Abstract: A rotary airfoil in a gas turbine engine is provided. The airfoil includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. An axial component of a center of gravity of a cross-section taken chordally toward the tip of the airfoil being upstream relative to an axial component of a center of gravity of a cross-section taken chordally toward the root of the airfoil. A method for forming such blade is also presented.

Abstract: A rotor for a gas turbine engine comprises a rotor having a hub and blades around the hub, and extending from the hub to tips. The tips include first and second tip portions between their respective tip leading edge and tip trailing edge. Tips are spaced from a rotational axis of the rotor by spans. A mean span of a first tip portion of a first blade is greater than a mean span of a corresponding first tip portion of a second blade. A mean span of a second tip portion the first blade is less than a mean span of a corresponding second tip portion of the second blade.

Abstract: A gas turbine engine casing apparatus includes annular first and second annular cases connected by a spigot joint. The spigot joint includes an annular projection of the first annular case fitted into an annular recess of the second annular case. A plurality of circumferentially spaced apart scallops are formed on one of surfaces of the annular projection or of the annular recess, and are located in selective circumferential locations adjacent respective enhanced stiff areas of the first and second annular cases.

Abstract: A fan for a gas turbine engine comprises blades distributed around a hub. The blades include first and second blades, having geometric parameters and/or material properties that differ from each other to frequency mistune the fan. The blades are distributed about the hub with at least one second blade between adjacent first blades. The leading edge of the airfoil of the second blades is disposed axially aft of the corresponding leading edge of the first blades in at least a portion of the blade span.

Abstract: A rotor for a gas turbine engine. The rotor includes blades circumferentially distributed around a hub. The blades have airfoils with a span defined between a root and tip, a chord defined between a leading edge and a trailing edge, and a thickness defined between a pressure side surface and suction side surface. The blades include first blades and second blades. The airfoil of the first blades has a first thickness distribution defining a first natural vibration frequency of the airfoils of the first blades. The airfoil of the second blades has a second thickness distribution defining a second natural vibration frequency different than the first natural vibration frequency. The first thickness distribution is different than the second thickness distribution along a radially-inner half of the span, and the first thickness distribution matches the second thickness distribution along a radially-outer half of the span.

Abstract: A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.

Abstract: A compressor rotor for a gas turbine engine includes a hub disposed about an axis of rotation and an outer surface forming a radially inner gaspath boundary, the outer surface defining a nominal hub diameter. A circumferential array of blades extends radially outwardly from the hub. A first inter-blade passage is defined between a first set of adjacent blades and has a first throat area. A second inter-blade passage is defined between a second set of adjacent blades and has a second throat area that is smaller than the first throat area. At least one scoop is disposed in the second inter-blade passage, the scoop defining a cavity extending radially into the outer surface of the hub relative to the nominal hub diameter.

Abstract: A rotor for a gas turbine engine comprises a rotor having a hub and blades around the hub, and extending from the hub to tips. The tips include first and second tip portions between their respective tip leading edge and tip trailing edge. Tips are spaced from a rotational axis of the rotor by spans. A mean span of a first tip portion of a first blade is greater than a mean span of a corresponding first tip portion of a second blade. A mean span of a second tip portion the first blade is less than a mean span of a corresponding second tip portion of the second blade.

Abstract: A gas turbine engine casing apparatus includes annular first and second annular cases connected by a spigot joint. The spigot joint includes an annular projection of the first annular case fitted into an annular recess of the second annular case. A plurality of circumferentially spaced apart scallops are formed on one of surfaces of the annular projection or of the annular recess, and are located in selective circumferential locations adjacent respective enhanced stiff areas of the first and second annular cases.

Abstract: An airfoil for a gas turbine engine comprises at least three successive sections defined between a root and a tip. A first decrease of maximum thickness-to-chord ratios defined by a difference between a maximum thickness-to-chord ratio at the root and a maximum thickness-to-chord ratio at a first spanwise position, a second decrease of the maximum thickness-to-chord ratios defined by a difference between the maximum thickness-to-chord ratio at the first spanwise position and a maximum thickness-to-chord ratio at a second spanwise position, a third decrease of the maximum thickness-to-chord ratios defined by a difference between the maximum thickness-to-chord ratio at the second spanwise position and a maximum thickness-to-chord ratio at the tip, the second decrease being greater than the corresponding first and third decreases.

Abstract: A compressor rotor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. A leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical. The leading edge dihedral angle has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.

Abstract: An airfoil in a gas turbine engine includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend spanwise from a root to a tip of the airfoil. The airfoil has a spanwise distribution of maximum thicknesses of chordwise cross-sections of the airfoil. The spanwise distribution of maximum thicknesses decreases from the root to the tip. In one aspect, the spanwise distribution is stepped between a first portion extending from the root and a second portion extending to the tip.

Abstract: A compressor rotor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. A leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical. The leading edge dihedral angle has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.

Abstract: A gas turbine engine casing apparatus includes annular first and second annular cases connected by a spigot joint. The spigot joint includes an annular projection of the first annular case fitted into an annular recess of the second annular case. A plurality of circumferentially spaced apart scallops are formed on one of surfaces of the annular projection or of the annular recess, and are located in selective circumferential locations adjacent respective enhanced stiff areas of the first and second annular cases.

Abstract: A gas turbine engine is described having a rotor with a plurality of rotor blades each extending radially between a blade root to a blade tip and axially between a leading edge and a trailing edge. An annular casing body housing the rotor blades includes an abradable segment of the inner surface facing the blade tips and having an abradable member. The abradable segment having one or more annular grooves, extending into the casing body from an inner surface thereof, and including an edge groove axially aligned with and facing the leading edge or the trailing edge of the rotor blades. The edge groove extends axially between a first position on the inner surface upstream of the leading edge or the trailing edge, and a second position on the inner surface downstream of the leading edge or the trailing edge.

Abstract: A gas turbine having an annular casing with a series of circumferentially spaced openings defined therethrough; a plurality of vanes extending radially inwardly though respective casing openings, an outer end of the vanes projecting radially outwardly from the casing through the respective openings and located within the respective openings by grommets, and an inner end of the vanes being mounted to an inner portion of the casing; a flexible, segmented strap extending around the annular casing, surrounding the projecting outer ends of the vanes; and a spring radially loading the flexible strap configured to apply a tension force to the flexible strap.

Abstract: An airfoil in a gas turbine engine includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend spanwise from a root to a tip of the airfoil. The airfoil has a spanwise distribution of maximum thicknesses of chordwise cross-sections of the airfoil. The spanwise distribution of maximum thicknesses decreases from the root to the tip. In one aspect, the spanwise distribution is stepped between a first portion extending from the root and a second portion extending to the tip.

Abstract: A compressor rotor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. A leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical. The leading edge dihedral angle has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.

Abstract: A rotary airfoil in a gas turbine engine is provided. The airfoil includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. An axial component of a center of gravity of a cross-section taken chordally toward the tip of the airfoil being upstream relative to an axial component of a center of gravity of a cross-section taken chordally toward the root of the airfoil. A method for forming such blade is also presented.

Abstract: A gas turbine having an annular casing with a series of circumferentially spaced openings defined therethrough; a plurality of vanes extending radially inwardly though respective casing openings, an outer end of the vanes projecting radially outwardly from the casing through the respective openings and located within the respective openings by grommets, and an inner end of the vanes being mounted to an inner portion of the casing; a flexible, segmented strap extending around the annular casing, surrounding the projecting outer ends of the vanes; and a spring radially loading the flexible strap configured to apply a tension force to the flexible strap.