A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.

The invention relates to a bleed air extraction device for a turbomachine, which has: an axial compressor, formed in a flow path and having at least one compressor stage, which comprises a rotor and a stator, and a bleed air duct, which is provided and designed to guide a bleed air flow branched off from the flow path of the axial compressor. In this case, the bleed air duct comprises an inlet opening, which is formed downstream of a stator of the axial compressor in the radially outer flow path boundary, an axially forward wall adjoining the inlet opening, and an axially rearward wall adjoining the inlet opening. Guide means are provided, which are provided and designed for the purpose of guiding at least a portion of the bleed air flow branched off from the flow path in the direction of the axially forward wall of the bleed air duct.

A rotor blade includes an airfoil having an airfoil shape. The airfoil shape has a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in one of Table I, Table II, Table III, Table IV, Table V, Table VI, Table VII, Table VIII, or Table IX. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances expressed in a unit of distance by multiplying the Cartesian coordinate values of X, Y and Z by a scaling factor of the airfoil in the unit of distance. The X and Y values, when connected by smooth continuing arcs, define airfoil profile sections at each Z value. The airfoil profile sections at Z values are joined smoothly with one another to form a complete airfoil shape.

A disclosed fan drive gear system includes a sun gear rotatable about an axis of rotation, a plurality of intermediate gears rotatable about an intermediate gear rotation axis in meshing engagement with the sun gear and a ring gear circumscribing the intermediate gears. A bearing assembly supports at least one of the plurality of intermediate gears and includes a first beam extending in a first direction and a second beam extending from an end of the first beam in a second direction. The bearing surface supported on the second beam such that first and second beams are configured to maintain the bearing surface substantially parallel to the intermediate gear rotation axis during operation.

A sub-assembly for a low-pressure compressor of an aircraft turbine engine includes a straightener provided with cantilevered vanes and a rotor hub having a cavity covered by an inner shroud opposite the vanes. Orifices are made in the inner shroud to allow an air flow to circulate in the downstream to upstream direction of the low-pressure compressor.

A vane assembly for a gas turbine engine which is a single unitary component that includes an aerofoil. A leading edge passageway is disposed proximal to a leading edge of the aerofoil and configured to receive a flow of a fluid therein. The vane assembly further includes a connecting passageway fluidly communicating the leading edge passageway with a trailing edge distribution passageway that is spaced apart from the leading edge, the leading edge passageway and a trailing edge of the aerofoil. The vane assembly further includes a plurality of trailing edge passageways disposed proximal to a pressure surface of the aerofoil and extending from the trailing edge distribution passageway towards the trailing edge along a chordwise direction. Each trailing edge passageway is configured to discharge the fluid through a corresponding passageway outlet disposed on the pressure surface and in fluid communication with a corresponding trailing edge passageway.

The invention relates to a bleed air extraction device for a turbomachine, which has: an axial compressor, formed in a flow path and having at least one compressor stage, which comprises a rotor and a stator, and a bleed air duct, which is provided and designed to guide a bleed air flow branched off from the flow path of the axial compressor. In this case, the bleed air duct comprises an inlet opening, which is formed downstream of a stator of the axial compressor in the radially outer flow path boundary, an axially forward wall adjoining the inlet opening, and an axially rearward wall adjoining the inlet opening. Guide means are provided, which are provided and designed for the purpose of guiding at least a portion of the bleed air flow branched off from the flow path in the direction of the axially forward wall of the bleed air duct.

The gas turbine engine includes a core engine, a low pressure turbine, a fan assembly, a gearbox, and a lubrication scavenge pump. The core engine includes a high pressure compressor, a combustor, and a high pressure turbine configured in a serial flow arrangement. The low pressure turbine is positioned axially aft of the core engine. The fan assembly is positioned axially forward of the core engine. The gearbox is positioned axially forward of the fan assembly. The lubrication scavenge pump is positioned forward of the gearbox.

A compressor section for a gas turbine engine according to an example of the present disclosure includes, among other things, a low pressure compressor including a plurality of rotor blades arranged about an axis, a high pressure compressor, and a core flowpath passing through the low pressure compressor. The core flowpath at the low pressure compressor defines an inner diameter and an outer diameter relative to the axis. The outer diameter has a slope angle relative to the axis.

A vane retainer for a gas turbine engine includes a retainer body, an inner surface of the retainer body to engage complimentary threads of a mating component, and a plurality of castellations located at an end of the retainer body for engagement with complementary installation tool features during installation of the vane retainer to the mating component. A gas turbine engine includes a combustor and a plurality of vane assemblies in fluid communication with the combustor. Each vane assembly includes a vane having an airfoil portion and a vane stem extending from the airfoil portion. The vane stem is inserted into a bushing. A vane retainer is engaged with the bushing and includes a retainer body, an inner surface of the retainer body to engage complimentary threads of the bushing, and a plurality of castellations located at an end of the retainer body for engagement with complementary installation tool features.