Abstract: A gas turbine engine includes: a turbomachine including a combustion section and an exhaust assembly arranged in serial flow order; and a fuel system. The fuel system includes: a hydrogen fuel tank for holding a hydrogen fuel in a liquid phase; and a hydrogen delivery assembly in thermal communication with the exhaust assembly. The hydrogen delivery assembly includes: a liquid hydrogen line in fluid communication with the hydrogen fuel tank; a gaseous hydrogen line in fluid communication with the combustion section of the gas turbine engine; and a heat exchanger tubing positioned within the exhaust assembly downstream of the liquid hydrogen line and upstream of the gaseous hydrogen line.

Abstract: Provided are fluid exchange apparatuses and methods of exchanging fluids between streams. Exemplary fluid exchange apparatus includes a first interleaved pathway with a plurality of first passages, and a second interleaved pathway with a plurality of second passages, in which the plurality of first passages and the plurality of second passages interleave with one another. Exemplary methods include directing a first fluid through a first interleaved pathway of a fluid exchange apparatus and directing a second fluid through a second interleaved pathway of a fluid exchange apparatus. The first fluid may flow from an upstream portion of a first duct into the first interleaved pathway and may discharge into a downstream portion of the second duct. The second fluid may flow from an upstream portion of the second duct into the second interleaved pathway and may discharge into a downstream portion of the first duct.

Abstract: A gas turbine engine includes a casing, a rotor assembly rotatably mounted within the casing, the rotor assembly comprising a forward rotor portion defining a forward rotor cavity and an aft rotor portion defining an aft rotor cavity, the aft rotor cavity being fluidly isolated from the forward rotor cavity, and a cooling system. The cooling system includes a forward cooling fluid supply, a forward supply line providing fluid communication between the forward cooling fluid supply and the forward rotor cavity, an aft cooling fluid supply, an aft supply line providing fluid communication between the aft cooling fluid supply and the aft rotor cavity, and a flow control system fluidly coupled to the forward supply line and the aft supply line for regulating a flow of forward cooling fluid and a flow of aft cooling fluid.

Abstract: A gas turbine engine includes a fan positioned at an engine central longitudinal axis, and a fan drive turbine located at the engine central longitudinal axis and configured to drive rotation of the fan. A gas generator is non-coaxial with the fan drive turbine and operably connected to the fan drive turbine such that exhaust from the gas generator drives rotation of the fan drive turbine. An auxiliary power core is located at the engine central longitudinal axis, and one or more bleed passages connect the gas generator and the auxiliary power core. The one or more bleed passages are configured to selectably combine a bleed airflow from the gas generator and an auxiliary core airflow at the auxiliary power core to direct the combined airflow to the fan drive turbine to increase output of the fan drive turbine.

Abstract: A power supply device that supplies power to a power load of a flying body, comprises: a power generation unit; a hollow housing including a reserving portion configured to reserve a fuel of the power generation unit, and a storage portion configured to store the power generation unit; and a connecting portion configured to connect the housing to an airframe of the flying body. The housing has a shape long in a front-and-rear direction of the flying body, and is arranged outside the airframe, and the reserving portion and the storage portion are arranged in a longitudinal direction of the housing.

Abstract: A reverse flow gas turbine engine has a low pressure (LP) spool and a high pressure (HP) spool arranged sequentially in an axial direction. The LP spool comprises an LP compressor disposed forward of an LP turbine and drivingly connected thereto via an LP compressor gear train. The HP spool comprises an HP compressor in flow communication with the LP compressor, and an HP turbine disposed forward of the HP compressor and drivingly connected thereto via an HP shaft.

Abstract: An oil pipe assembly for a gas turbine engine. The oil pipe assembly includes a first pipe that defines a first fluid passage between an oil supply and a bearing chamber, and a second pipe that houses the first pipe and defines a second fluid passage between the first pipe and the second pipe that is supplied with cooling air. The oil pipe assembly also includes a restrictor that extends from the second pipe and restricts the passage of fluid from the second fluid passage before it flows into a breather. Pressure and temperature sensors) are located adjacent the restrictor to detect and measure changes in air pressure and air temperature adjacent the restrictor from which a controller identifies whether a leak has occurred in the first pipe or the second pipe. A method for detecting a leak in the oil pipe assembly, and a gas turbine are also described.

Abstract: An additively manufactured attritable engine includes a compressor section, a combustion section, a turbine section, and an engine case wall, which surrounds the compressor section, the combustion section, and the turbine section. The engine case wall includes a first cavity embedded in the engine case wall that defines an injector that is in fluid communication with the combustion section. The engine case wall includes a second cavity embedded within the engine case wall and defines a fuel feed passage that is in thermal communication through the exterior surface of the engine case wall.

Abstract: An aircraft with an integrated boundary layer ingesting propulsion having a mechanically-distributed propulsion system. The mechanically-distributed propulsion system may include an engine to generate a mechanical drive power, a drive shaft, a direction-reversing transmission, and a propulsor fan. The drive shaft may be operatively coupled to the engine to receive the mechanical drive power. The direction-reversing transmission may have a first rotating shaft and a second rotating shaft, the first rotating shaft operatively coupled to the drive shaft to receive the mechanical drive power, which is configured to redirect the mechanical drive power received at the first rotating shaft from a first direction to face a second direction at the second rotating shaft. The propulsor fan may be coupled to the second rotating shaft to convert the mechanical drive power into thrust.

Abstract: A method of securing a fuel injector seal assembly includes placing a fuel nozzle within a bore of a seal, and selectively limiting withdrawal of the fuel nozzle from the bore based on the circumferential position of the fuel nozzle relative to the bore.

Abstract: An additive manufactured article for a gas turbine having a body with two lateral surfaces elongated in a first direction; at least a nested duct housed within the lateral surfaces and elongated in the first direction; a structure so that the nested duct is structurally connected to an attachment within the lateral surfaces, wherein at least the body, the duct and the structure are manufactured by an additive manufacturing process and the structure includes an array of ribs attached to the duct in order to compensate differential elongation along the first direction of the duct with respect to the attachment of the ribs by flexural deformation.

Abstract: In accordance with one aspect of the disclosure, a stream diverter for a gas turbine engine is disclosed. The stream diverter may include a first air duct, a second air duct, a third air duct, and a door operatively associated with the second and third air ducts of the gas turbine engine. The door may have at least an open position allowing air from the second air duct to flow into the third air duct and a closed position preventing air from flowing between the ducts.

Abstract: The present disclosure relates to a method of operating a supersonic aircraft comprising the steps of: at takeoff, positioning a slidable plug-cowl assembly disposed within a nozzle and behind an engine in an aft position such that a front surface of a plug is aft of an exit plane of a cowl, to thereby reduce noise during takeoff and maintain engine efficiency; and after takeoff, re-positioning the slidable plug-cowl assembly to a forward position such that the front surface of the plug is not disposed aft the exit plane of the cowl.

Abstract: A sealing device is coupled to a first structure and is configured to seal a gap between the first structure and a second structure. The sealing device includes a translatable cartridge including a first wall and a second wall parallel to each other and a third wall perpendicular to and coupled to the first and second walls. A force applying device is configured to apply force on the cartridge in a direction toward the second structure such that the cartridge is exerted against the second structure. A ceramic insulation layer disposed at the third wall such that the ceramic insulation layer is between the third wall and the second structure, and an abrasion layer disposed on the ceramic insulation layer.

Abstract: A gas turbine configured to prevent eccentricity of a rotor (14) due to heat is installed with a strut (23), an outer diffuser (24), an inner diffuser (25), a strut cover (26), and a partition wall (28), wherein the gas turbine includes an inflow hole (31) for cooling air (W), a first flow passage (R1) formed between a casing wall (21) and the outer diffuser (24), a second flow passage (R2) formed between the strut (23) and the strut cover (26), a third flow passage (R3) formed between the inner diffuser (25) and the partition wall (28), and an outflow hole (51) installed in the inner diffuser (25).

Abstract: A gas turbine engine for mounting under a wing of an aircraft has a propulsor that rotates on a first axis, and an engine core including a compressor section, a combustor section, and a turbine section, with the turbine section being closer to the propulsor than the compressor section. The engine core is aerodynamically connected to the propulsor and has a second axis. A nacelle is positioned around the propulsor and engine core. The nacelle is attached to the wing of the aircraft. A downstream end of the nacelle has at least one pivoting door with an actuation mechanism to pivot the door between a stowed position and a horizontal deployed position in which the door inhibits a flow to provide a thrust reverse of the flow.

Abstract: A system for supplying a working fluid to a combustor includes a combustion chamber and a flow sleeve that circumferentially surrounds at least a portion of the combustion chamber. A tube provides fluid communication for the working fluid to flow through the flow sleeve and into the combustion chamber, wherein the tube comprises an axial centerline. A first set of injectors are circumferentially arranged around the tube and angled radially with respect to the axial centerline of the tube, wherein the first set of injectors provide fluid communication for the working fluid to flow through a wall of the tube.

Abstract: A gas turbine engine having a ramburner is disclosed. The ramburner is disposed downstream of a gas turbine engine combustor and receives an engine exhaust flow from the gas turbine engine combustor. The ramburner also accepts a bypass air. Fuel is injected into the ramburner and a combustion reaction is auto-initiated based upon local gas temperatures. No mechanical flame holders need be used. A slidable valve may be used to vary the amount of bypass air into the ramburner. A movable cowl and a plug nozzle form an exit flow path of the gas turbine engine. The movable cowl can be positioned to vary a throat area and exit area of the gas turbine engine based upon the operation of the ramburner, which may be influenced by the amount of bypass air entering the ramburner.

Abstract: An exemplary igniter assembly for a turbomachine includes an igniter extending along an axis from an igniting end portion to an opposing end portion. The igniting end of the igniter is configured to be received within a bore of a tube. An annular seal restricts flow between the igniting end and the tube.

Abstract: A gas turbine engine augmentor includes at least one fluid based augmentor initiator defining a chamber in flow communication with a source of air and a source of fuel. The chamber includes a plurality of ejection openings in flow communication with an exhaust flowpath. The at least one fluid based augmentor initiator is devoid of any exhaust flowpath protrusions thereby minimizing any pressure drops and loss of thrust during dry work phase of operation. The source of fuel is operable for injecting fuel into the chamber such that at least a portion of the fuel flow is ignited at the plurality of ejection openings to produce a plurality of fuel-rich hot jets radially into the exhaust flowpath.

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

Abstract: An aircraft gas turbine engine has a row of FLADE fan blades disposed radially outwardly of and drivingly connected to a fan in the engine's fan section. The FLADE fan blades extend across a FLADE duct circumscribing the fan section. A two dimensional air discharge passage is in fluid flow communication with the FLADE duct and with FLADE air upstream and downstream discharge slots in a divergent flap of a two dimensional exhaust nozzle. A valve fully closes the upstream slot when the downstream slot is fully opened and fully opens the upstream slot when the downstream slot is fully closed. The upstream and downstream slots may be located upstream and downstream respectfully of a nozzle discharge area in the nozzle. A sliding deck slides aft or down to open upstream slot and close downstream slot and slides forward or up to close upstream slot and open downstream slot.

Abstract: An augmentor for a gas turbine engine includes an augmentor spray bar with a spray bar outlet, and an augmentor spray bar tip. The spray bar tip includes a tip body, a tip bushing and a tip support strut. The tip body includes a first flow passage extending therethrough between a tip inlet and a tip outlet, wherein the tip inlet is connected to the spray bar outlet. The tip bushing includes a bushing bore. The tip body extends through the bushing bore defining a second flow passage between the tip body and the tip bushing. The tip support strut connects the tip body to the tip bushing across the second flow passage.

Abstract: This invention relates to devices for discharging high pressure exhaust, and in particular to pulse detonation engines. More specifically, the invention describes a rotary pulse detonation engine having a rotary valve system. The rotary valve includes a generally-triangular rotor having rotor tips within a rotor chamber having trochoid inner end surfaces and side surfaces. The rotor defines three working chambers defined by the rotor tips contacting the rotor surfaces. In operation, the rotor tips move in a circumferential direction around the rotor chamber as the rotor spins. During operation, each of the working chambers will sequentially pass through an intake interval, compression interval, expansion interval, and an exhaust interval to create and detonate compressed fuel air mixtures for effective release to an exhaust chamber and nozzle thereby creating a pulsed detonation sequence.

Abstract: A protocol for assembling a fuel shut off pin valve assembly for a gas turbine engine. The protocol outlines a specific sequence of events in order to ensure failsafe incorporation of a fuel shut off valve assembly within the low pressure turbine area and specifically within the engine casing of the gas turbine.

Abstract: A flame-holder device for an afterburner of a turbofan, the turbofan including a first inner annular casing and a second inner annular casing, these defining a passage for a primary flow, and an outer annular casing defining with the first inner annular casing a passage for a secondary flow, including at least one arm-support made of metallic material configured to be fixed to the outer casing by an upper plate and at least one flame-holder arm of a one-piece structure of composite material including two mutually attached walls arranged to define a groove whose profile is approximately a V. The walls support, on their upper part situated in the secondary flow, a fixing mechanism for fixing to the arm-support. The upper parts are approximately planar and are positioned facing each other.

Abstract: A panel for use in a gas turbine engine exhaust case is disclosed. The panel has an airfoil section and a flow diverting structure adjacent a leading edge, wherein the flow diverting structure directs fluid flow into an area of the airfoil that lacks sufficient internal pressure for cooling fluid flow.

Abstract: An exemplary igniter assembly for a turbomachine includes an igniter extending along an axis from an igniting end portion to an opposing end portion. The igniting end of the igniter is configured to be received within a bore of a tube. An annular seal restricts flow between the igniting end and the tube.

Abstract: An example turbomachine assembly includes a plurality of fuel spraybars spaced about a turbine exhaust inner case of the turbomachine. At least one fuel injector assembly is associated with each of the fuel spraybars. The at least one fuel injector assembly may include a fuel nozzle having a fuel delivery conduit. The at least one fuel injector assembly includes a seal member that is biased away from the fuel nozzle in a direction. The fuel nozzle is configured to limit movement of the seal member in the direction.

Abstract: A system and method of reducing noise caused by jet engines is provided. The method comprises the steps of using the afterburner to heat the exhaust gas flow while simultaneously reducing power to the core engine. Together these two operations reduce the pressure of the exhaust gas in the nozzle area while holding the exhaust gas velocity constant, which maintains engine thrust while decreasing engine noise. The method may be supplemented by altering the location of the afterburner flames to create an inverted exhaust velocity profile, thereby decreasing engine noise even further.

Abstract: An afterburner chamber includes a burner ring having housed therein an annular fuel strip that is pierced by orifices and that is mounted inside an annular thermal protection tube, and an ignition spark plug penetrating into the inside of the burner ring. The protection tube and the burner ring have orifices for passing air. The chamber also includes a device for ventilating the orifices of the fuel strip that are situated in the immediate vicinity of the spark plug.

Abstract: A system comprises a bypass duct, a turbine case, a turbine exhaust ring and a splash plate. The turbine case is formed along a radially inner margin of the bypass duct, and the turbine exhaust ring is coaxially disposed within the turbine case. The splash plate extends axially along the turbine exhaust ring, radially spaced between the turbine exhaust ring and the turbine case. There are cooling fluid apertures in the turbine case to provide cooling fluid flow from the bypass duct onto the splash plate, and impingement holes in the splash plate to provide impingement flow onto the turbine exhaust ring.

Abstract: An augmentor for a gas turbine engine includes an augmentor spray bar with a spray bar outlet, and an augmentor spray bar tip. The spray bar tip includes a tip body, a tip bushing and a tip support strut. The tip body includes a first flow passage extending therethrough between a tip inlet and a tip outlet, wherein the tip inlet is connected to the spray bar outlet. The tip bushing includes a bushing bore. The tip body extends through the bushing bore defining a second flow passage between the tip body and the tip bushing. The tip support strut connects the tip body to the tip bushing across the second flow passage.

Abstract: A flame holder for an afterburner duct of a jet engine, including an arm in the form of a gutter forming a cavity, a shield for protecting the cavity of the arm from heat, and an air supply baffle housed in the cavity, is disclosed. The arm, the protective shield, and the air supply baffle are held together by a one-piece spacer shoe.

Abstract: Integrally formed, multi zone fuel manifolds for gas turbine engines and methods of construction are provided. The fuel manifold includes a plurality of annular channel members positioned adjacent each other, aligned and coupled together at the inner and outer peripheral sides thereof, with each pair of channel members defining a separate fluid conduit or passage within the manifold assembly. The fuel manifold can be constructed of any number of stacked and secured channel members to provide a compact multi zone fuel manifold having the desired number of fluid conduits or passages with a minimum number of sealed joints.

Abstract: Methods and apparatus are provided for operating a gas turbine engine. In a first operational mode, the gas turbine engine generates thrust using the propulsion turbine and not the afterburner when it is commanded to generate a thrust between at least a first thrust magnitude and a second thrust magnitude, and generates thrust using the propulsion turbine and the afterburner when it is commanded to generate thrust greater than the second thrust magnitude. In a second operational mode, the gas turbine engine generates thrust using the propulsion turbine and the afterburner when it is commanded to generate a thrust greater than the first thrust magnitude. The steady state thrust-versus-throttle position response has a substantially constant linear slope that is set to be two times the similar slope of the first operational mode.

Abstract: A gas turbine engine augmenter has a gas flowpath. A number of vanes extend into the gas flowpath. A number of augmenter fuel conduits have outlets along at least some of the vanes. At least one burner discharge outlet is along at least one of the vanes for discharging a pilot gas.

Abstract: A gas turbine engine comprises a fan drive gear system, a low spool connected to the fan drive gear system, and a high spool disposed aft of the low spool. The low spool comprises a rearward-flow low pressure compressor disposed aft of the fan drive gear system, and a forward-flow low pressure turbine disposed aft of the low pressure compressor. The high spool comprises a forward-flow high pressure turbine disposed aft of the low pressure turbine, a combustor disposed aft of the high pressure turbine, and a forward-flow high pressure compressor disposed aft of the combustor.

Abstract: A gas turbine engine augmentor includes at least one fluid based augmentor initiator defining a chamber in flow communication with a source of air and a source of fuel. The chamber includes a plurality of ejection openings in flow communication with an exhaust flowpath. The at least one fluid based augmentor initiator is devoid of any exhaust flowpath protrusions thereby minimizing any pressure drops and loss of thrust during dry work phase of operation. The source of fuel is operable for injecting fuel into the chamber such that at least a portion of the fuel flow is ignited at the plurality of ejection openings to produce a plurality of fuel-rich hot jets radially into the exhaust flowpath.

Abstract: A supersonic combustion apparatus including a fixed geometric nozzle having a converging area, throat, and a diverging area, at least one movable combination of a fuel injector and an oxygen injector where the combined fuel injector and the oxygen injector is located within the divergent area of the fixed geometric nozzle, and an exit plane adjacent and downstream to the diverging area. The exit plane Mach speed is varied by heat addition in the diverging area by introduction of a combustible fuel through the fuel injector and oxygen through the oxygen injector and then axially aligning and positioning the combination of the fuel injector and the oxygen injector along the length of the diverging area to obtain a stabilized flame at the exit plane.

Abstract: A cooling system for a turbine exhaust assembly comprises an annular case, a flowpath ring and a splash plate. The flowpath ring is coaxially disposed within the annular case. The splash plate extends axially between the annular case and the flowpath ring. A plurality of cooling fluid apertures is formed in the annular case, in order to provide cooling fluid flow onto the splash plate. A plurality of impingement holes is formed in the splash plate, in order to provide impingement flow onto the flowpath ring.

Abstract: A gas turbine engine augmentor includes an annular trapped vortex cavity pilot having a cavity forward wall, a cavity radially outer wall, and a cavity aft wall, an annular cavity therebetween, and cavity fuel injector tubes operably disposed through the outer wall into the cavity. Circumferentially spaced apart radial flameholders with integral spraybars and/or radial spraybars interdigitated with the radial flameholders radially inwardly into an exhaust flowpath of the augmentor just forward and upstream of the trapped vortex cavity pilot at a radially outer portion of a combustion zone of the exhaust flowpath. An annular trapped dual vortex cavity pilot version is operable for producing trapped dual counter-rotating inner and outer vortices of fuel and air mixtures.

Abstract: A flame holder for an augmentor includes geometry to produce large radial variations in heat release to prevent coupling with modes within the exhaust duct that produce combustor screech. A surface of the flame holder varies axially in a radial direction to produce a radial phase variation in heat released from the flame holder. The radially changing surface produces radially non-uniform and unsteady heat release that does not couple with acoustic modes within the exhaust duct.

Abstract: A device for mounting a dividing wall for separating the main air stream and the bypass air in a bypass turbojet engine afterburner is disclosed. The device includes an attachment unit which attaches the upstream end of the wall to a guide vane casing and a supporting unit which supports the downstream end of the wall provided on flame holder arms of the afterburner radially on the inside of the dividing wall.

Abstract: A device attaching an afterburner to a turbojet casing. The device comprises a transverse plate for attachment to the casing and at least two side brackets for supporting two legs of the afterburner. The brackets and the two legs extend longitudinally. The afterburner includes a flame-arrestor arm with the two legs.

Abstract: A turbine engine trailing edge box for an afterburner has corrugations formed to guide and support a spraybar. The trailing edge box is formed of sheet metal to create corrugations in the walls. The spraybar is supported and guided by the corrugations. The corrugations are shaped and sized to hold the spraybar, but to prevent shifting of the spraybar during engine operation. In addition, the corrugations reinforce the walls of the trailing edge box to provide structural support.

Abstract: A protocol for assembling a fuel shut off pin valve assembly for a turbo fan engine. The protocol outlines a specific sequence of events in order to ensure failsafe incorporation of a fuel shut off valve assembly within the low pressure turbine area and specifically within the engine casing of the turbo fan.

Abstract: An annular trapped vortex cavity assembly segment comprising includes a cavity forward wall, a cavity aft wall, and a cavity radially outer wall therebetween defining a cavity segment therein. A cavity opening extends between the forward and aft walls at a radially inner end of the assembly segment. Radially spaced apart pluralities of air injection first and second holes extend through the forward and aft walls respectively. The segment may include first and second expansion joint features at distal first and second ends respectively of the segment. The segment may include a forward subcomponent including the cavity forward wall attached to an aft subcomponent including the cavity aft wall. The forward and aft subcomponents include forward and aft portions of the cavity radially outer wall respectively. A ring of the segments may be circumferentially disposed about an axis to form an annular segmented vortex cavity assembly.

Abstract: Systems and methods reducing thermo-acoustic coupling of a gas turbine engine augmentors are provided. In this regard, a representative method includes: determining acoustic resonances and heat release phase relationships associated with the augmentor; and determining relative axial positions of a fuel injector and a flame holder of the augmentor that result in reduced thermo-acoustic coupling of the augmentor during operation.

Abstract: A gas turbine engine fueling system includes a number of spray bars having conduits extending through associated vanes. A number of fuel injector nozzles are along each conduit. Each of the nozzles are positioned to discharge an associated fuel stream from one of the sides of the associated vane. A number of wear members are each mounted relative to the associated one of the nozzles for a range of motion relative thereto. The wear members movably cooperate with the associated vane to accommodate operating deflection and/or tolerances of the spray bars and vanes.