122). The front mount assembly (148) is disposed axially coincident with the guide vane structure (36). The front mount assembly (148) includes at least one pylon attachment linkage (154).

An aircraft powerplant (20) assembly includes a first electric device (108A) and a cooling system (36). The first electric device (108A) includes a first device heat exchanger (160A), a first device housing (118A) and first electrical circuitry (120A). The cooling system (36) includes a liquid circuit (138), an air circuit (136) and a system heat exchanger (140). The cooling system (36) is configured to cool the first electrical circuitry (120A) using a liquid flowing in the liquid circuit (138) and air flowing in the air circuit (136). The liquid circuit (138) extends through the system heat exchanger (140) and the first electric device (108A). The first device heat exchanger (160A) is configured to transfer heat energy from the first electrical circuitry (120A) into the liquid in the liquid circuit (138). The air circuit (136) extends through the system heat exchanger (140). The system heat exchanger (140) is configured to transfer at least some of the heat energy from the liquid in the liquid circuit (138) into the air in the air circuit (136).

A method of inspecting erosion of a coating (16) on a component (10) at an outlet (26) of a through-hole (12) formed by waterjet drilling includes positioning a camera (42) to face the outlet (26); positioning a light source (46) at a first position (L1) from an inlet (24) opposite the outlet (26) of the through-hole (12); capturing a first image data of the outlet (26) and surrounding coating (16) with the camera (42) while the light source (46) is at the first position (L1); positioning the light source (46) at a second position (L2) from the inlet (24), wherein the first position (L1) is located between the inlet (24) and the second position (L2); capturing a second image data of the outlet (26) and surrounding coating (16) with the camera (42) while the light source (46) is at the second position (L2); and determining an amount of erosion of the coating (16) at the outlet (26) by comparing the first image data to the second image data.

An alignment tool including a base ring having a base ring inside diameter bounded by a base ring inner surface, the base ring comprising a top surface opposite a bottom surface; a stabilizer ring having a stabilizer ring inside diameter bounded by a stabilizer ring inner surface, the stabilizer ring comprising a stabilizer ring upper surface opposite a stabilizer ring lower surface; a biasing member in operative communication between the base ring top surface and the stabilizer ring lower surface; a gap formed between the base ring and the stabilizer ring; and at least one pull tab formed from the base ring.

A system for evaluating fatigue for a component design may be collectively operable to execute an evaluation environment, which may be operable to access a baseline component design associated with a baseline component. A baseline set of finite elements may be associated with a volume of the baseline component design. The evaluation environment may be operable to access a second component design associated with a second component. A second set of finite elements may be associated with a volume of the second component design. The evaluation environment may be operable to calculate at least one stress-based MAC value for the second component design based a comparison between nodal stresses associated with respective corner nodes of the baseline set of finite elements and nodal stresses associated with respective corner nodes of the second set of finite elements. A method for evaluating fatigue for a component design is also disclosed.

A start mechanism for a propulsion system including a receptacle in operative communication with a case; at least one portable source of energy disposed within the receptacle; a flow channel fluidly coupled to the receptacle; and a turbine flow path within the case, the turbine flow path fluidly coupled to the flow channel; wherein the at least one portable source of energy is configured to produce at least one of electrical energy and a source of chemical energy.

An assembly for an aircraft propulsion system (20) includes a housing structure (54). The housing structure (54) includes an engine case assembly (106) and a nacelle (108). The nacelle (108) is disposed radially outboard of the engine case assembly (106). The nacelle (108) forms at least a portion 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 nacelle (108) includes a nacelle interior skeletal structure (140) mounted to the engine case assembly (106) at a forward mount (146) and an aft mount (148).

An aircraft powerplant (20) assembly includes a rotor, a shroud (124), a casing (126), a structure (104) and an air system (102). The structure (104) circumscribes the casing (126) with a plurality of air plenums (144) formed by and disposed radially between the structure (104) and the casing (126). The air plenums (144) are arranged circumferentially around the axis (34) in an array. The air plenums (144) include a first plenum (144A) and a second plenum (144B). The air system (102) is configured to control a clearance (148) between the shroud (124) and an array of rotor blades (130) using air. The air system (102) includes a first air circuit (110) and a second air circuit (110). The first air circuit (110) is configured to direct a first flow of the air into the first plenum (144A) to cool a first section of the casing. The second air circuit (110) is configured to direct a second flow of the air into the second plenum (144B) to cool a second section of the casing (126).

An operating method is provided for an aircraft. During this method, a first control signal is provided to drive rotation of a first propulsor rotor of a first propulsion system of the aircraft using a first engine of the first propulsion system. Operation of the first engine is monitored to provide first engine data. A degraded condition of the first engine is identified based on the first engine data. In response to the identifying of the degraded condition, a second control signal is provided to drive the rotation of the first propulsor rotor at a first rate using a first electric machine of the first propulsion system. The first rate is equal to or greater than fifty percent of a second rate at which a second propulsor rotor of a second propulsion system of the aircraft is being rotated.

A gas turbine engine (20) is provided that includes first and second spools (37, 39), an accessory gearbox (44), and a spool coupling device (46). The accessory gearbox, AG, (44) is in drive communication with the second spool (39). The accessory gearbox (44) includes an AG output shaft. The spool coupling device, SCD, (46) is in drive communication with the first spool (37) and the AG output shaft. The spool coupling device (46) includes an SCD accessory dock (66) and a transmission (62) that is disposable in first or second spool engaged configurations. In the first spool engaged configuration, rotational drive from the first spool (37) is provided to the SCD accessory dock (66), and in the second spool engaged configuration, rotational drive from the second spool (39) is provided to the SCD accessory dock (66).

A turbine engine rotor has a central longitudinal axis and includes: a central shaft having an outer diameter seal groove (200); a disk stack having a number of disks encircling the shaft; and a seal (120) captured in the groove (200) and facing or engaging an inner diameter surface (170) of one of the disks. The seal (122) is wound about the central longitudinal axis for at least 400°.

A turbine vane (94) includes a vane body (100) and a baffle (102). The vane body includes an inner side forming an internal cavity (122). The vane body further forms at least one retention feature recess (172) on the inner side. The baffle is disposed within the internal cavity and includes a tubular body which subdivides the internal cavity into an outer cavity portion and an inner cavity portion. The outer cavity portion is formed between the inner side and the baffle tubular body. The inner cavity portion is circumscribed by the tubular body. The tubular body includes at least one protruding retention feature (150). The at least one protruding retention feature is disposed within the at least one retention feature recess. The at least one protruding retention feature disposed within the at least one retention feature recess obstructs radial movement of the baffle relative to the vane body.

An assembly is provided for an aircraft powerplant (20). This assembly includes an aircraft engine, a control system (98) and a conduit (112). The control system includes an electronic engine controller (104), a first electric device and a first electric cable electrically coupling the first electric device to the electronic engine controller. The electronic engine controller is configured to control operation of the aircraft engine using the first electric device. The conduit extends longitudinally from a first end (114) of the conduit to a second end (116) of the conduit. The first electric cable extends longitudinally in a bore (126) of the conduit between the electronic engine controller and the first electric device. The conduit is configured to shield the first electric cable from electromagnetic interference generated outside of the conduit.

A device (60) may include an inner sleeve (66) having a cylindrical body (72) defined by a radially inner surface (74) and a radially outer surface (76). The device may include an outer sleeve (68) positioned outside of the inner sleeve (66), the outer sleeve (68) including at least one slit (82) extending axially along the outer sleeve (68). A device may include a rubber component (70) disposed between the radially outer surface (76) and the outer sleeve (68), wherein the rubber component (70) circumscribes at least a portion of the inner sleeve (66).

A system for and method of operating a propulsion system for an aircraft is provided. The propulsion system includes a gas turbine engine (20), an accessory gearbox (42), and an electric motor (48). The gas turbine engine (20) includes a spool (37) having a fan rotor (28) with a plurality of fan blades (28A), and a pitch control mechanism (58) engaged with the fan blades (28A). The method includes: providing rotational drive from the spool (37) to the accessory gearbox (42); using the accessory gearbox (42) to drive the electric motor (48) to produce electrical energy, wherein the electric motor (48) has a predetermined maximum rotational speed limit; and controlling the propulsion system to follow a predetermined descent portion (60) of a flight profile for the aircraft, wherein controlling the propulsion system includes controlling the pitch control mechanism (58) to alter an amount of thrust produced by the fan blades (28A) by altering a pitch of the fan blades (28A).

An optical inspection method (300) includes capturing one or more low-speed images (414) of a plurality of fans in a fan rotor, storing the one or more low-speed images (414) in a historical database (420) of an optical inspection system (400), and using the one or more low-speed images (414) in an inspection algorithm (500) of the optical inspection system (400) to produce an inspection result (440). Capturing the one or more low-speed images (414) for the plurality of fans in the fan rotor may include capturing the one or more images (404) of the plurality of fans while the fans are rotating at less than 100 rpm using a camera (402). Using the one or more low-speed images (414) in the inspection algorithm (500) may include using the one or more low-speed images (414) as reference images (404) for the inspection algorithm (500) to determine a presence of an irregularity on the plurality of fans.

An airfoil (102) extends longitudinally between a leading edge (120) and a trailing edge (122), laterally between a first side (124A) and a second side (124B), and spanwise from a base (108) to a tip (112). The airfoil (102) includes a body section (126), a tip section (128) and a channel (130). The body section (126) extends spanwise from the base (108) to a shelf (136), and laterally between the first side (124A) and the second side (124B). The tip section (128) is connected to the body section (126), extends spanwise from the shelf (136) to the tip (112) and laterally between the second side (124B) and a sidewall (140). The channel (130) projects spanwise into the airfoil (102) from the tip (112) to the shelf (136) and laterally into the airfoil (102) from the first side (124A) to the sidewall (140). The tip section (128) is formed with a density that varies along the channel (130).

A device may include a mount bushing assembly (60) including an inner sleeve (66) having a cylindrical body (72) defined by a radially inner surface (74) and a radially outer surface (76), an outer sleeve (68) positioned outside of the inner sleeve (66), and a rubber component (70) disposed between the radially outer surface (76) and the outer sleeve (68). The device may include a mounting pin (65) received in the inner sleeve (66) and connected to a mounted structure, wherein the outer sleeve (68) includes a strain relieving structure (82) at an outer peripheral surface (80) along the outer sleeve (68), and wherein the rubber component (70) circumscribes at least a portion of the inner sleeve (66).

An acoustic treatment proximate a fan stage (12) including the acoustic treatment (42) disposed in at least one location along surfaces exposed to a fan discharge airflow downstream from a fan within the fan stage (12).

An assembly is provided for an aircraft propulsion system (20). This assembly includes a fan rotor (44), an engine core (38), an engine flowpath, a bypass flowpath (42), an inter-flowpath structure (108) and an air circuit (130). The inter-flowpath structure (108) includes a structure leading edge, a structure inner side and a structure outer side. The inter-flowpath structure (108) projects axially to the structure leading edge next to and downstream of the fan rotor (44). The structure inner side borders the core flowpath (40). The structure outer side borders the bypass flowpath (42). The air circuit (130) includes a circuit inlet (132) fluidly coupled to the core flowpath (40) and a circuit outlet (134) fluidly coupled to the bypass flowpath (42). The air circuit (130) extends within the inter-flowpath structure (108) from the circuit inlet (132) to the circuit outlet (134). The circuit inlet (132) is disposed along the structure inner side and is upstream of the compressor section (33). The circuit outlet (134) is disposed along the structure outer side.

An aircraft powerplant assembly (22) includes a gearbox housing (116), a power transfer shaft (166), a first gear system (146), a second gear system (158) and a power coupler (104). The second gear system is operationally independent of the first gear system. The power coupler is disposed within the interior of the gearbox housing. The power coupler is arranged with the power transfer shaft and is disposed between the first gear system and the second gear system. The power coupler is configurable in a first arrangement and a second arrangement. In the first arrangement, the power coupler operatively couples the first gear system to the power transfer shaft and operatively decouples the second gear system from the power transfer shaft. In the second arrangement, the power coupler operatively couples the second gear system to the power transfer shaft and operatively decouples the first gear system from the power transfer shaft.

An aircraft system (20) includes a powerplant (24) configured to drive rotation of a propulsor rotor (22) during a first mode and a second mode. The powerplant (24) includes a turbine engine (38), a first electric machine (98A) and a power coupler (112). A first rotating structure includes a first turbine rotor disposed along a core flowpath (54). The first electric machine (98A) is configured as an electric motor during the second mode. The first electric machine (98A) is disposed between the propulsor rotor (22) and the turbine engine (38). The first electric machine (98A) includes a first electric machine rotor (102A) coupled to the propulsor rotor (22). The power coupler (112) is disposed between the first electric machine (98A) and the turbine engine (38). The power coupler (112) is configured to couple the first rotating structure to the first electric machine rotor (102A) during the first mode and to decouple the first rotating structure from the first electric machine rotor (102A) during the second mode.

A thin walled structure (100) includes a primary structure wall (104) connected to a shielding wall (102) with a plurality of support pins (106) that extend from the shielding wall (102) to the primary structure wall (104). A method for printing a thin walled structure (100) includes the steps of: modeling, using distortion analysis software, a method of making a primary structure wall (104) to generate a distortion analysis; inputting the distortion analysis into a digital model of the primary structure wall (104) to determine a desirable layout and thickness for a support wall (102) and a plurality of support pins (106) that will connect with the primary support structure wall (104) to form a thin walled structure (100); integrating the desirable layout and thickness for a support wall (102) and a plurality of support pins (106) into a build model of the primary structure wall (104) to create a supported build model; and printing, using the supported build model, the thin walled structure (100).

A fan inlet case strut 62 including a strut structural section 70; a strut cover 74 disposed over a portion of the strut structural section 70; a strut airfoil section formed with a high pressure surface and a low pressure surface connected at a leading edge and a trailing edge and extending from a first end to a second end, a camber line that extends between the leading edge and the trailing edge and a flow feature 88 formed at a strut cover termination.

An aircraft system includes a propulsion system (24A,24B) mounted to an airframe (22). The propulsion system (24A,24B) includes a fan rotor (82), an engine core (72), an outer housing structure (62), an inner housing structure (60), a bypass flowpath (77) and a core flowpath (76). The engine core (72) is configured to drive rotation of the fan rotor (82) about an axis (54). The engine core (72) includes a compressor section (67), a combustor section (68) and a turbine section (69). The outer housing structure (62) houses the fan rotor (82) and forms an outer peripheral boundary of the bypass flowpath (77). The inner housing structure (60) forms an inner peripheral boundary of the bypass flowpath (77). The core flowpath (76) includes a core inlet (112) and a core exhaust (116). The core flowpath (76) extends longitudinally through the compressor section (67), the combustor section (68) and the turbine section (69) from the core inlet (112) to the core exhaust (116). The core inlet (112) is located axially forward of the fan rotor (82) and radially between the axis (54) and the airframe (22).

A gas turbine engine (20) includes a turbine section (28) having at least one turbine rotor. A compressor section (24) includes at least one compressor rotor, and a propulsor (119). The at least one turbine rotor is connected to drive an epicyclic gear reduction to in turn drive the propulsor (119) at a slower speed than the at least one turbine rotor. The epicyclic gear reduction includes a sun gear (110) for receiving a drive input from the at least one turbine rotor and having gear teeth engaging gear teeth (104) on a plurality of intermediate gears (102). The intermediate gears (102) are mounted within a carrier (112). The teeth (104) on the intermediate gears (102) also engage teeth (106) on a ring gear (100). One of the carrier (112) and the ring gear (100) is connected to drive a propulsor drive shaft (120) to in turn drive the propulsor (119), and the other of the carrier (112) and the ring gear (100) connected to a static structure, with the carrier (112) connected to either the propulsor drive shaft (120) or the static structure through a torque frame (111). The torque frame (111) is connected to the carrier (112) at a radially outer location through a splined connection.

A gas turbine engine (20) includes a static housing (71) defining a bearing compartment (58). A lubricant seal assembly (68) is adjacent to the static housing (71) to bound the bearing compartment (58). A tube (80) is routed through the static housing (71) and is configured to convey fluid. The tube (80) includes a first end (96) and a second end (98). The first end (96) includes a grooved neck (108) which includes an annular body (110) with an inner circumferential surface (112) and an outer circumferential surface (114) surrounding the inner circumferential surface (112). The outer circumferential surface (114) includes a plurality of longitudinal grooves (116). The second end (98) is threaded into the static housing (71). A method of installing a tube (80) within a static housing (71) of a mid-turbine frame (57) of a gas turbine engine (20) is also disclosed.

A gas turbine engine is provided that includes a fan section, a compressor section, a combustion section, a turbine section, an annular fan bypass air path, a bleed air structure (54), an annular compartment (64), and an anti-icing system (68). The compressor section has a low (30A) and high (30B) pressure compressor sections. The fan bypass air path is defined by inner (52B) and outer (52A) bypass wall structures. An outer radial boundary structure defines an outer radial boundary of a core gas path within the low pressure compressor section (30A). The bleed air structure is disposable in an open state or in a closed state. The annular compartment (64) is defined by the inner bypass wall structure (52B), the outer radial boundary structure, and a firewall (66). The anti-icing system (68) is disposable in an activated state or in a deactivated state. In the activated state the anti-icing system (68) is configured to mitigate ice accumulation within the annular compartment (64).

The disclosure describes assembly clips (200; 300; 400; 500) for use in attaching brush seals (100) to static hardware components (160; 165) during the assembling of gas turbine engines (20). The clips (200; 300; 400; 500) are made of plastic which melts and vaporize during operation of the engine (20). The temporary clips (200; 300; 400; 500) provide an efficient means for attaching brush seals (100) to subassemblies of gas turbine engines (20) and which reduce or eliminate the potential for foreign object damage or domestic object damage risks.

An assembly (100) for a gas turbine engine (20), the assembly (100) comprising: a CMC vane (105) having a CMC vane inner diameter platform (145), wherein the CMC vane inner diameter platform (145) includes an upper sealing interface surface (144a) comprising an upper surface (140a) of a groove (140); one of a tangential onboard injector (110) or an inner air seal outer diameter, wherein said one of the tangential onboard injector (110) or the inner air seal outer diameter includes a lower sealing interface surface (144b) comprising a lower surface (140b) of the groove (140), wherein the CMC vane inner diameter platform (145) and said one of the tangential onboard injector (110) or the inner air seal outer diameter are disposed in contact to form an interface (144) comprising the upper sealing interface surface (144a) and the lower sealing interface surface (144b); and a seal (135; 200; 300; 400) disposed within the groove (140) formed by the upper surface (140a) and the lower surface (140b).

An assembly includes a compressor section (33), a core flowpath (40), a bypass flowpath (42), a flowpath wall and an air circuit (124). The flowpath wall forms an inner peripheral boundary of the bypass flowpath (42). The flowpath wall includes a wall base (92) and a wall protrusion (94). The wall base (92) extends axially along and circumferentially about the axis. The wall protrusion (94) projects radially out from the wall base (92). The wall protrusion (94) extends axially along the axis from a protrusion first end to a protrusion second end. The protrusion first end is upstream of the protrusion second end along the bypass flowpath (42). The air circuit (124) extends from a circuit inlet (126) to a circuit outlet (128). The circuit inlet (126) is fluidly coupled to the core flowpath (40). The circuit outlet (128) is fluidly coupled to the bypass flowpath (42). The circuit outlet (128) is formed in the wall protrusion (94) axially between the protrusion first end and the protrusion second end.

An aircraft propulsion system (20) assembly includes a fan rotor (44), an inner flowpath (40), an outer flowpath (42) and an inter-flowpath structure (122). The fan rotor (44) includes a plurality of fan blades (48) arranged circumferentially about an axis (26). The outer flowpath (42) includes an outer flowpath inlet (94) downstream of the fan rotor (44). The outer flowpath inlet (94) is radially outboard of an inner flowpath inlet (90). The inter-flowpath structure (122) is disposed radially between and partially forms the inner flowpath inlet (90) and the outer flowpath inlet (94). The inter-flowpath structure (122) includes a splitter (128) and an actuator (142). The splitter (128) forms a leading edge (134) of the inter-flowpath structure (122). The actuator (142) is configured to translate the splitter (128) axially along the axis (26) between a first position and a second position.

A key for a keylock insert including a key body having a key forward end opposite a key aft end; the key body has a base side and an interference side opposite the base side; the key body has a first lateral side opposite a second lateral side relative to a key body centerline; the key has a tang portion proximate the key forward end; the key has a double peak portion adjacent the tang portion and proximate the key aft end; the key has a transition region located between the tang portion and the double peak portion; wherein the transition region is configured as a sloped surface; the double peak portion includes a first peak extending from the base side and a second peak adjacent the first peak relative to the key body centerline.

A ceramic matrix composite (CMC) airfoil component can be made by forming a stack of fabric plies with a plurality of plies adhered together with a tackifier. The stack of fabric plies is machined into an airfoil shape to form an airfoil core (220) with an orientation of the fabric plies selected based upon strength and/or performance requirements of the airfoil component. The airfoil core (220) is then overwrapped with an outer mold layer (230) formed of one or more plies of fabric to form a preform (200). The preform (200) is then densified, such as via melt infiltration (MI) and/or chemical vapor infiltration (CVI), to form the CMC airfoil component such as a turbine vane or a turbine blade.