The disclosure is towards an inlet cowl for a turbine engine including a surface defining an inlet with a flow path and a method towards controlling the airflow in the flow path. The inlet cowl further includes an inlet lip and inner and outer barrels. The inlet lip confronts the inner barrel at a junction defining a gap.

An air intake system comprising an air duct suitable for providing airflow to the inside of an aircraft, preferably to an auxiliary power unit; an inlet arranged at one end of the air duct; a skin surrounding the inlet; a plurality of slots arranged on the skin; a driving arrangement, a flap door connected to the driving arrangement, and a plurality of fins connected to the driving means. The driving arrangement is configured for moving the flap door between at least two positions, the positions being a closed position wherein the flap door closes the inlet, and an opened position wherein the flap door is driven away from the closed position. The driving arrangement is also configured for moving the plurality of fins such that the plurality of fins protrudes through the slots.

An assembly for a turbomachine through which an air flow passes, includes a stator including guide vanes extending radially in relation to a longitudinal axis, at least one inter-vane platform extending between the radially outer ends of two circumferentially consecutive guide vanes, each inter-vane platform including an outer surface that faces the axis, a heat exchanger located downstream of the stator in relation to a direction of circulation of the air flow in the turbomachine during operation, this stator including a heat exchange surface extending in the extension of an inter-vane platform. At least one inter-vane platform located in the upstream extension of the heat exchange surface is provided with at least one turbulence generator on its outer surface.

A rotating machine includes a rotating body rotatably supported in a casing; a rotor blade fixed to an outer peripheral portion of the rotating body; a stator blade arranged on a downstream side in a fluid flow direction with respect to the rotor blade and fixed to an inner peripheral portion of the casing; a sealing device arranged between the inner peripheral portion and a front end of the rotor blade; a swirling flow generation chamber provided in the casing on the downstream side from the sealing device along a circumferential direction of the rotating body; and guiding members provided in the swirling flow generation chamber along a radial direction of the rotating body and in the circumferential direction at predetermined intervals. The swirling flow generation chamber has a wall surface located on the downstream side from an edge of the stator blade on an upstream side.

Provided is a gas turbine vane and blade assembly in which lattice structures are installed between an impingement plate and an effusion plate. The gas turbine vane and blade assembly is capable of enhancing cooling efficiency in an impingement/effusion cooling technique.

In addition, the gas turbine vane and blade assembly can be manufactured using an additive manufacturing technique, and the lattice structures are capable of replacing supports that are used during an additive manufacturing process, and improving not only structural rigidity and stability but also cooling performance.

A turbine blade structure for improving aerodynamic efficiency of low-pressure turbine blades, including: a suction side, a pressure side, multiple dimples and a blade body. The suction side is an outer convex side of the blade body. The pressure side is an inner concave side of the blade body. The dimples are arranged on the suction side in pairs. Each dimple forms an inclination angle β with an air flow. The air flow includes a first fluid and a second fluid, and the energy of the first fluid is lower than that of the second fluid. Each dimple sucks the first fluid at a first end when the air flow passes a surface of the blade body, and allows the first fluid to spirally flow along an inclined direction in each dimple to form a spiral vortex, and discharge the first fluid through a second end.

An aircraft engine has: a first component and a second component coaxially mounted about a central axis; a flow passage extending within an annular gap defined radially between the first component and the second component, the flow passage fluidly connecting a first zone to a second zone; a seal disposed in the flow passage between the first zone and the second zone, the seal extending from a seal base secured to the first component to a seal end radially spaced apart from the seal base; a deflector located downstream of the seal relative to a first flow flowing from the first zone to the second zone through the seal, the deflector extending from a deflector base secured to the second component to a deflector end radially spaced apart from the deflector base; and a radial gap defined radially between the seal end and the deflector end.

A gas turbine engine including: an exhaust diffuser including an inner tube and an outer tube that form therebetween an annular exhaust passage; a bearing chamber formed radially inside the inner tube for accommodating a bearing that supports a rotor of a turbine; a plurality of hollow struts extending across the exhaust passage; an oil introduction passage extending through one of the struts for introducing oil to be supplied to the bearing chamber; an oil drain passage extending through one of the struts for draining the oil from an exhaust oil inlet opened on a bottom surface of the bearing chamber; and an oil discharge passage for discharging a portion of the oil having passed through the oil introduction passage toward the oil drain inlet.

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 extending circumferentially around the axis X and comprising an inner wall, which faces the axis X in order to guide an inner air flow, and an outer wall, which is opposite the inner wall, for guiding an external air flow, the walls being connected by a leading edge and an inner partition so as to delimit an annular cavity. The air intake comprises 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 internal cavity, the air intake comprising at least one disruption member of the external air flow, positioned upstream of the ventilation orifice, which extends outwardly from the outer wall.

A hybrid rocket engine with a vortex flow field injection system that produces a high-speed sustained vortex flow field is described. The hybrid rocket engine includes a generally cylindrical injection chamber with an inner circumference to comprise an outer edge of a solid propellant grain in the hybrid rocket engine. The engine also includes an injection system that has a throttle valve and an injector that injects injection fluid into the engine and produces a vortex flow-field for the injected fluid. The injector includes at least one primary feed line that distributes the injection fluid throughout a pre-swirl chamber and multiple orifices along an inner edge of the injection chamber. The pre-swirl chamber connects to the injection chamber and at least one of the primary feed lines and redirects a primary fluid flow of the injected fluid from a primary axial direction to a centrifugal direction.