A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A turbine housing defining a pair of volutes with respective outlets divided by a divider wall, includes a diffuser space in the gas flow path between the volutes and the turbine wheel. The diffuser space has an upstream portion having a smaller axial extent than a downstream portion of the diffuser space. The widening of the diffuser space tends to direct exhaust gas entering the diffusion space from at least one side of the divider wall towards the corresponding axial end of the diffuser space. Thus reduces the tendency of this gas to interrupt the flow into the diffuser space of exhaust gas from the other inlet volute.

A turbine wheel, in particular in a charging device for use in an internal combustion engine, is specified, wherein the turbine wheel (10) comprises a plurality of blades (12) on a hub (16) that forms a rear wall (14), wherein adjacent blades (12) form an inlet surface (18) having two leading edges (20) and an outlet surface (22) having two trailing edges (24) and situated substantially axially inward, wherein a surface (26) of a blade (12) is configurable by way of an angle (T) and a length (Z0) of a plurality of curvatures (30, . . . , 38) situated next to one another between the leading edge (20) and the trailing edge (24), wherein, for each of the curvatures (30, . . . , 38), the angle (T) of the leading edge (20) initially increases or remains constant and then decreases as the length (Z0) increases so as to form a maximum (40, 40, 40)

A gas turbine engine assembly includes a fan. A diameter of the fan has a dimension D. The fan has a pressure ratio of greater than 1.20 and less than 1.45. A leading edge on an inlet portion of a nacelle is within a first reference plane oriented at an oblique angle. A forward most portion on the fan blade leading edges is in a second reference plane. A length of the inlet portion has a dimension L different at a plurality of locations on the inlet portion. A geared architecture has a gear reduction ratio of greater than 2.3, a bypass ratio is greater than 10, and a low pressure turbine includes a pressure ratio greater than 5:1. A dimensional relationship of UD is between 0.25 and 0.45. The leading edge on the inlet portion is further from the second reference plane near the top of the assembly.

According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is between 0.20 and 0.40.

A gas turbine engine component according to an exemplary aspect of the present disclosure includes, among other things, a body having a first outer face meeting a second outer face at an intersection, the body having a plurality of apertures extending from an opening in the first outer face to an opening on the second outer face; and a coating filling at least a portion of the plurality of apertures.

A thermal barrier coating (TBC) with depth-varying material properties is formed on a turbine component. Exemplary depth-varying material properties include physical ductility, strength and thermal resistivity that vary from the TBC layer inner to outer surface. Exemplary ways to modify physical properties include application of plural separate overlying layers of different material composition or by varying the applied material composition during the application of the TBC layer. Various embodiment described herein also apply a calcium-magnesium-aluminum-silicon (CMAS)-retardant material over the TBC layer to retard reaction with or adhesion of CMAS containing combustion particulates to the TBC layer. In other embodiments the CMAS retardant material is also applied within engineered groove features (EGFs) that are formed in the TBC surface.

According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is no more than 0.45.

A turbine engine casing flow-path segment that is locally diffusing, followed by a flow-path segment contracting in the vicinity of a fan blade. This contraction accelerates the fluid flow axially forward of the fan blade leading edge at the tip and converges with the linear flow-path aft of the fan blade leading edge but forward of the fan blade trailing edge. More diffused fluid flow results in increased flow capacity of the fan, and increased fan efficiency.

An airfoil for a gas turbine engine, the airfoil comprising a wall having a first surface, a second surface, and a passageway extending through the wall from a first opening in the first surface to a second opening in the second surface, the passageway having one or more sections between the first opening and the second opening, the one or more sections in fluid communication with each other, the plurality of sections comprising a first diffuser section providing a first change in cross-sectional area within the passageway, a second diffuser section providing a second change in cross-sectional area within the passageway, a flow conditioning section, and an edge section having two surfaces set opposite each other across the passageway, the two surfaces extending along the passageway substantially in parallel to one another, the edge section being located adjacent to the second opening.