Abstract: An inlet assembly for a high-mach engine includes a gas turbine core, an inlet turbine, and a core-flow director. When in a closed position, the core-flow director forces air to interact with the inlet turbine before entering the gas turbine core.

Abstract: An environmental control system for an aircraft includes a higher pressure tap to be associated with a higher compression location in a main compressor section associated with an aircraft engine, and a lower pressure tap to be associated with a lower pressure location in the main compressor section associated with the aircraft engine. The lower pressure location being at a lower pressure than said higher pressure location. The lower pressure tap communicates to a first passage leading to a downstream outlet, and having a second passage leading into a compressor section of a turbocompressor. The higher pressure tap leads into a turbine section of the turbocompressor such that air in the higher pressure tap drives the turbine section to in turn drive the compressor section of the turbocompressor. A turbine outlet receives airflow exhausted from the turbine section. A compressor outlet receives airflow exhausted from the compressor section.

Abstract: A combustor includes: a combustion liner having a combustion chamber formed therein; and a fuel injector mounted to a top portion of the combustion liner, and including a fuel injection member having a plurality of fuel injection annular portions and an air guide member including a plurality of combustion air annular portions that guide air for combustion. The fuel injection annular portions and the combustion air annular portions are arranged concentrically and alternately. The fuel injector injects fuel and air into the combustion chamber. Each of the fuel injection annular portions includes a plurality of fuel injection holes that are open in a radial direction thereof, and each of the combustion air annular portions includes a plurality of air guide grooves that are open in an axial direction thereof, and guide the air to the fuel jetted from the fuel injection holes.

Abstract: A method and apparatus using a cooler to selectively cool P3 air directed into an impeller rear cavity of a gas turbine engine during engine high power levels, may include connection of the cooler to a low pressure compressor bleed-off valve apparatus that is modulated to be closed during engine high power levels to create a pressure differential over the bleed-off valve apparatus, to drive a stream of low pressure compressor air as a cooling work fluid under such pressure differential to selectively flow through the cooler.

Abstract: Systems and methods for adjusting airflow distortion in a gas turbine engine using a translating inlet assembly are provided. In one embodiment, a core engine of a gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. The compressor section can include an inner flowpath surface. A core casing can enclose the core engine. A forward end of the core casing can include a translating inlet assembly moveable between a first position and a second position. The translating inlet assembly and the inner flowpath surface can together define an inlet to an engine airflow path. A translating inlet assembly can define a first inlet area in the first position and a second inlet area in the second position, the first inlet area being greater than the second inlet area.

Abstract: Fuel used as a coolant in an aircraft can be thermally conditioned for active thermal management of the airframe and engine. The fuel can be thermally conditioned using the residual cooling capacity of a power and thermal module (PTM), providing flexibility of thermal system design, or via a compact engine-mounted turbo cooler, to maximize system efficiency. The fuel can be stored in a thermal reserve tank to provide a missionized heat sink capable serviceable for periodic high heat flux equipment. The cooling and provision of cooled fuel to aircraft components can be intelligently controlled to provide efficient cooling and effectively unlimited ground hold times.

Abstract: A wall assembly that may be for a combustor of a gas turbine engine includes a liner having a hot face that defines a combustion chamber, an opposite cold face, and a plurality of effusion holes. A shell of the assembly is spaced outward from the cold face and includes a plurality of impingement holes each having a centerline orientated substantially normal to the cold face. A plurality of cooling member arrays of the liner each include a first plurality of members that may be pins projecting outward from the cold face to conduct heat out of the liner. Each array is spaced between adjacent effusion holes and is symmetrically orientated about the respective centerline.

Abstract: The jet engine includes an inlet and a combustor. The inlet (11) takes air. The combustor burns fuel with the air. The combustor includes an injector (20). The injector (20) has an opening (31) through which the fuel is injected. The injector (20) includes a self-extinguishing member (32). The self-extinguishing member (32) self-extinguishes with time in a flight so that the injection direction of the fuel is modified.

Abstract: A combustor of a gas turbine engine includes a multiple of liner panels mounted to the support shell, at least one of the multiple of liner panels includes a first impingement cavity that operates at a first pressure and a second impingement cavity that operates at a second pressure different than the first pressure. A method of cooling a wall assembly within a combustor of a gas turbine engine includes directing air through a support shell and a liner panel that defines a first impingement cavity and a second impingement cavity. The first impingement cavity operates at a first pressure and the second impingement cavity operates at a second pressure that is different than the first pressure.

Abstract: Apparatus and methods for heating pre-combustor air in a gas turbine engine are disclosed. In one embodiment, a gas turbine engine configured for heating the pre-combustor air comprises a compressor for pressurizing air received in the gas turbine engine; an electric heater configured to heat the compressed air; a combustor in which the heated compressed air is received, mixed with fuel and ignited for generating combustion gas; and a turbine for extracting energy from the combustion gas. Joule heating may be used to heat the compressed air upstream of the combustor.

Abstract: Hybrid aircraft propulsors having electrically-driven augmentor fans are disclosed. An example apparatus includes a turbofan having a core engine and a ducted fan to be rotated via the core engine. The ducted fan includes a plurality of ducted fan blades arranged circumferentially around the core engine and circumscribed by a nacelle. The example apparatus further includes an augmentor fan having an augmentor hub ring and a plurality of augmentor fan blades. The augmentor fan blades are arranged circumferentially around the augmentor hub ring and project outwardly relative to an outer surface of the nacelle. The augmentor fan is to rotate separately from the ducted fan. The example apparatus further includes an electrical drive to rotate the augmentor hub ring in response to a supply of electrical energy provided to the electrical drive.

Abstract: Systems and methods for generating an oblique shock in a supersonic inlet are disclosed. The system can comprise an inlet with a slot disposed at an oblique angle to the main incoming air stream. High-pressure air can be provided through the slot into the main air stream. The high-pressure air can be introduced at a high enough pressure ratio—i.e., the ratio of pressure of the air stream from the slot to the pressure for the main flow—such that an aerodynamic ramp is created in the main air flow. The aerodynamic ramp, in turn, can cause one or more oblique shock waves to eventually slow the main air stream velocity to a subsonic speed prior to the face of the engine. Systems and methods for controlling the slot pressure ratio to create these shocks are also disclosed.

Abstract: A gas turbine engine having a combustor, turbine and transition duct to channel hot gas from combustor to turbine. The transition duct has an internal surface on which the hot gas impinges causing a varying temperature profile. A thermal barrier coating is located on the internal surface having a first and second thermal barrier coating patch. The first patch having a first thickness located on the internal surface and within a first area subject to a higher temperature than an uncoated part and bounded by a first isotherm of a first temperature. The second patch having a second thickness located on the internal surface and within a second area subject to a higher temperature than the uncoated part and bounded by a second isotherm of a second temperature. The second temperature is higher than the first temperature and the second thickness is thicker than the first thickness.

Abstract: A fuel injector for injecting alternate fuels having a different energy density in a gas turbine is provided. A first fuel supply channel (18) may be fluidly coupled to a radial passage (22) in a plurality of vanes (20) that branches into passages (24) (e.g., axial passages) to inject a first fuel without jet in cross-flow injection. This may be effective to reduce flashback in fuels having a relatively high flame speed. A mixer (30) with lobes (32) for injection of a second fuel may be arranged at the downstream end of a fuel delivery tube (12). A fuel-routing structure (38) may be configured to route the second fuel within a respective lobe so that fuel injection of the second fuel takes place radially outwardly relative to a central region of the mixer. This may be conducive to an improved (e.g., a relatively more uniform) mixing of air and fuel.

Abstract: An apparatus and method for a turbine engine having an engine core. A casing can contain a compressor section, a combustor section and a turbine section in axial flow arrangement. An air conduit can extend at least partially between the compressor section and the turbine section. A compliant interface can include a nut and a biasing device mount the air conduit to the casing.

Abstract: A combustor of a jet engine includes a fuel injector, an igniter for igniting a gas mixture of air and fuel, and a flame holder. The igniter is disposed in the flame holder. After an activation of the igniter, the igniter disappears, and a space after the disappearance functions as a flame-holding space.

Abstract: Wind-funneling systems for gas turbines are disclosed. Air travels through a wind funnel where it is compressed, and then flows into a gas turbine that is fueled by a hydrocarbon fuel source such as natural gas. The wind funnels have a controlled volume with an opening facing the wind that concentrates air power and energy and directs the force into a constricted outlet that feeds into the gas turbine. Compressed air from the wind funnel may enter the front compressor section of the gas turbine at relatively high density and force. As a result, the gas turbine does not have to use as much energy to pull the air in, which creates fuel savings. The compressed air from the wind funnel then flows to the combustion section of the gas turbine where oxygen from the wind-compressed air is used to combust the hydrocarbon fuel supplied to the gas turbine.

Abstract: A combustor of a gas turbine engine includes a combustor shell having a diffuser side facing a diffuser chamber and a combustor side facing a combustor chamber. The combustor may include a first combustor panel coupled to the combustor side of the combustor shell and a second combustor panel coupled to the combustor side of the combustor shell. A gap may be defined between the first combustor panel and the second combustor panel and the combustor shell may include a gap impingement hole that is directly open to and is configured to deliver cooling air directly to the gap. In various embodiments, the combustor further includes a throttle plate coupled to the diffuser side of the combustor shell.

Abstract: A plate heat exchanger comprising a plurality of plates having a plane peripheral zone, an inner zone having sinusoidal undulations and two chimneys positioned at two opposite corners of the plates. Modules are formed by assembling two plates that make contact via the undulation troughs and the peripheral zones. The modules are stacked so as to make contact via the inlet and outlet chimneys. Each module may thus deform independently, in particular at the undulation troughs and ridges without transmitting stress to the other modules of the heat exchanger. In addition, the heat exchanger may comprise a tie rod in each inlet and outlet pipe so as to withstand the static pressure of the fluids flowing in the pipes.

Abstract: The present disclosure relates generally to turbomachinery in which a rotating component is disposed adjacent a counter-rotating component in order to achieve a relative rotational velocity that is higher than would be achieved with a rotating component disposed adjacent to a stationary component.