A fan blade anti-icing system comprises a fan hub and a fan blade extending radially outwardly from the fan hub. The fan blade has a base and an airfoil extending radially outwardly from the fan base. The airfoil having a leading edge, a trailing edge, a convex side surface between the leading and trailing edge and a concave side surface between the leading and trailing edge. The fan blade further has a radial passage extending from a blade air inlet in the blade base in communication with a source of heated air, and a rearwardly directed passage in communication with the radial passage and having a blade air outlet forward of the trailing edge and oriented tangentially to the convex side surface or concave side surface of the airfoil.

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 airflow straightening assembly for a turbine engine comprising: a cylindrical platform (15) centered on an axis (X-X), at least one straightener blade (20) extending from the platform, a structural unit (30) extending radially relative to the axis, and a 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) or the protruding mechanical member, the fairing having a three-dimensional surface defined by: at least one upstream end point (Ai, A.sub.e) located axially upstream from the mechanical member (40) relative to the direction of air flow in the turbine engine, and at 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 point (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 or said points being separated by at most 0.1 C.sub.QGVin which C.sub.OGV is the chord of the straightener blade (20).

Various embodiments of a vortex hybrid motor are described herein. In some embodiments, the vortex hybrid motor may include a combustion zone defined by a fuel core and/or motor housing. The combustion zone may include an upper zone and a central zone that each contribute to thrust created by the vortex hybrid motor. In some embodiments, an injection port configuration is described that includes a proximal injection port that may be controlled for modulating a delivery of an amount of oxidizer for adjusting an oxidizer-to-fuel ratio. In some embodiments, a fuel core configuration is described that provides radially varying gradients of fuel in order to achieve desired thrust profiles. In some embodiments, the fuel core may include a support structure and/or a proximal end of a nozzle of the vortex hybrid motor may extend into the fuel core.

A cooling circuit for a gas turbine engine comprises a gas turbine engine component having a first portion connected to a second portion via a curved surface. An inlet is formed in or near one of the first and second portions to receive a cooling air flow. An outlet is formed in or near the other of the first and second portions to direct cooling flow along a surface of the gas turbine engine component. At least one cooling path extends between the inlet and the outlet and has at least one cooling path portion that conforms in shape to the curved surface. A gas turbine engine and a method of forming a cooling circuit for a gas turbine engine are also disclosed.

A gas turbine engine comprises a high speed spool and a low speed spool. A low speed power takeoff is driven by the low speed spool and a high speed power takeoff is driven by the high speed spool. The low speed power takeoff drives a plurality of low speed driven accessories about low speed spool drive axes and the high speed power takeoff drives a plurality of high speed driven accessories about high speed spool drive axes. The low speed spool drive axes are generally tangential to an envelope defined about a center axis of the engine, and the high speed spool drive axes are generally tangential to an envelope defined about the center.

An air turbine starter having a housing, a turbine, a drive shaft, and at least one vane. The housing having an inlet, an outlet and a curvilinear flow path extending between the inlet and the outlet. The at last one vane is located within a portion of the curvilinear flow path, and includes an outer wall extending between a root and a tip in a span-wise direction and between a leading edge and a trailing edge in a chord-wise direction. In one aspect, the vane is arranged to define an acute axial angle that is non-constant in the chord-wise or span-wise direction. In another aspect, the vane is arranged to define an acute tangential angle that is non-constant in the chord-wise or span-wise direction.

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.

A gas turbine engine comprises a high speed spool and a low speed spool. A low speed power takeoff is driven by the low speed spool and a high speed power takeoff is driven by the high speed spool. The low speed power takeoff drives a plurality of low speed driven accessories about low speed spool drive axes and the high speed power takeoff drives a plurality of high speed driven accessories about high speed spool drive axes. The low speed spool drive axes are generally tangential to an envelope defined about a center axis of the engine, and the high speed spool drive axes are generally tangential to an envelope defined about the center.

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.