Abstract: The invention relates to an air intake of an aircraft turbojet engine nacelle extending along an axis X, in which an air flow circulates from upstream to downstream, the air intake comprising an inner wall and an outer wall which are connected by a leading edge and an inner partition so as to delimit an annular cavity, the air intake comprising means for injecting at least one hot air flow into the inner cavity and at least one ventilation orifice formed in the outer wall in order to allow the hot air flow to escape after heating the inner cavity, the air intake comprising at least one acoustic member positioned in the inner cavity facing the ventilation orifice so as to modify the acoustic resonance frequencies and/or to attenuate the acoustic waves formed in the inner cavity via the ventilation orifice.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter. The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module including a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the fan assembly. Each heat transfer element may be individually and independently fluidly isolated from the remaining heat transfer elements.

Abstract: A compressor bleed apparatus includes: a compressor comprising one or more rotors mounted for rotation about a central axis and enclosed in a compressor casing; a bleed slot passing through the compressor casing; an outer wall defining, in cooperation with the compressor casing, a plenum surrounding the compressor casing; at least one offtake pipe communicating with the plenum; and wherein at least one of the plenum and the bleed slot has a non-axisymmetric structure.

Abstract: An air start unit for starting and servicing jet engines in aircraft and other flying machines, wherein at least one compressor is provided for generating air and a power supply for supplying electrical power consumers. The air start unit drives the compressor with at least one electric motor, with the electric motor drawing the electrical energy for operation from a high-voltage battery. The electric motors then in turn drive the compressor, which generates compressed air from ambient air for the purpose of starting and servicing the jet engines in aircraft and other flying machines. With an appropriate battery capacity, the air start unit can also supply power consumer of these aircraft and other flying machines via the power supply.

Abstract: A combustion chamber for a turbine engine, in particular for an aircraft turbojet engine or turboprop engine. The combustion chamber includes a radially outer annular shroud, a radially inner annular shroud coaxial with the radially outer shroud, and an end wall connecting the radially outer shroud and the radially inner shroud. The combustion chamber further includes a first annular sealing member coaxial with said radially inner and outer shrouds. The first annular sealing member is radially interposed between the end wall and the radially outer shroud.

Abstract: A hydrogen/oxygen stoichiometric combustion turbine system includes: a high-pressure steam turbine (2); a low-pressure steam turbine (3); and a heater (5) disposed between the high-pressure and low-pressure steam turbines. The heater (5) has a combustion portion (53) in which stoichiometric combustion of hydrogen and oxygen is caused, and a mixing portion (55) configured to mix discharged steam (S4) from the high-pressure steam turbine (2) with combustion gas (R) from the combustion portion (53) and to supply the obtained product to the low-pressure steam turbine (3).

Abstract: Systems and methods to increase intake air flow to a gas turbine engine of a hydraulic fracturing unit when positioned in an enclosure may include providing an intake expansion assembly to enhance intake air flow to the gas turbine engine. The intake expansion assembly may include an intake expansion wall defining a plurality of intake ports positioned to supply intake air to the gas turbine engine. The intake expansion assembly also may include one or more actuators connected to a main housing of the enclosure and the intake expansion assembly. The one or more actuators may be positioned to cause the intake expansion wall to move relative to the main housing between a first position preventing air flow through the plurality of intake ports and a second position providing air flow through the plurality of intake ports to an interior of the enclosure.

Abstract: The cooling system may comprise: an electric machine; a first conduit including a cable housing and an inlet; a plurality of conductive cables extending from the electric machine, the plurality of conductive cables disposed at least partially in the cable housing; and an electric fan disposed in the first conduit, the cooling system configured to passively flow air through the first conduit to cool the plurality of conductive cables during operation of the gas turbine engine, and the electric fan configured to actively cool the plurality of conductive cables after an engine shutdown of the gas turbine engine.

Abstract: A variable area nozzle assembly for a gas turbine engine includes a fixed structure surrounding an exhaust duct extending along a nozzle centerline. The fixed structure further includes a first lateral side and a second lateral side opposite the first lateral side. The variable area nozzle assembly further includes a nozzle. The nozzle includes a nozzle outlet including a nozzle outlet cross-sectional area. The variable area nozzle assembly further includes a thrust reverser system including a first thrust reverser door and a second thrust reverser door. The first thrust reverser door is pivotably mounted to the fixed structure at the first lateral side. The second thrust reverser door is pivotably mounted to the fixed structure at the second lateral side. The first thrust reverser door and the second thrust reverser door define a portion of the nozzle outlet of the nozzle.

Abstract: In an embodiment, a gas turbine system is provided. The gas turbine system may include a supplementary air system and a flow diffuser configured to receive compressed air from the supplementary air system. The gas turbine system further comprises a gas turbine compressor downstream the flow diffuser, a combustor downstream the gas turbine compressor, a turbine downstream the combustor, and a turbine haft. The turbine shaft comprises a thrust bearing and a clutch. When the gas turbine system is operating, the gas turbine compressor does not consume electricity.

Abstract: A combined combustor and recuperator is formed with the recuperator surrounding the combustor. Cold gas conduits (14, 16, 20) through the recuperator follow along involute paths toward the combustor. Hot has conduits (26) through the recuperator follow counterflow paths along corresponding involute curves outward from the combustor. The openings (18) in the combustion chamber wall through which cold gas enters the combustion chamber may be directed to impart flow direct to the cold gas to support particular desired behaviour of the cold gas in the portions of the combustion chamber concerned, e.g. supporting a stable vortex flame, enhancing mixing, providing a protective barrier layer.

Abstract: A gas supply system is provided herein. The gas supply system includes a fuel oxygen reduction unit having a circuit defining a gas flowpath for a flow of a stripping gas. A reservoir is in selective fluid communication with the fuel oxygen reduction unit and is configured to store a portion of the stripping gas from the circuit. The reservoir is further configured to be in selective fluid communication with the fuel system component when installed in a vehicle to provide the stored portion of the stripping gas to the fuel system component in response to detection of a purge condition.

Abstract: An Ericsson cycle turbine engine. The Ericsson cycle turbine may comprise: a centrifugal gas compressor, shaft, at least one heat exchanger, and a reaction turbine. The centrifugal gas compressor may function as a spinning wheel trompe and may be fed with a gas-liquid mixture. The centrifugal gas compressor may separate a gas from the gas-liquid mixture and compress that gas via centrifugal acceleration. The shaft may couple to the downstream end of the centrifugal gas compressor and may have an annular space to permit the compressed gas to travel therein. The heat exchanger may introduce heat to the compressed gas, such that isothermal expansion is approached. The reaction turbine may couple to the downstream end of the shaft and may rotate the shaft when releasing the compressed gas against a plurality of vanes. The liquid may be mercury, oil, or water. The gas may be helium, air, argon, or ammonia.

Abstract: A combustor nozzle capable of injecting fuel uniformly, a combustor including the combustor nozzle, and a gas turbine including the combustor are provided. The combustor nozzle includes a premix passage in which air and fuel are mixed, wherein the premix passage includes a plurality of pegs configured to inject fuel and guide an air flow and a plurality of mixing bars configured to mix air and fuel injected from the plurality of pegs, the mixing bar having a twisted structure.

Abstract: A housing for a gas turbine engine accessory includes an enclosure having a drive end with an opening and a closure. The closure is seated in the opening, defines therethrough a shaft aperture and has a v-band flange extending about a periphery of the closure to limit load carried by the enclosure when the engine accessory is mounted to an accessory gearbox mount in a cantilevered arrangement. Engine accessories for gas turbine engines and mounting arrangements for gas turbine engine accessories are also described.

Abstract: A conduit for transport of a fluid, the conduit comprising: a wall extending around and along an axis extending parallel to a direction of bulk fluid flow. The wall having an inner surface defining an interior of a channel through which fluid flows; and a plurality of projections extending from the inner surface of the wall. The plurality of projections extend around the axis, in a plane perpendicular to the axis; and wherein the projections have a height perpendicular to the axis into the channel, and the height is arranged such that the projections modify the flow of fluid at a boundary layer of the fluid adjacent the wall. The conduit is arranged to carry a first fluid; the conduit includes an opening for introducing a second fluid into the channel such that the first and second fluid are mixed downstream of the opening and the plurality of projections are provided upstream of the opening, in a direction of fluid flow.

Abstract: An aircraft component is used for an aircraft gas-turbine engine. The aircraft component includes an annular part, a flange, and a boss. The annular part has an outer circumferential surface. The flange is formed at one end portion of the annular part in an axial direction. The boss projects from the outer circumferential surface of the annular part to the radial direction. On a section cut along an axial direction of the annular part, the outer circumferential surface of the annular part between the flange and the boss has a taper part that is formed into a tapered shape in which plate thickness becomes thicker from the flange toward the boss.

Abstract: The present invention provides a flexible hose for connecting a fuel manifold to a burner of a gas turbine engine. The flexible hose includes a metal convolute tube, an elongate member or members located in grooves formed in the inner surface of the convolute tube, and a pressure-containing sheath outside the convolute tube. The flexible tube has end connectors fluidly-tightly joined to respective ends of the hose for connection at one end of the hose to the fuel manifold of the gas turbine engine, and at the other end of the hose to the burner of the gas turbine engine.

Abstract: A heat exchange system includes: a gas line through which a gas to be cooled flows; a first heat exchanger disposed in the gas line and configured to cool the gas through heat exchange with a refrigerant; a refrigerant introduction line for introducing the refrigerant into the first heat exchanger; a refrigerant discharge line for discharging the refrigerant after cooling the gas from the first heat exchanger; a recirculation line for recirculating at least a part of the refrigerant flowing through the refrigerant discharge line into the refrigerant introduction line; and a flow-rate adjustment unit for adjusting a flow rate of the refrigerant flowing through the recirculation line so that a temperature of the refrigerant introduced into the first heat exchanger from the refrigerant introduction line is not lower than a threshold.

Abstract: An anti-icing system for an aircraft nacelle includes an inner duct and an outer duct. The inner duct extends through a first bulkhead between a gas turbine engine and a plenum of the nacelle. The inner duct is configured to deliver heated gases from the gas turbine engine to the plenum. The outer duct includes a first end and a second end. The outer duct circumferentially encompasses at least a portion of the inner duct on a first side of the bulkhead. At least one of the first end and the second end of the outer duct is mounted to at least one of the respective first bulkhead and a second bulkhead by a bellows forming an annular seal between the respective bulkhead and the outer duct.