Abstract: Provided is a vehicle capable of suppressing vibration of a drive unit having an engine and a motor to be small and having high vehicle motion performance. The drive unit for travel of the vehicle includes an engine and a motor. The engine is a rotary engine and has an engine output shaft and an engine housing. The motor has a motor output shaft, a rotor, and a stator that are accommodated in a motor housing. The engine housing of the engine and the motor housing of the motor are directly joined to each other. The engine output shaft and the motor output shaft are directly joined to each other. The rotor of the motor has: a main body section that generates a rotational magnetic field with the stator; and a weight section that functions as a balance weight of the engine.

Abstract: A fixture for transferring turbine blades may include: a first body having an arcuate radially inward surface shaped to contact a rotor of a turbomachine, and a radially outward surface including a plurality of dovetail slots therein shaped to engage the dovetails of the plurality of turbine blades; and a first alignment aperture extending axially through the first body relative to a centerline axis, and positioned for alignment with a portion of the rotor wheel such that the plurality of dovetail slots of the first body are substantially axially aligned with the plurality of dovetail slots of the rotor wheel for at least partial transfer of the turbine blade thereto from the fixture, wherein the dovetails of the plurality of turbine blades are slidably removable from the plurality of dovetail slots of the first body for guided insertion into the plurality of dovetail slots of the rotor wheel.

Abstract: A closure having enhanced oxygen barrier capabilities includes a cap body formed from a first material, the cap body defining a cavity therethrough, a stopper formed from a second material, the second material being different than the first material and disposed within the cavity, and a barrier layer formed from a third material, the third material being different than at least the second material and associated with at least one of the cap body and the stopper. A container assembly including a closure and a method of forming the closure is also provided.

Abstract: A gas turbine engine turbine has a high pressure turbine configured to rotate with a high pressure compressor as a high pressure spool in a first direction about a central axis and a low pressure turbine configured to rotate with a low pressure compressor as a low pressure spool in the first direction about the central axis. A power density is greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/cubic inches. A fan is connected to the low pressure spool via a speed changing mechanism and rotates in the first direction.

Abstract: A turboprop engine including an annular outer case including a mount ring for attachment to an aircraft along a plane to transfer loads from a propeller and having an annular flange extending radially inwardly therefrom within the plane, an annular turbine support case received within the outer case and connected to the outer case only through a direct connection with the annular flange allowing a limited relative pivoting motion between the turbine support case and the mount ring, and a turbine section including a rotor closely surrounded by a shroud with an annular tip clearance being defined therebetween, the shroud being directly connected to and located by the turbine support case. A method of isolating a turbine shroud from propeller loads in a propeller engine is also disclosed.

Abstract: An integrated single-piece exhaust system (SPEX) with modular construction that facilitates design changes for enhanced aerodynamics, structural integrity or serviceability. The SPEX defines splined or curved exhaust path surfaces, such as a series of cylindrical and frusto-conical sections that mimic curves. The constructed sections may include: (i) a tail cone assembly fabricated from conical sections that taper downstream to a reduced diameter; or (ii) an area-ruled cross section axially aligned with one or more rows of turbine struts; or both features. Modular inner and outer diameter inlet lips enhance transitional flow between the last row blades and the SPEX, as well as enhance structural integrity. Modular strut collars have large radius profiles between the SPEX annular inner diameter and outer diameter flow surfaces, for enhanced airflow and constant thickness walls for uniform heat transfer and thermal expansion. Scalloped mounting flanges enhance structural integrity and longevity.

Abstract: A rotor (10) has a basic rotor body (12) and a plurality of rotating blades (14) mounted on the basic rotor body (12). The rotating blades (14) in this case are mounted rigidly or non-detachably, in particular, cohesively, on the basic rotor body (12). Thus, at least one rotating blade (14) has at least one integral sealing element (24), by means of which a root intermediate space (22) is sealed in the region of a blade root (20) radially underneath a blade platform (18) of the rotating blade (14). In addition, the invention relates to a method for producing a rotor (10), in particular, for an aircraft engine, in which a plurality of rotating blades (14) is mounted on a basic rotor body (12), wherein at least one rotating blade (14) having at least one integral sealing element (24) is mounted on the basic rotor body (12).

Abstract: An air intake system for use with a gas turbine is provided. The air intake system includes a weather hood coupled to a filter house. The air intake system also includes a prefilter assembly configured within said weather hood, wherein said prefilter assembly includes at least three air filtration units, including at least one moisture separator and at least one coalescing filter.

Abstract: One embodiment includes a fuel manifold segment for supplying fuel to a fuel injector. The fuel manifold segment contains a fuel line surrounded by a first firesleeve. A pigtail line connects to the fuel line. A connector on the pigtail line connects to a fuel injector inlet fitting. A second firesleeve surrounds the first firesleeve. A cuff surrounds the pigtail line, portion of first firesleeve, and portion of fuel line to which the pigtail line is connected. A boot then surrounds the connector.

Abstract: A method of fabricating a mixer for a gas turbine engine is provided. The method includes forming a forward end and an aft end of the mixer, and forming an annularly undulating contour that defines a plurality of core immersion lobes and a plurality of bypass immersion lobes between the forward end and the aft end. The plurality of bypass immersion lobes includes a first bypass immersion lobe and a second bypass immersion lobe. The first bypass immersion lobe has a first crown contour line extending from the forward end to the aft end of the mixer, and the second bypass immersion lobe has a second crown contour line extending from the forward end to the aft end of the mixer. The first crown contour line is different than the second crown contour line.

Abstract: A method for producing a metal reinforcement for protecting a leading edge of a compressor blade of composite, including: creating a core that has a shape of an internal cavity of the reinforcement; creating an insert made of an alloy of a hardness greater than that of the reinforcement; shaping sheet metal by stamping with the creation, upstream of the core, a cavity between the metal sheets, which cavity is configured to accept the insert, positioning the sheets around the core with the insert placed in the cavity and securing the assembly together; creating a vacuum and closing the assembly by welding; consolidation by hot isostatic pressing; cutting the assembly to extract the core and separate the reinforcement; creating an external profile of the reinforcement by a final machining operation that reveals a material of the insert.

Abstract: A drive assembly for a torque converter is provided. The drive assembly includes a turbine shell including a rounded portion and an outer radial extension on an outer circumference of rounded portion; and a tab plate bonded to the outer radial extension of the turbine shell by adhesive. A method of forming a drive assembly for a torque converter is also provided.

Abstract: The invention concerns a method for manufacturing of a gas turbine engine component (37) comprising an outer ring structure (42, 47), an inner ring structure (41), and a plurality of circumferentially spaced elements (46, 46a, 46b) extending between the inner ring structure (41) and the outer ring structure (42), wherein a primary gas channel for axial gas flow is defined between the elements (46, 46a, 46b), and wherein the component (37) has an inlet side for gas entrance and an outlet side for gas outflow. The invention is characterized in that the method comprises the step of machining a one-piece metal blank as to form a one-piece part (47) comprising: a portion (46b) of each of said elements (46), wherein said portion (46b) relates to a portion of an extension length of the elements (46) between said ring structures (41, 42); and a ring-shaped member (42) that connects said element portions (46b) and that is intended to form part of one of the ring structures.

Abstract: Disclosed is an offset rotational non-reciprocating-piston internal combustion engine in which an outer ring rotates on a one axis and an inner disk rotates on a second offset axis. Pistons are connected to the outer ring, while cylinders and other devices for operating the engine are mounted within an inner disk that rotates on a second axis that is offset from the axis of the outer ring. The inner disk and outer ring rotate together, such that the pistons create conditions of compression and explosive expansion within the cylinders without vibrating reciprocal piston motion. A unique fuel injection system is also disclosed that provides a variable fuel pressure that is created by centrifugal forces on the fuel. Because of the rotating inner disk and outer ring, advantages are taken of centrifugal force and gravity to distribute fuel, air, oil, high-voltage current, and cooling air in the engine.

Abstract: A rotatory internal combustion engine is provided. The engine includes a power module, wherein all moving parts of the engine are positioned within the power module, a housing including an intake and an exhaust, wherein the housing is configured to retain the power module, and a sleeve configured to couple in the housing, the sleeve including a sleeve intake and a sleeve exhaust.

Abstract: A method of forming a microchannel cooled component is provided. The method includes forming at least one microchannel within a surface of a relatively planar plate. The method also includes placing a relatively planar cover member over the surface having the at least one microchannel formed therein. The method further includes adhering the relatively planar cover member to the relatively planar plate. The method yet further includes curving the microchannel cooled component by pressing the relatively planar cover member with a forming component for at least a portion of a time period of adhering the relatively planar cover member to the relatively planar plate.

Abstract: A method for coating a turbine engine component comprises the steps of: providing a turbine engine component having at least one sacrificial attachment on a first side; grasping the turbine engine component via the at least one sacrificial attachment to position a first surface of the turbine engine component relative to a source of coating material; and applying a coating to said first side.

Abstract: A gage tool (700) for inspecting tolerances of a gas turbine engine seal plate (430) includes a base plate (710) arid a top plate (720). The base plate (710) includes abase opening (712) and a slot (711) with an annular shape sized to receive a seal plate (430). The top plate (720) Includes a top opening (723). The base opening (712) and top opening (723) each provide access to a portion of a seal plate (430) within the gage tool (700). The gage tool (700) also includes a first gage (750) with a probe tip for measuring a variation in an outer interlacing surface (433) of a seal plate (430) and a seal slide gage (740) for measuring a force required to rotate Ore seal plate (430) within the gage tool.

Abstract: According to one aspect of the invention, a method is provided for modifying an airfoil shroud located at a tip of an airfoil of a airfoil, the airfoil shroud having a first end edge, a second end edge, a leading edge and a trailing edge. The method includes locating a reference location in the first end edge of the airfoil shroud, the reference location being proximate a seal rail extending circumferentially from the substantially horizontal surface and forming a relief cut in the airfoil shroud to remove the reference location, wherein a modifying of the airfoil shroud is complete following forming of the relief cut.

Abstract: Embodiments of the disclosure include a combustor assembly. The combustor assembly may include one or more fuel plenums. The combustor assembly may also include one or more fuel distribution plates disposed within the fuel plenums. Moreover, the combustor assembly may include a number of mixing tubes disposed at least partially within the fuel plenums and extending through the fuel distribution plates. In certain aspects, the mixing tubes may each include a reduced diameter about the fuel distribution plates to form an annulus therebetween.

Abstract: A hanger assembly for use between a first duct and a second duct having an opening therein includes a spring having a first end and a second end, a first mount for attaching the first end to the first duct, a second mount for attaching the second end to the second duct, the second mount having an area greater than the opening, and a central aperture therethrough.

Abstract: An exemplary gas turbine engine component assembly includes, among other things, a case having a generally cylindrical outer surface. At least one synchronizing ring is received at least partially about the outer surface of the case. A plurality shims situated between an inner surface on the at least one synchronizing ring and the outer surface of the case at a plurality of shim locations are circumferentially spaced about the case outer surface. Each shim location includes at least one of the shims. A radial thickness of the shims at a first one of the shim locations is different than the radial thickness of the shims at a second one of the shim locations.

Abstract: A method for manufacturing an aircraft component according to one embodiment of this disclosure includes providing a machining system including a controller, at least one sensor, and a tool for machining. The method further includes providing an aircraft component, and machining the aircraft component with the tool based on feedback from the at least one sensor.

Abstract: A boroscope (60) has a first end (62) and a second end (64) and the first end (62) of the boroscope (60) has an optical fibre (66) and light source (68). A working head (70) is attached to the first end (62) of the boroscope (60). The working head (70) has an electrical motor (72) and a tool (74) is attached to and is arranged to be driven by the electrical motor (72) and the boroscope (60) carries a cable (76) extending from the electrical motor (72) to the second end (64) of the boroscope (60).

Abstract: A method for top-loaded assembly of an internal combustion engine includes providing a carrier base having a lower portion configured as a crankcase and an upper portion having a number of upwardly directed assembly provisions such as a crankshaft bay and at least one cylinder deck. Arranging the crankshaft bay and cylinder deck so that the engine may be assembled from a position above the engine, without the necessity of attaching connecting rods or crankshaft main bearing bolts from underneath the engine.

Abstract: A heat exchanger assembly for use in a gas turbine engine includes a bypass valve and at least one body portion. The body portion includes at least one de-congealing inlet channel in flow communication with the bypass valve, a plurality of cooling channels in flow communication with the bypass valve and the at least one de-congealing inlet channel, and at least one de-congealing outlet channel in flow communication with the bypass valve and the at least one de-congealing inlet channel. The bypass valve is configured to deliver a fluid between the at least one de-congealing inlet channel and the plurality of cooling channels during a first mode of operation to facilitate reducing a temperature of the fluid. The bypass valve is further configured to deliver the fluid between the at least one de-congealing inlet channel and the at least one de-congealing outlet channel during a second mode of operation.

Abstract: An apparatus is provided and includes a combustor, a transition piece and a flow conditioner. The combustor has a head end and defines a first flowpath along which a main flow is directable to flow in a downstream direction from the head end. The transition piece is disposable downstream from the combustor, has a forward end and defines a second flowpath along which the main flow is directable to flow in the downstream direction from the forward end. The flow conditioner includes an inner member coupled to the transition piece forward end, an outer member and aerodynamic elements supportively disposed between the inner and outer members to be interactive with a flow of fluid proceeding toward the combustor head end.

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion, a cooling circuit disposed within the body portion and including at least a first cavity and a microcircuit in fluid communication with the first cavity. A plunged hole intersects at least a portion of the microcircuit.

Abstract: In a gas turbine engine, a fan rotor and a compressor rotor are connected to a joint which is attached to a shaft. The fan rotor is connected to the joint by a plurality of fasteners extending through mounting openings in the joint and through apertures in the fan rotor. A standard retaining ring is attached to the joint to cover a portion of an enlarged head of each of the respective fasteners in order to prevent removal of the respective fasteners from the mounting openings in the joint before the respective fasteners are connected to the fan rotor.

Abstract: An airfoil assembly including an endwall and an airfoil extending from the into a gas flow path. The endwall includes upstream and downstream edges, and is defined on a platform structure having a front surface extending radially in a direction of a thickness of the platform structure. At least one fluid injection passage extends through the platform structure in a direction from the upstream edge toward the downstream edge of the endwall. The fluid injection passage has an outlet opening defined at the endwall and an inlet opening in fluid communication with a pressurized fluid source. The fluid injection passage extends at a shallow angle relative to a plane of the endwall wherein the fluid injection passage defines a passage axis passing through the front surface and the endwall for effecting energization of a boundary layer between the outlet opening and the airfoil leading edge.

Abstract: A method is disclosed for manufacturing a tie shaft for aero or land based gas turbine engine. The method includes flow forming a tie shaft preform to produce a tie shaft. In one example, the tie shaft includes a nickel alloy cylindrical wall having a length to diameter ratio of at least 6:1, wherein the diameter is an average outer diameter. The wall includes a minimum effective strain of 0.3 in/in (7.6 mm/mm), and a grain size is in the range of G4 to G16 per ASTM E112. The wall includes a roll formed threaded surface having a thread roughness of less than 1260 ?in (32 microns).

Abstract: A method for manufacturing a fan blade shield having a sheath cavity made from a hard metal material is disclosed The method may comprise: performing a sheath cavity grinding step for the fan blade shield; and performing a sheath cavity bottom grinding step for the fan blade shield. A method for forming a fan blade is disclosed. The method may comprise: obtaining a fan blade body; obtaining a fan blade shield made from a hard metal material and having a sheath cavity match a leading edge of the fan blade body, the fan blade shield being obtained by a process comprising: performing a sheath cavity grinding step for the fan blade shield; and performing a sheath cavity bottom grinding step for the fan blade shield; and attaching the fan blade shield to the leading edge of the fan blade body to produce the fan blade.

Abstract: A gas turbine engine includes a combustor, a first turbine section in fluid communication with the combustor, a second turbine section in fluid communication with the first turbine section, and a mid-turbine frame located axially between the first turbine section and the second turbine section. An oil system includes a first oil system component that houses oil. A first connector mechanically connects the first oil system component to the mid-turbine frame.

Abstract: An exhaust system for use with a steam turbine is provided. An exhaust hood includes an input and an output, the input receiving fluid from the steam turbine. The exhaust hood includes a first side wall that extends between the input and the output. The first side wall includes an aperture. An ejector is coupled to the exhaust hood. The ejector includes inlets and an outlet. At least one of the inlets receives fluid from the exhaust hood via the aperture.

Abstract: An air bleed system for a jet engine is provided. The air bleed system is located in the low pressure compressor chamber and comprises a movable bleed valve and a stationary, annular bleed case having a forward section and an aft section and ligaments connecting the two sections and providing structural support to the bleed case. The ligaments are machined from the bleed case into an aerodynamic shape that maximizes the flow of air through the bleed ports.

Abstract: According to one aspect of the invention, a method of manufacturing an impingement sleeve for a turbine engine combustor includes placing a first metal sheet in a first die, pressing the first metal sheet in the first die to form a first vertical half of the impingement sleeve, placing a second metal sheet in a second die and pressing the second metal sheet in the second die to form a second vertical half of the impingement sleeve. The method also includes forming holes in the first vertical half and second vertical half while the first vertical half and second vertical half remain separate, positioning the first vertical half and second vertical half about a transition piece and welding the first vertical half to the second vertical half to form the impingement sleeve.

Abstract: A turbine and a method of mitigating out-of-roundness effects at a turbine is disclosed. An inner turbine shell and an outer turbine shell of the turbine is provided. The inner turbine shell is coupled to the outer turbine shell using a ring insert. The ring insert is segmented into a plurality of ring insert segments that reduce a transfer of load from the outer turbine shell to the inner turbine shell to mitigate out-of-roundness of the inner turbine shell.

Abstract: Processes for fabricating secondary structures of gas turbine engines from polymer-based materials, and secondary structures formed thereby. The processes entail performing an additive manufacturing technique to produce a secondary structure of a gas turbine engine. The additive manufacturing technique directly produces the secondary structure from a polymer-based material to have a complex three-dimensional shape characterized by portions that lie in different planes.

Abstract: The present invention provides reduced fatigue stress levels in a stressed area of an engine crankcase by lowering the outer edges of the crankshaft bearing cap fracture-split lines without changing the central axis location of the crankshaft in the crankcase. This is accomplished by causing the fracture-split lines of the bearing caps to angle downward from the horizontal plane of the crankshaft axis to a slightly lower position at the outer edges of the bearing caps. This lowers the outer edges of the bearing caps sufficiently to allow larger radii to be formed on the lateral connectors of the crankcase support webs with the adjacent sidewalls of the crankcase. Accordingly, the fatigue stress levels in the crankcase web members are reduced without requiring a significant change in the crankcase dimensions.

Abstract: It is intended to provide a rotary machine provided with an atmosphere relief mechanism by which damage of a rupture disc when attaching the disc can be prevented and which can withstand use over a long period of time. The rotary machine is provided with a casing and an atmosphere relief mechanism 20 for closing an opening 12 for atmospheric release formed in the casing and relieving pressure of inner fluid to atmosphere when a pressure in the casing rises. The atmosphere relief mechanism 20 includes: a rupture disc 24 configured to rupture when the pressure in the casing reaches a predetermined pressure; a pair of annular holding parts 22, 27 arranged to hold an outer edge of the rupture disc 24 from both sides; a plurality of fastening members 29 for fastening the pair of annular holding parts 22, 27 to each other so as to hold the rupture disc 24 between the pair of annular holding parts 22, 27; and an annular spacer part 25 provided along an outer circumference of the rupture disc 24.

Abstract: Disclosed herein are apparatuses, methods, and systems for transition piece contouring. In an embodiment, a high thermal stress section of a transition piece and a low thermal stress section of a transition piece is a determined The low thermal stress section of the transition piece may be contoured to intercept hot gas flow.

Abstract: A gas turbine engine includes an engine core outer casing and a fan nacelle spaced radially outwardly relative to the engine core outer casing to define a bypass duct. A plurality of drag links is used to pivot blocker doors into a flow blocking position in the bypass duct when a thrust reverser is deployed. The plurality of drag links is located within the bypass duct in an area of non-uniform flow defined by a plurality of local airflow angles. Each drag link is individually configured to align with one of the local flow angles.

Abstract: A component for a gas turbine engine includes a wall and a cooling hole extending through the wall. The wall has a first surface and a second surface. The cooling hole includes a metering section extending downstream from an inlet in the first surface of the wall and a diffusion section extending from the metering section to an outlet in the second surface of the wall. The diffusion section includes a first plurality of lobes diverging longitudinally and laterally from the metering section on a first side of a centerline axis of the cooling hole and a second plurality of lobes diverging longitudinally and laterally from the metering section on a second side of the centerline axis.

Abstract: A combustor assembly for use with a gas turbine. The combustor assembly may include a liner, an impingement sleeve disposed about the liner, and an airflow channel defined between the liner and the impingement sleeve. One or more holes may be disposed through the impingement sleeve, and one or more tubulators may be disposed within the airflow channel.

Abstract: A raft assembly for a gas turbine engine is provided. The raft assembly includes a rigid raft formed of a rigid material that has an electrical system and/or a fluid system embedded therein. The raft assembly further includes one or more clamps for mounting tubular members to the raft. The or each clamp has a first clamp block and a second clamp block which, in use, clamp together to grip a tubular member between the blocks. The first block is fixed to the raft. The or each clamp further has a fastener operatively extending between the blocks.

Abstract: A method is provided for machining a slot having a slot base in a turbine engine rotor disk. The method includes forming a preform slot that extends along a longitudinal axis through the rotor disk. The preform slot includes a preform slot endwall connected laterally between a first slot sidewall and a second slot sidewall. A cutting tool is rotated about its axial centerline, which is substantially parallel to the longitudinal axis. The preform slot endwall is machined with the rotating cutting tool to form the slot base.

Abstract: A turbine bucket for use with a turbine engine. The turbine bucket includes an airfoil that extends between a root end and a tip end. The airfoil includes an outer wall that defines a cavity that extends from the root end to the tip end. The outer wall includes a first ceramic matrix composite (CMC) substrate that extends a first distance from the root end to the tip end. An inner wall is positioned within the cavity. The inner wall includes a second CMC substrate that extends a second distance from the root end towards the tip end that is different than the first distance.

Abstract: One or more support elements radially extend through one or more openings defined in a turbine engine casing and are configured to centralize and at least partially support a rotor assembly of the engine during an engine disassembly or assembly procedure. The support elements are configured to transfer any rotor assembly weight loads to an engine casing while a bearing support of the rotor assembly is absent or removed.

Abstract: A rotor assembly for a gas turbine engine including a circumferential array of regularly spaced apart features provided on the disk which are each configured for receiving a balancing weight. The circumferential array of features has an angular orientation relative to the circumferential array of blades such that the features are each located at an at least substantially same offset angle from a stacking line of a respective adjacent blade.

Abstract: A method of manufacturing a multistage exhaust turbocharger having a high-pressure stage turbocharger and a low-pressure stage turbocharger includes providing a high-pressure turbine housing of the high-pressure stage turbocharger, the high-pressure turbine housing being integrally formed with the exhaust manifold by casting or welding, assembling constituent parts of the high-pressure stage turbocharger using the high-pressure turbine housing as a reference, after the assembling constituent parts of the high-pressure stage turbocharger, attaching the low-pressure stage turbocharger to the high-pressure stage turbocharger by attaching a low-pressure turbine housing to a low-pressure turbine connection flange, and after the attaching the low-pressure stage turbocharger, attaching an air supply channel to connect between a supply air outlet of the low-pressure stage turbocharger and a supply air inlet of the high-pressure stage turbocharger, wherein the low-pressure turbocharger is disposed below the high-press