The invention relates to a ventilation device for a turbomachine turbine casing, comprising a plurality of line sets (16) configured to spray air over the turbine casing, the line sets being arranged next to one another, each line set comprising a main ring (161) in which air circulates, the main ring (161) comprising orifices (17) configured to spray a stream of air towards the turbine casing, the line set comprising a shield (162) configured to isolate the main ring (161) from a stream of air returning from the turbine casing towards the line sets after having been sprayed towards the turbine casing, the said shield (162) enveloping the main ring (161) and having orifices aligned with the orifices of the main ring (161).

The proposed invention is an air-flow straightening assembly for a turbine engine comprising:a cylindrical platform (15) centred on an axis (X-X),at least one straightener blade (20) extending from the platform,a structural arm (30) extending radially relative to the axis, anda mechanical member (40) protruding from the platform (15), said mechanical member (40) being one of the group comprising:a radial shaft,an angle transmission box of a radial shaft,an electric, hydraulic or pneumatic connection element,an intermediate gear driving a radial shaft, the straightening assembly further comprising a fairing (50) of the protruding mechanical member, the fairing having a three-dimensional surface defined by:at least one upstream end pont (Ai, A.sub.e) located axially upstream from the mechanical member (40) relative to the direction of air flow in the turbine engine, andat least one downstream end point (Ci, C.sub.e) located axially downstream from the mechanical member, the three-dimensional surface being tangential to the platform at the upstream and downstream end points (Ai, A.sub.e, Ci, C.sub.e), and having a larger cross-section measured along an axis (Y-Y) orthogonal to the first, and in which the three-dimensional surface is further defined by two lateral end points (B.sub.i, B.sub.e) corresponding to the ends of said larger cross-section respectively on the pressure side and suction side of the structural arm (30), the axial positions of said points being separated by at most 0.1 C.sub.OGVin which C.sub.OGV is the chord of the straightener blade (20).

Apparatus for mounting a component in a gas turbine engine, the apparatus comprising: a casing for a gas turbine engine, the casing including a flange; a first connector to connect a component of the gas turbine engine to the casing at a first location adjacent to the flange; and a second connector to connect the component of the gas turbine engine to the casing at a second location adjacent to the flange.

An airfoil for a gas turbine engine, the airfoil includes a wall that has a leading edge and a trailing edge and at least partially defining a boundary of a leading edge cavity radially along the leading edge. A cooling jet structure is operatively associated with a portion of the wall proximate the leading edge and is configured to direct a cooling fluid tangent to the portion of the wall.

A support for carrying a tube for discharging oil-laden air from a turbine engine, the support including a radially inner annular portion for mounting around the tube, and fins extending outwards in a radial plane from the annular portion, each fin forming an angle relative to the radial direction. The fins have fastener zones at their outer peripheries, the fastener zones being inclined in the axial direction of the support so as to be suitable for fastening to an exhaust cone of the turbine engine.

A method of bleeding air from a gas turbine engine having a compressor section with a downstream centrifugal compressor is described, which includes bleeding air from a first location upstream of the centrifugal compressor's outlet, when the compressor is operating in a first pressure range, and bleeding air from a second location upstream of the centrifugal compressor's outlet, when the engine is operating in a second pressure range.

The present invention relates to a cooling-air injection casing (2) for cooling a bladed rotor disc of a turbine, in particular a high-pressure turbine, of a turbomachine, this casing extending around a longitudinal axis (X-X) and being traversed by at least one channel (20) forming an air injector, the channel (20) comprising an inlet mouth (201) and an outlet mouth (202). This casing is noteworthy in that the channel (20) comprises a primary section (203) which extends in an axial plane (P) from the inlet mouth (201) to an elbow (206), and a secondary section (207) which extends from this elbow (206) to the outlet mouth (202), the secondary section (207) having a progressive variation of its orientation with a tangential component between the elbow (206) and the outlet mouth (202), in that the channel (20) has a reduction in cross section between the inlet mouth (201) and a neck (204), and in that the channel (20) has at least one corrugation (205) in its primary section (203), such that the outlet mouth (202) is situated closer to the longitudinal axis (X-X) than the inlet mouth (201).

A gas turbine engine includes a turbine section that includes a turbine rotor arranged in a plenum. A compressor section includes a compressor rotor assembly that has spaced apart inner and outer portions that provide an axially extending cooling channel. Compressor blades extend radially outward from the outer portion which provides an inner core flow path. A rotor spoke is configured to receive a first cooling flow and fluidly connect the outer portion to the cooling channel. The compressor rotor assembly has a coolant exit that is in fluid communication with the cooling channel. The compressor rotor assembly is configured to communicate the first cooling flow to the turbine rotor. A bleed source is configured to provide a second cooling flow. A combustor section includes an injector in fluid communication with the bleed source. The tangential onboard injector is configured to communicate the second cooling flow to the turbine rotor.

A method of transferring a fluid flow from a static component to a rotor of a gas turbine engine and a modulated flow transfer system are provided. The modulated flow transfer system includes an annular inducer configured to accelerate the fluid flow in a substantially circumferential direction in a direction of rotation of the rotor. The system further includes a first fluid flow supply including a compressor bleed connection, a feed manifold formed of bendable tubing, and a feed header extending between the compressor bleed connection and the feed manifold. The feed header includes a modulating valve configured to control an amount of fluid flow into the feed manifold. The system also includes a flow supply tube that extends between the feed manifold and the inducer and is couplable to at least one of the plurality of first fluid flow inlet openings through a sliding piston seal.

An injector for introducing hydrogen and gas into a combustion chamber of a gas turbine engine includes a convergent-divergent nozzle head and a hydrogen/gas feed nozzle. The convergent-divergent nozzle head is arranged along a nozzle axis and has an upstream end defining an inlet mouth. The hydrogen/gas feed nozzle includes a mixing chamber arranged along the nozzle axis upstream of the inlet mouth, first feed conduits configured to feed hydrogen into the mixing chamber, and second feed conduits configured to feed gas into the mixing chamber. Each of the first feed conduits and the second feed conduits are tangentially oriented to the mixing chamber.