Fuel injectors having a fuel-air heat exchange system and methods thereof for use in a gas turbine engine. The fuel-air heat exchanger allows heat transfer between a flow of cooling air used to cool components of the engine and a flow of fuel used to drive the engine to transfer heat to the flow of fuel and cool the cooling air.

The present disclosure relates to a rocket engine, and more particularly, a rocket engine with an integrated combustor head and turbopump in which a turbopump of the rocket engine is formed integrally with a combustor head.

A bearing housing for supporting a bearing of a gas turbine engine. The bearing housing comprises an annular body extending around an axis and having an oil inlet and an oil outlet. The bearing housing further comprises an oil manifold integrated into the annular body, the oil manifold having an inner oil channel defined in the annular body and extending circumferentially around the axis, the inner oil channel in fluid communication with the oil inlet and at least one oil jet directed towards the bearing.

A core engine article includes a combustor liner defining a combustion chamber therein and a turbine nozzle. The combustor liner includes a plurality of injector ports, and the plurality of injector ports have a shape that tapers to a corner on a forward side of the injector ports. The turbine nozzle includes a plurality of airfoils. The combustor liner and turbine nozzle are integral with one another. A method of making a core engine article is also disclosed.

An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a split case structure. The split case structure includes a first wall, a second wall, a first case segment and a second case segment. The first wall extends axially along and circumferentially about an axial centerline. The second wall extends axially along and circumferentially about the axial centerline. The second wall is radially outboard of and axially overlaps the first wall. The first case segment is configured to form a first portion of the first wall and a first portion of the second wall. The second case segment is configured to form a second portion of the first wall and a second portion of the second wall. The second case segment is circumferentially adjacent and attached to the first case segment at a joint.

The invention relates to an assembly (601) which comprises a front frame (602) of a thrust reverser structure of an aircraft nacelle and at least one cascade (603). The assembly is characterized in that, when the cascade is fixed to the frame, the axial distance (L) that separates the first blading (605) upstream of the cascade from a covering area (606) between the extension of the cascade and the front frame is greater than or equal to once the mean height (h) of the cascade.

Disclosed is a turbine arrangement for a supercharging device. The turbine arrangement comprises a turbine housing, a turbine wheel and a guide device. The turbine housing defines a turbine spiral and a turbine outlet. The turbine wheel is arranged in the turbine housing between the turbine spiral and the turbine outlet. The guide device comprises a carrier ring and multiple guide blades. The guide blades are arranged on the carrier ring fixedly in a predetermined orientation. The guide device is arranged in an inflow channel between the turbine spiral and the turbine wheel such that, during operation, fluids are conducted from the turbine spiral through the inflow channel over the guide blades onto the turbine wheel.

A heat exchanger for a gas turbine engine includes a heat exchanger body having a first surface and a second surface oriented at least partially at an oblique angle relative to the first surface. The heat exchanger body defines a plenum extending between the first and second surfaces. Furthermore, the heat exchanger body defines a fluid passage extending through the second surface such that the fluid passage is in fluid communication with the plenum. The fluid passage, in turn, includes first and second portions. The first portion intersects the plenum at an intersection and defines a line of projection extending normal to the second surface. The second portion defines a line of projection extending normal to the first surface. The fluid passage further includes a curved portion extending from the first portion to the second portion.

A turbomachine turbine casing that extend around an axis and includes an annular wall and a cooling device. The annular wall is provided with a casing hook which extends in radial protrusion from an inside of the annular wall. The casing hook allows a mounting, on the turbomachine turbine casing, of ring segments disposed circumferentially end to end around the axis. The cooling device includes a collector duct intended to convey cooling air, the collector duct extending circumferentially around the annular wall. The collector duct has a cooling air inlet and a cooling air outlet. The collector duct and the annular wall have a common portion, which delimits the collector duct and from which the corresponding casing hook extends.

Described is a rotor disk (40) for a compressor (29, 32) of a gas turbine, in particular an aircraft gas turbine (10), the rotor disk having a main body (42), at least one rotor arm (44) projecting from the main body (42) in the axial direction (AR), the rotor arm (44) having, in a sectional view taken in a sectional plane defined by the axial direction (AR) and the radial direction (RR) a beginning portion (44a) merging into the main body (42); an end (44e) portion remote from the main body (42) and forming a kind of free end in the axial direction (AR), the beginning portion (44a) and the end portion (44e) being interconnected by an intermediate portion (44z), characterized in that the intermediate portion (44z) is curved with at least one radius of curvature (Ri, Ra).