Abstract: A gas turbine engine includes a blade track, a support assembly, and a seal system. The blade track is arranged around a central axis of the gas turbine engine to direct gasses through the gas turbine engine. The support assembly is arranged around the blade track to support the blade track relative to the central axis. The seal system is configured to block fluid communication between pressurized cavities formed in the support assembly.

Abstract: A turbofan engine has a fan portion in fluid communication with a core stream and a bypass stream of air separated by splitters disposed both upstream and downstream of the fan portion. A fluid passage is defined between the splitters. The turbofan engine has a plurality of high pressure fluid jets originating from a trailing edge of the upstream splitter, the jets restricting the migration of the core stream into the bypass stream through the fluid passage.

Abstract: Methods are provided for creating a fluidic barrier between the core stream and the bypass stream in a turbofan engine. A method comprises compressing the bypass and core streams with a fan between an upstream splitter and a downstream splitter which divides the bypass and core streams, and imparting a first momentum into the air stream proximate the fan in a region between the core and bypass streams and the upstream and downstream splitters to form a fluid barrier, wherein the first momentum of the air stream in the region is higher than a second momentum of the air stream adjacent the fluid barrier.

Abstract: A turbofan engine has a fan portion in fluid communication with a core stream and a bypass stream of air separated by splitters disposed both upstream and downstream of the fan portion. A fluid passage is defined between the splitters. The turbofan engine has a plurality of surface interruptions on a core side surface of the upstream splitter and a plurality of vortices originating from the surface interruptions, thereby restricting the migration of the core stream into the bypass stream through the fluid passage.

Abstract: A turbofan engine has a fan portion in fluid communication with a core stream and a bypass stream of air separated by splitters disposed both upstream and downstream of the fan portion. A fluid passage is defined between the splitters. The turbofan engine has a plurality of vortex generators, each of the vortex generators positioned on the leading edge of a respective fan blade proximate the upstream splitter and the core stream restricting the migration of the core stream into the bypass stream through the fluid passage.

Abstract: A system for directing the flow of a fluid and controlling the rate of flow of the fluid. The system comprises a channel for directing the flow of the fluid; at least a pair of articulating vanes positioned within the channel for controlling the flow rate of the fluid within the channel; and a linkage between the vanes coupling the articulation of each of the vanes to the other of the vanes, wherein each vane imparts a force on the linkage when the relative angle of attack is greater than zero, wherein the force imparted on the linkage by one of the vanes is at least partially cancelled by the force imparted on the linkage by the other of the vanes during the articulation of the vanes.

Abstract: An integrated propulsion system comprising at least two gas turbine engines, at least one fan, and a transmission assembly coupling the at least two gas turbine engines to the at least one fan wherein the at least two gas turbine engines are disposed within a main body of an airframe comprising the main body and a pair of wings, and wherein the number of gas turbine engines is greater than the number of fans.

Abstract: A turbomachine comprises a hub and a plurality of blade elements. Each blade element comprises a blade, a platform, and a tang. The plurality of blade elements are arranged circumferentially around the hub, each interlocking or affixed with an adjacent blade element and retained in position by the hub. Each blade elements formed from a single stamped blank to provide an inexpensive method of manufacture, for low cost turbomachinery.

Abstract: A propulsive force imparted to an object is vectored using rotatable members arranged in one or more arrays disposed in the path of a fluid ejected by a fluid accelerator unit, such as air ejected by a fan driven by a gas turbine engine. The propulsive force is vectored by changing the rotation of one or more of the rotatable members.

Abstract: A composite turbomachine comprises a hub comprised of fiber and resin and a plurality of blade assemblies. Each blade assembly comprises a blade, a base, and a tang. The plurality of blade assemblies are arranged circumferentially around the hub, each interlocking with an adjacent blade assembly and retained in position by the hub and a band overwrapping the respective tang of each of the plurality of blade assemblies.

Abstract: A compressor shroud assembly in a turbine engine having a dynamically moveable shroud for encasing a rotor segment. The rotor segment comprises a bladed disc. The compressor shroud assembly maintains a clearance gap between the shroud and the blade tips of the bladed disc. The assembly comprises a static compressor casing, a shroud mounted to the casing, and an actuator mounted to the casing. The shroud comprises a flowpath boundary member spaced radially inward from the casing, the flowpath boundary member being moveable relative to the casing in a radial direction. The actuator is coupled to the shroud for effecting the movement of the flowpath boundary member.

Abstract: The present disclosure provides a multiple use threaded fastener having a plurality of redundant locking features. The fastener includes a threaded hollow lock bolt having a tapered internal wall constructed to threadingly engage a threaded aperture in a housing. A threaded lock stud having a tapered external wall is constructed to fit at least partially within the hollow lock bolt and threadingly engage the threaded aperture of the housing and engage the tapered wall of the hollow lock bolt.

Abstract: A turbofan engine has a fan portion in fluid communication with a core stream and a bypass stream of air separated by splitters disposed both upstream and downstream of the fan portion. A fluid passage is defined between the splitters. The turbofan engine has a plurality of high pressure fluid jets originating from the low pressure side of the fan blades, the jets restricting the migration of the core stream into the bypass stream through the fluid passage.

Abstract: An expandable exhaust cone assembly is described which is able to move from a collapsed position to an expanded position.

Abstract: A compressor for use in a gas turbine engine is disclosed herein. The compressor includes a case, vanes, and a vane actuation system. The vanes are arranged circumferentially adjacent to one another about a central axis inside the case and each one of the vanes is configured for rotation about a vane axis generally perpendicular to the central axis. The vane actuation system is configured to cause rotation of at least one of the vanes about the vane axis.

Abstract: A gas turbine engine includes a turbine shroud assembly arranged around a turbine wheel assembly. The turbine shroud assembly includes an annular turbine case, a blade track, and a forward case. The turbine case is arranged around a central axis of the turbine assembly. The blade track is located radially between the turbine case and the central axis. The forward case is coupled to the turbine case and is arranged to block axial movement of the blade track relative to the turbine case.

Abstract: An axial compressor comprises a plurality of compressor stages positioned axially adjacent one another within a casing. Each of the plurality of compressor stages comprise a rotor segment and a banded stator segment positioned axially adjacent the rotor segment. An annulus is formed between the casing and an outer flowpath ring of the stator segment. A vane of each stator segment comprises a member extending into the annulus.

Abstract: A composite rotatable assembly for an axial flow compressor comprises a spool having a plurality of blade assemblies arranged in stages on the spool and attached thereto by a wound band. Each blade assembly comprises a blade and a base, with the base having a forward tang extending axially forward of a leading edge of the blade and an aft tang extending axially aft of a trailing edge of the blade. The band is wound over at least a portion of the forward and aft tangs of the plurality of blade assemblies to hold the blade assemblies to the spool under centrifugal loading. An abradable layer may be added over the wound band.

Abstract: An axial compressor comprises a plurality of compressor stages positioned axially adjacent within a casing. Each of the plurality of compressor stages comprise a rotor segment and a banded stator segment positioned axially adjacent the rotor segment. An annulus is formed between the casing and an outer flowpath ring of the stator segment. A pathway is provided that establishes an air flowpath between an entry pathway in a first stage and an exit pathway in a second stage.

Abstract: An axial flow compressor comprises a tubular casing which encases a rotatable shaft, a pair of rotor segments coupled to the rotatable shaft and each comprising a bladed disc, and a banded stator segment disposed between the pair of rotor segments and comprising a plurality of stator vanes extending between an outer flowpath ring and an inner flowpath ring. A method of assembling an axial flow compressor comprises installing a rotor segment inside a tubular compressor casing, installing a vane segment adjacent the installed rotor segment, and repeating the steps of installing a rotor segment and vane segment until a desired number of rotor segment and vane segment pairs are installed.