A turbine blade having a gas film cooling structure with a composite irregular groove. The turbine blade has a hollow structure, and a plurality of first grooves which are recessed grooves are provided on an outer surface thereof. A plurality of discrete holes A extending to an inner surface of the turbine blade are provided at the groove bottom of each first groove. The first groove is an irregular groove, and includes at least two portions in a depth direction. A portion having a depth H.sub.1 from the groove bottom of the first groove is a first portion, and the rest thereof is a second portion. At least one side surface of the second portion is formed by expanding in lateral direction from a corresponding side surface of the first portion.

An engine component assembly with a first substrate having a hot surface in thermal communication with a hot combustion gas flow and a cooling surface, with the cooling surface being different than the hot surface, a second substrate having a first surface in fluid communication with a cooling fluid supply and a second surface, different from the first surface, facing and spaced from the cooling surface to define at least one interior cavity. At least one cavity is provided within the first substrate defining a cavity surface where at least a portion of the cavity surface is directly opposite the second surface. At least one cooling aperture extending through the second substrate from the first surface to the second surface, and defining a streamline along which a cooling fluid passes from the cooling fluid supply to the at least one interior cavity.

An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S.sub.tip, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition M.sub.rel, wherein M.sub.rel is between 0.4 and 0.93, L/D is between 0.2 and 0.45 and S.sub.tip is from −1 to 0.1.

A casing assembly for a gas turbine engine includes a plurality of vane casing segments. Each vane casing segment includes an arcuate member and a pair of split-line flanges. Each split-line flange includes a first axial flange end and a second axial flange end. Each split-line flange is fixedly coupled to an adjacent split-line flange of an adjacent vane casing segment. Each vane casing segment includes at least one row of stator vanes welded to the arcuate member. Each split-line flange is at least partially and circumferentially inclined relative to a rotational axis of the gas turbine engine, such that the first axial flange end is circumferentially offset from the second axial-flange end. Further, each split-line flange includes an intersecting portion, such that at least the intersecting portion of each split-line flange is circumferentially inclined relative to the rotational axis.

A stator vane for a turbomachine, including an airfoil having a leading edge and a trailing edge, which are interconnected by a suction side and a pressure side, the airfoil extending substantially in a radial direction between an inner platform and an outer platform, wherein, when viewed in an axial direction of the turbomachine toward the trailing edge, the trailing edge extends so as to be inclined at a first angle to the pressure side radially on an outside in relation to the radial direction at a circumferential position at which the trailing edge meets the outer platform, and wherein the first angle between the trailing edge and a tangent to the outer platform at a transition from the outer platform to the trailing edge is between 72° and 84°.

A stator vane 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, Table IX, Table X, Table XI, or Table XII. 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.

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 comprising an inner wall facing the axis X and an outer 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 to allow the hot air flow to escape after heating the inner cavity, the ventilation orifice comprising an upstream edge, the circumferential profile of which is discontinuous in order to generate turbulences, and a downstream edge, the radial profile of which is aerodynamic in order to limit the formation of pressure fluctuations.

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.

An impeller includes Z blades, where Z is an integer equal to 5 or more, arranged in a circumferential direction of the impeller and extending radially, and pitch angles between adjacent blades are all different. In terms of an arbitrary pitch angle θ, when a pitch angle α1 adjacent to the pitch angle θ, and a pitch angle α2 adjacent to the pitch angle θ, different from the pitch angle α1, satisfy a relation, α1<α2, a pitch angle β1 different from the pitch angle θ adjacent to the pitch angle α1 and a pitch angle β2 adjacent to the pitch angle α2, different from the pitch angle θ, satisfy a relation, β2<β1.

Various aspects of the disclosure are directed to providing structures that define a radial gap between an impeller and a pump casing element that facilitates minimizing the movement of fluid into the radial gap in a manner that lessens the impact, and consequent degradation, of the inner surface of the pump casing element by movement of abrasive particulates out of the radial gap, which is accomplished by providing a suction inlet arrangement of an impeller and pump casing element that are angled from the eye of the impeller to the outer periphery of the impeller in a direction away from the back shroud or drive side of the impeller and toward a first end of the pump casing in which fluid is introduced into the pump casing.