Abstract: A method includes removing a casting shell and core from a cast component, which may be a gas turbine blade. The method further includes utilizing a focused removal technique, such as a water jet or laser drill, to remove a portion of a virtual pattern cast (VPC) shell from the cast component. The cast component is then exposed to a leaching solution and high pressure water wash to remove an internal core material and a portion of the VPC shell remainder from the cast component. The method further includes exposing the cast component to a high agitation leaching solution and to the high pressure water wash for a minimal time. An electroless nickel-boron coating is then applied to the cast component, and an electrolytic palladium coating is further applied to the cast component. The cast component is further exposed to a high agitation leaching solution for an extended period.

Abstract: An aircraft may have a heat generating component and an engine, at least one of which generates a heat load, and a thermal management system to cool the heat load. The engine may have a duct and an engine fan configured to draw an inlet air stream into an inlet portion of the duct, where at least a portion of the inlet air stream may be used as an engine air stream. The thermal management system may include a cooling circuit configured to circulate a fluid through the heat load such that at least a portion of it may be transferred to the fluid, a heat exchanger configured to enable heat transfer between the fluid and a cooling air stream, and a pumping device. The pumping device may be configured to draw the cooling air stream through the heat exchanger and into a portion of the engine air stream.

Abstract: A gas turbine engine includes a compressor assembly that is rotationally coupled to a shaft, the compressor assembly having a centrifugal impeller and a shroud covering a bladed portion of the centrifugal impeller. The compressor assembly includes a diffuser that is attached to the shroud via a pair of flanges, the diffuser including a strut that is mounted through an aft-extending leg to a base of an intercase. A sealing assembly is attached to the diffuser and is attachable to a transition duct that is positioned to receive air from the diffuser. The sealing assembly is configured to prevent air from passing through the sealing assembly while allowing relative motion to occur between the transition duct and the diffuser.

Abstract: A gas turbine engine includes an inner shaft extending axially along the gas turbine engine, a plurality of disks extending radially inwardly and toward the inner shaft, at least one hole in at least one of the plurality of disks, and an obstruction positioned between the inner shaft and an end of the disk having the at least one hole, such that a bore flow that flows along an axial length of the inner shaft is obstructed from flowing along the shaft by the obstruction, and forced to flow radially outward from the obstruction, through the at least one hole, and radially inward toward the inner shaft.

Abstract: One embodiment of the present disclosure is a unique gas turbine engine having an electrical machine for generating electrical power. Other embodiments include other unique gas turbine engines having electrical machines for generating electrical power. Still other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines having electrical machines for generating electrical power.

Abstract: One embodiment of the present disclosure is a unique particle separator. Another embodiment is a unique aircraft. Another embodiment is a unique inertial particle separator. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for aircraft, engines and particle separators for aircraft and engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present disclosure will become apparent from the description and figures provided herewith.

Abstract: One embodiment of the present invention is a unique aircraft. Another embodiment is a unique external pod for an aircraft. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for aircraft and external pods for aircraft. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: One embodiment of the present invention is an integrated aero-engine flowpath structure. Another embodiment is a method of manufacturing integrated aero-engine flowpath structure. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for aero-engine flowpath structures. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: In one embodiment, a nozzle of a gas turbine engine may be provided having a coanda injector and a fluidic injector which operate together to provide for a change in exhaust flow direction. The fluidic injector may be coincident with or downstream of the coanda injector and both may be used in high pressure ratio operations of the nozzle. The fluidic injector may be positioned opposite the coanda injector and, when activated, may provide for a region of separated flow on the same side of the nozzle as the fluidic injector. The coanda injector may provide additional momentum to an exhaust flow flowing through the nozzle and may encourage the flow to stay attached on the coanda injector side of the nozzle.

Abstract: One embodiment of the present invention is a unique flight vehicle. Another embodiment is a unique propulsion system. Another embodiment is a unique thrust vectoring system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for flight vehicles, propulsion systems and thrust vectoring systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: A discharge nozzle is provided having a centerbody and a conduit enclosing the centerbody and is operable to receive exhaust flow from a turbine discharge of a gas turbine engine. In one form the centerbody is symmetric and extends along the flow direction, rising away from a reference axis to a maximum position before falling and converging on the reference axis. A set of rings is disposed between the centerbody and the conduit and are positioned between the centerbody and the conduit.

Abstract: The present invention provides in one embodiment a unique gas turbine engine. Another embodiment is a unique gas turbine engine combustion system. Still another embodiment is a unique gas turbine engine combustor. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine engine combustion systems and combustors.

Abstract: A fastener is disclosed including a threaded male fastener and a shear bushing for coupling thin walled structures together. The shear bushing having an internal through opening and a shear ring constructed to project across an interface of adjacent walls is operable for absorbing shear loads transmitted into the fastener from a fastened structure. The male fastener includes a threaded portion extending through the shear ring for engaging a threaded female receiver.

Abstract: One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique cooling system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for cooling one or more objects of cooling. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: One embodiment of the present invention is a damping system for rotating machinery such as gas turbine engines. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for damping systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique gas turbine engine main engine rotor. Still another embodiment is a unique method for assembling a gas turbine engine main engine rotor. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine rotors. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: A particle separator for a gas turbine engine is disclosed. The particle separator includes flow dividers operable to divide flow in a gas turbine engine particle separator and flow scavengers operable to scavenge flow in a gas turbine engine particle separator.

Abstract: The present invention provides engine case systems for gas turbine engines. Embodiments of the engine case systems include a plurality of ring cases and a tie-bolt case. The ring cases may house blades and vanes for the compressor and/or for the turbine. The ring cases may be stacked in series, and may be clamped together by a tie-bolt case.

Abstract: A diffuser is provided having a number of diffuser channels defined by a top wall, bottom wall, and two sidewalls. The sidewalls can be constructed according to an involute of a circle which provides constant interwall distance between the sidewalls along the length of the diffuser channel. The top and bottom walls, however, diverge from one another along the length of the diffuser channel to provide a desired diffusion.

Abstract: The present invention provides a variable pressure ratio gas turbine. In an embodiment, the engine is configured to selectively bypass all or a portion of a high pressure compressor and a high pressure turbine. In an embodiment, during bypass operation, all or a portion of the flow that would otherwise pass through the high pressure compressor and the high pressure turbine may be routed through a respective compressor bypass duct and turbine bypass duct.