Abstract: An airfoil for a gas turbine engine includes an airfoil with pressure and suction sides that are joined at leading and trailing edges. The airfoil extends a span from a support to an end in a radial direction. 0% span and 100% span positions respectively correspond to the airfoil at the support and at the end. The leading and trailing edges are spaced apart from one another an axial chord in an axial direction. A cross-section of the airfoil at a span location has a diameter tangent to the pressure and suction sides. The diameter corresponds to the largest circle fitting within the cross-section. A ratio of the diameter to the axial chord is at least 0.4 between 50% and 95% span location.

Abstract: A strut for a gas turbine engine includes an airfoil section extending in a spanwise direction between a first platform and a second platform, extending in a chordwise direction between a leading edge and trailing edge to define a chord length, and extending in a thickness direction between a first side and a second side to define a chord width. Exterior surfaces of the airfoil section define a leading portion between the leading edge and a widest location of the airfoil section relative to the thickness direction, and a trailing portion between the widest location and the trailing edge. The exterior surfaces establish a respective exterior contour for each span position between a 0% span position and a 100% span position. The exterior surfaces define a plurality of dimples in the leading portion.

Abstract: A component for a gas turbine engine includes a platform that has a radially inner side and a radially outer side. A root portion extends from the radially inner side of the platform. An airfoil extends from the radially outer side of the platform. The airfoil includes a pressure side that extends between a leading edge and a trailing edge. A suction side extends between the leading edge and the trailing edge. A curvature inflection point is located between 30% and 70% of an axial chord length of the airfoil.

Abstract: An airfoil for a gas turbine engine includes an airfoil with pressure and suction sides that are joined at leading and trailing edges. The airfoil extends a span from a support to an end in a radial direction. 0% span and 100% span positions respectively correspond to the airfoil at the support and at the end. The leading and trailing edges are spaced apart from one another an axial chord in an axial direction. A cross-section of the airfoil at a span location has a diameter tangent to the pressure and suction sides. The diameter corresponds to the largest circle fitting within the cross-section. A ratio of the diameter to the axial chord is at least 0.4 between 50% and 95% span location.

Abstract: An airfoil for a gas turbine engine includes an airfoil with pressure and suction sides that are joined at leading and trailing edges. The airfoil extends a span from a support to an end in a radial direction. 0% span and 100% span positions respectively correspond to the airfoil at the support and at the end. The leading and trailing edges are spaced apart from one another an axial chord in an axial direction. A cross-section of the airfoil at a span location has a diameter tangent to the pressure and suction sides. The diameter corresponds to the largest circle fitting within the cross-section. A ratio of the diameter to the axial chord is at least 0.4 between 50% and 95% span location.

Abstract: A strut for a gas turbine engine includes an airfoil section extending in a spanwise direction between a first platform and a second platform, extending in a chordwise direction between a leading edge and trailing edge to define a chord length, and extending in a thickness direction between a first side and a second side to define a chord width. Exterior surfaces of the airfoil section define a leading portion between the leading edge and a widest location of the airfoil section relative to the thickness direction, and a trailing portion between the widest location and the trailing edge. The exterior surfaces establish a respective exterior contour for each span position between a 0% span position and a 100% span position. The exterior surfaces define a plurality of dimples in the leading portion.

Abstract: A stator vane assembly includes a plurality of distinct vane segments that each respectively include a first platform, a second platform and at least one vane airfoil connected at opposed ends thereof to the first platform and the second platform. The first platforms meet at distinct first joints with each other, and the second platforms meet at distinct second joints with each other such that the plurality of vane segments forms an annular structure.

Abstract: A low pressure turbine engine component for use in an engine, for propelling a vehicle such as an aircraft is formed by a disk portion and a plurality of low pressure turbine blades extending outwardly from the disk portion. Each of the low pressure turbine blades has an airfoil portion with an axial chord length and a trailing edge. The low pressure turbine blades are spaced apart so that there is a pitch-to-chord ratio greater than 1.4, wherein the pitch is a distance between the trailing edges of adjacent ones of the low pressure turbine blades and the chord is the axial chord length of the blades.

Abstract: A component for a gas turbine engine includes a platform that has a radially inner side and a radially outer side. A root portion extends from the radially inner side of the platform. An airfoil extends from the radially outer side of the platform. The airfoil includes a pressure side that extends between a leading edge and a trailing edge. A suction side extends between the leading edge and the trailing edge. A curvature inflection point is located between 30% and 70% of an axial chord length of the airfoil.

Abstract: An inlet guide vane having varied trailing edge geometry is provided. The inlet guide vane may have an inner diameter edge opposite an outer diameter edge, and may comprise a leading edge part and a trailing edge part. The trailing edge part may comprise a trailing edge length defining a length between a forward edge and an aft edge. The trailing edge length may vary in length. The trailing edge part may be displaced from the leading edge part at a stagger angle. The stagger angle may vary in displacement.

Abstract: A gas turbine engine component is provided. The gas turbine engine component comprises a main body having a leading edge and a leading edge wall including an elongated transition portion extending between the leading edge and a proximate flowpath surface of the main body. The elongated transition portion has an axial length that is greater than a radial height. A gas turbine engine is also provided.

Abstract: A gas turbine engine has a fan rotor, a first compressor rotor and a second compressor rotor. The second compressor rotor compresses air to a higher pressure than the first compressor rotor. A first turbine rotor drives the second compressor rotor and a second turbine rotor. The second turbine drives the compressor rotor. A fan drive turbine is positioned downstream of the second turbine rotor. The fan drive turbine drives the fan through a gear reduction. The first compressor rotor and second turbine rotor rotate as an intermediate speed spool. The second compressor rotor and first turbine rotor together as a high speed spool. The high speed spool, the intermediate speed spool, and the fan drive turbine configured to rotate in the same first direction.

Abstract: A rotor blade for a gas turbine engine includes a platform section between a root section and an airfoil section, the platform section having a non-axisymmetric surface contour.

Abstract: A turbomachine includes an annular flow path section between a plurality of radially extending stator vanes and a plurality of radially extending rotor blades. At least a first portion of the flow path section has a circumferentially varying outer periphery which includes a recess circumferentially alternating with axisymmetric surfaces.

Abstract: One exemplary embodiment of this disclosure relates to a system including an airfoil having a static portion, a moveable portion, and a seal between the static portion and the moveable portion. The seal is moveable separate from the static portion and the moveable portion.

Abstract: A gas turbine engine component is provided. The gas turbine engine component comprises a main body having a leading edge and a leading edge wall including an elongated transition portion extending between the leading edge and a proximate flowpath surface of the main body. The elongated transition portion has an axial length that is greater than a radial height. A gas turbine engine is also provided.

Abstract: An array of components in a gas turbine engine includes first and second structures respectively including first and second surfaces that are arranged adjacent to one another to provide a gap. The first and second surfaces respectively have first and second rounded edges at the gap that are arranged in staggered relationship relative to one another.

Abstract: An airfoil assembly has at least one cooling hole in an aft edge of at least one platform for cooling at least one of an axially downstream airfoil root and/or tip region. The airfoil assembly may be a high pressure turbine first stage vane coupled with a combustor operating at a low Pattern Factor.

Abstract: An airfoil is disclosed comprising a distal edge, a proximal edge, a pressure side surface extending between the distal edge and the proximal edge and between a leading edge and a trailing edge, the leading edge forward of the trailing edge, wherein a first cooling hole and a second cooling hole are disposed in the pressure side surface, wherein the first cooling hole is oriented at a first angle relative to the distal edge and the second cooling hole is oriented at a second angle relative to the proximal edge, wherein the first angle and the second angle have bilateral symmetry about a plane, wherein the plane is located at a point at least one of more than or less than the midpoint between the distal edge and proximal edge.

Abstract: A turbomachine includes an annular flow path section between a plurality of radially extending stator vanes and a plurality of radially extending rotor blades. At least a first portion of the flow path section has a circumferentially varying outer periphery which includes a recess circumferentially alternating with axisymmetric surfaces.