A turbofan engine is provided. The turbofan engine includes a fan comprising a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; a nacelle surrounding and at least partially enclosing the fan; an inlet pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and a means for reducing ice buildup or ice formation on the inlet pre-swirl feature, the means in communication with the inlet pre-swirl feature.

To provide a method of designing a blade of an axial flow fluid machine that has a blade surface whose radius of curvature is continuous at a leading edge thereof and has a high aerodynamic performance. The method includes a step of determining a pressure surface curve and a suction surface curve as curves capable of first to third order differentiations at respective connection points to a leading edge curve, that is, a pressure surface connection point and a suction surface connection point, and a step of forming the leading edge curve as a fifth order Bezier curve that is defined by a first control point, a second control point, a third control point, a fourth control point, a fifth control point and a sixth control point. The first control point is the suction surface connection point. The sixth control point is the pressure surface connection point. Provided that an intersection of a tangent to the pressure surface curve at the pressure surface connection point and a tangent to the suction surface curve at the suction surface connection point is referred to as a point ahead of the blade, the second and fifth control points are points that internally divide line segments connecting the point ahead of the blade to the suction surface connection point and the pressure surface connection point, respectively. The third and fourth control points are points having coordinates that are determined as solutions of simultaneous equations obtained by applying continuity conditions for first to third differential coefficients at the first and sixth control points to a fifth order Bezier function.

A turbomachine vane includes a blading made of composite material with fibrous reinforcement densified by a matrix and an integrated metal leading edge, the blading extending in a longitudinal direction, the leading edge being formed by a metal foil overmolded onto the blading, the foil having a lower wing and an upper wing which extend respectively on the lower and upper faces of the blading while matching an aerodynamic profile of the vane. One of the lower wing and the upper wing has a positioning portion extending in the longitudinal direction, the portion having a flat inner face and an increasing thickness away from the leading edge, and being housed in a correspondingly shaped groove in the blading.

A nozzle vane for a variable geometry turbocharger has an airfoil including a leading edge, a trailing edge, a pressure surface, and a suction surface at least in a center position in a blade height direction. The airfoil satisfies 0≤W.sub.max/L<0.03, where W.sub.max is a maximum value of a distance from a line segment connecting the trailing edge and a fixed point on the pressure surface at a 40% chord position from the trailing edge toward the leading edge to a given point on the pressure surface between the trailing edge and the fixed point, and L is a length of the line segment.

A blade includes an airfoil defined by a pressure side outer wall and a suction side outer wall connecting along leading and trailing edges and form a radially extending chamber for receiving a coolant flow. A rib configuration may include: a leading edge transverse rib connecting the pressure side outer wall and the suction side outer wall and partitioning the radially extending chamber into a leading edge passage within the leading edge of the airfoil and a central passage adjacent to the leading edge passage. One or both camber line ribs connect to a corresponding pressure side outer wall and suction side outer wall at a point aft of the leading edge transverse rib causing the central passage to extend towards one or both of the pressure side outer wall and the suction side outer wall, resulting in a flared center cavity aft of the leading edge.

A turbomachine includes a rotor and a stator, the stator having a plurality of profiled structures, each profiled structure being elongated in a direction of elongation in which the profiled structure has a length exposed to an airflow, and having a leading edge and/or a trailing edge, at least one of which is profiled and has, in said direction of elongation, serrations defined by a succession of peaks and troughs and having a geometric pattern transformed, over at least a part of said length exposed to the airflow, by successive scaling, via multiplicative factors, in the direction of elongation and/or transverse to the direction of elongation. The geometric pattern, as defined with reference to a radial distribution of the integral scale of the turbulence, evolves in a non-periodic manner.

A cooling assembly includes a coolant chamber disposed inside an airfoil of a turbine assembly that directs coolant inside the airfoil. The airfoil extends between a leading edge and a trailing edge along an axial length of the airfoil. Inlet cooling channels are fluidly coupled with the coolant chamber and direct the coolant in a direction toward a trailing edge chamber of the airfoil. The trailing edge chamber is fluidly coupled with at least one inlet cooling channel. The trailing edge chamber is disposed at the trailing edge of the airfoil and includes an inner surface. The inlet cooling channels direct at least a portion of the coolant in a direction toward the inner surface of the trailing edge chamber. One or more outlet cooling channels direct at least a portion of the coolant in one or more directions away from the trailing edge chamber.

Disclosed examples include a retrofit fan frame assembly, comprising: a leading edge adjustment component coupleable to an airfoil, the leading edge adjustment component of variable chord length, the variable chord length to increase and then decrease along a radial length from a hub end of the airfoil to an opposite tip end of the airfoil; and an attachment mechanism configured to couple the leading edge adjustment component to a leading edge of the airfoil.

The invention relates to a diffuser for a radial compressor, comprising a flow channel defined by a first side wall and a second side wall, a diffuser vane ring with a plurality of diffuser vanes that are at least partially arranged in the flow channel, each of the diffuser vanes having a pressure side and a suction side, a plurality of diffuser passages, said diffuser passages being formed between every two adjacent diffuser vanes of the plurality of diffuser vanes, and circulation openings, each circulation opening connecting the flow channel to a diffuser cavity, at least two circulation openings being associated with one diffuser passage, and a circulation opening associated with a diffuser passage being fluidically connected to another circulation opening associated with the same diffuser passage or to a circulation opening associated with another diffuser passage, via the diffuser cavity.

A blade for a gas turbine engine comprises an aerofoil body having a suction side, a pressure side, and a trailing edge. An internal cooling passageway is provided in the aerofoil body, and an ejection slot in fluid communication with the cooling passage and provided at the trailing edge of the aerofoil body. The ejection slot is defined between a pressure side wall and a suction side wall. Both the suction side wall and the pressure side wall include a mid-section and a trailing edge section adjacent the mid-section, and the thickness of the suction side wall and the pressure side wall reduces to define a taper with a wedge angle less than or equal to 20 degrees.