Abstract: Thrust reversers and methods to provide reverse thrust are disclosed. An example thrust reverser includes a cascade to direct air from an interior volume of an engine nacelle to generate reverse thrust with respect to a direction of travel, and an air extractor to extract a portion of air approaching the cascade from within the interior volume.

Abstract: A multipoint fuel injection system includes a plurality of fuel manifolds. Each manifold is in fluid communication with a plurality of injectors arranged circumferentially about a longitudinal axis for multipoint fuel injection. The injectors of separate respective manifolds are spaced radially apart from one another for separate radial staging of fuel flow to each respective manifold.

Abstract: A system includes a combustor cap assembly for a multi-tube fuel nozzle. The combustor cap assembly includes a support structure defining an interior volume configured to receive an air flow, a plurality of mixing tubes disposed within the interior volume, wherein each of the plurality of mixing tubes comprises a respective fuel injector and is individually removable from the combustor cap assembly, an air distributor disposed within the interior volume and configured to distribute the air flow received by the interior volume to each of the plurality of mixing tubes, and a combustor cap removably coupled to the support structure.

Abstract: A fuel nozzle in a combustion turbine engine that includes: a fuel plenum defined between an circumferentially extending shroud and axially by a forward tube-sheet and an aft tube-sheet; and a mixing-tube that extends across the fuel plenum that defines a passageway connecting an inlet formed through the forward tube-sheet and an outlet formed through the aft tube-sheet, the mixing-tube comprising one or more fuel ports that fluidly communicate with the fuel plenum. The mixing-tube may include grooves on an outer surface, and be attached to the forward tube-sheet by a connection having a fail-safe leakage path.

Abstract: A hanger for suspending a liner within an exhaust duct of a gas turbine engine comprises a bracket, a disk and a rod. The bracket is for connecting to an exhaust duct liner of the gas turbine exhaust system. The disk is slidably engaged with the bracket. The rod is for connecting to an exhaust duct of the gas turbine engine. The rod is pivotably connected to the disk.

Abstract: A gas turbine engine fuel supply system includes a primary gear pump and a secondary gear pump. The primary gear pump always actively delivers fuel to the downstream fuel system, and is sized to supply 100% of the burn flow needed at a select low demand condition. The secondary gear pump is sized to make up the remainder of the flow at high demand conditions, and actively delivers fuel to the downstream fuel system only during those conditions. To supply discharge fuel pressures in excess of gear pump capability, a supercharger pump is disposed upstream of the primary and secondary gear pumps. The supercharger pump is preferably activated only during high demand conditions as an additional energy conservation measure.

Abstract: A system includes a gas turbine combustor configured to combust a fuel and an oxidant, such as O2 and O2 mixtures. The system also includes an aerodynamic peg disposed in the gas turbine combustor. The aerodynamic peg includes a first passage configured to convey a first fluid into the gas turbine combustor and a second passage configured to convey a second fluid into the gas turbine combustor. The first fluid and second fluid are different from one another.

Abstract: This invention is a packaged propellant air-induced variable thrust rocket engine that has a vast number of uses and applications for this invention. The primary purpose of the device described here is to provide a light weight, torque and vibration free thrust generator for the propulsion of aircraft. This device will facilitate the fabrication of very light weight aircraft because of the lack these forces. This device can also be used anywhere high velocity air flow and or the resulting thrust is needed. The invention uses aerodynamic principles to compress and accelerate the incoming air, prior to it being heated and accelerated by a short duration burst of thermal and kinetic energy from discrete packets of a mixture of oxidizable fuels. The heated and accelerated air then expands as it travels thru the device providing thrust.

Abstract: An air intake system suitable for a supersonic vehicle is disclosed. The system includes a channel comprising an inlet and a side wall and a plenum coupled to the side wall. The plenum is configured to accept a flow of coolant. In certain embodiments, the coolant is the waste coolant from an on-board electronics cooling system. The system also includes a porous region in the side wall configured to allow a flow of bleed air from the channel through the porous region of the side wall into the plenum so as to aid the transition to supersonic flow. In certain embodiments, the flow of the bleed air is reduced at supersonic speeds by pressurization of the plenum with the coolant.

Abstract: A combustor section for a gas turbine engine includes a combustor vane which extends at least partially into a combustion chamber.

Abstract: An aeronautical engine including a fuel pumping device including a high pressure fuel pump including an inlet connected to a low pressure fuel conduit and an outlet connected to a main circuit for feeding high pressure fuel, an electric device starting the engine, and a device cooling the electric starting device connected to the pumping device to ensure cooling by circulation of fuel. The cooling device is supplied with fuel by a pump that includes an inlet connected to the pumping device, upstream from the high pressure pump, and that is driven by an electric motor independently of the high pressure pump.

Abstract: A hanger for suspending a liner within an exhaust duct of a gas turbine engine comprises a bracket, a washer, a rod and a cap. The bracket is for connection to an exhaust duct liner of a gas turbine exhaust system. The washer is connected to the bracket at a slip joint. The rod is connected to the washer at a ball joint. The cap for is for connecting to an exhaust duct of the gas turbine exhaust system. The cap is connected to the rod.

Abstract: Certain embodiments include a resonator for the head end of a turbine combustor. The resonator is partially defined by first and second plates disposed about a fuel nozzle in the head end of the turbine combustor. The resonator is further configured to receive an air flow which extends through the resonator and into the fuel nozzle.

Abstract: A method of managing aircraft exhaust includes providing hot air at a hot air mass flow rate, providing cold air at a cold air mass flow rate, and mixing the hot air and the cold air at a variable hot air mass flow rate to cold air mass flow rate ratio, wherein the variable hot air mass flow rate to cold air mass flow rate ratio is selectively maintained independent of at least one of (1) a variation in the hot air mass flow rate and (2) a variation in a translational speed of the aircraft.

Abstract: A splitter apparatus for a gas turbine engine includes: a splitter including: an annular outer wall which defines a convex-curved leading edge at a forward end thereof; an annular floorplate positioned radially inboard of the outer wall; and an annular first bulkhead spanning between the outer wall and the floorplate. The outer wall, the floorplate, and the bulkhead collectively define an annular splitter plenum positioned adjacent the leading edge of the outer wall. At least one exhaust passage formed in the outer wall extends past the floorplate and communicates with the exterior of the splitter. At least one jumper tube assembly passes through the first bulkhead, each configured to pass air flow from the exterior of the splitter into the plenum.

Abstract: A splitter apparatus for a gas turbine engine includes: a splitter including: an annular outer wall which defines a convex-curved leading edge at a forward end thereof; an annular floorplate positioned radially inboard of the outer wall; and an annular first bulkhead spanning between the outer wall and the floorplate. The outer wall, the floorplate, and the bulkhead collectively define an annular splitter plenum positioned adjacent the leading edge of the outer wall. At least one exhaust passage formed in the outer wall extends past the floorplate and communicates with the exterior of the splitter. At least one jumper tube assembly passes through the first bulkhead, each configured to pass air flow from the exterior of the splitter into the plenum.

Abstract: A flow divider valve for a gas turbine engine fuel supply system is provided. The flow divider valve is configured, whenever fuel pressure at its inlet port is between a first predetermined pressure and a second predetermined pressure, to supply the fuel to a start nozzle manifold outlet port at a substantially constant flow rate, and to direct any fuel flow in excess of the substantially constant flow rate to a run nozzle manifold outlet port.

Abstract: An aircraft jet engine includes an exhaust-gas nozzle having a device configured to blow out an exhaust gas in a pulsating manner into an exhaust-gas stream so as to reduce noise. The exhaust-gas nozzle includes openings distributed along a circumference of the exhaust-gas nozzle and disposed upstream from a nozzle outlet. The openings communicate with the device.

Abstract: A fuel nozzle is provided and includes a nozzle body defining first and second interior regions for providing a supply of first and second fluids, a collar defining a third interior region and radial slots permitting radial ingress of a third fluid to the third interior region and a nozzle tip interposed between the nozzle body and the collar. The nozzle tip defines an annular slot, first discrete passageways by which the first fluid is communicated from the first interior region to the annular slot, second discrete passageways by which the first fluid is communicated from the annular slot to the radial slots, and third discrete passageways by which the second fluid is communicated from the second interior region to the radial slots.

Abstract: Systems and methods for driving an oil cooling fan (36) of a gas turbine engine (10) during different modes of operation of the gas turbine engine (10) are described. A system may include a coupling device (40) configured to: transmit motive power from a power turbine shaft (22) of the gas turbine engine (10) to the oil cooling fan (36) during a first mode of operation where the power turbine shaft (22) is turning, and to decouple the oil cooling fan (36) from the power turbine shaft (22) during a second mode of operation where the power turbine shaft (22) is prevented from turning. An alternate source (42) of motive power may be configured to drive the oil cooling fan (36) during the second mode of operation.