Abstract: A gas turbine engine includes a propulsor, a compressor section and a turbine section, at least one of the propulsor, compressor section and turbine section having at least one row of blades mounted in a rotor. There is a plurality of circumferentially spaced blades in the at least one row. The plurality of blades in the at least one row each having a platform with an airfoil extending radially outwardly of the platform. The airfoils in a first group of said plurality of rotating blades has a different shape than the airfoils in a second group of the plurality of blades.

Abstract: An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, major portions of leading edges of the downstream stator vanes circumferentially overlapped by the upstream stator vanes, the downstream stator vanes including: a first pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a first pitch, and a second pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a second pitch different than the first pitch.

Abstract: An aircraft engine, has: an upstream stator having upstream stator vanes distributed about a central axis; and a downstream stator having downstream stator vanes distributed about the central axis, the downstream stator located downstream of the upstream stator, a number of the upstream stator vanes different than a number of the downstream stator vanes, the downstream stator vanes including: a first vane, a major portion of a leading edge of the first vane circumferentially overlapped by one of the upstream stator vanes; and a second vane differing from the first vane by a geometric parameter, the geometric parameter causing the second vane to have one or more of: a stiffness greater than that of the first vane, and a major portion of a leading edge of the second vane circumferentially overlapped by another one of the upstream stator vanes.

Abstract: An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, major portions of leading edges of the downstream stator vanes circumferentially overlapped by the upstream stator vanes, the downstream stator vanes including: a first pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a first pitch, and a second pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a second pitch different than the first pitch.

Abstract: Assemblies and methods for directing a flow of air through a compressor section of a gas turbine engine are disclosed. A method includes receiving the flow of air over a strut extending radially across a substantially annular gas path of the gas turbine engine, and at least partially confining and modifying a strut wake generated in the flow of air by the strut using one or more variable orientation guide vanes.

Abstract: An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, the downstream stator vanes including: a first vane made of a first material, a major portion of a leading edge of the first vane circumferentially overlapped by one of the upstream stator vanes, and a second vane made of a second material having a greater stiffness, strength, and/or ductility than that of the first material, a major portion of a leading edge of the second vane exposed via a spacing defined between two of the upstream stator vanes.

Abstract: An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, the downstream stator vanes including: a first vane made of a first material, a major portion of a leading edge of the first vane circumferentially overlapped by one of the upstream stator vanes, and a second vane made of a second material having a greater stiffness, strength, and/or ductility than that of the first material, a major portion of a leading edge of the second vane exposed via a spacing defined between two of the upstream stator vanes.

Abstract: An aircraft engine, has: an upstream stator having upstream stator vanes distributed about a central axis; and a downstream stator having downstream stator vanes distributed about the central axis, the downstream stator located downstream of the upstream stator, a number of the upstream stator vanes different than a number of the downstream stator vanes, the downstream stator vanes including: a first vane, a major portion of a leading edge of the first vane circumferentially overlapped by one of the upstream stator vanes; and a second vane differing from the first vane by a geometric parameter, the geometric parameter causing the second vane to have one or more of: a stiffness greater than that of the first vane, and a major portion of a leading edge of the second vane circumferentially overlapped by another one of the upstream stator vanes.

Abstract: An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, major portions of leading edges of the downstream stator vanes circumferentially overlapped by the upstream stator vanes, the downstream stator vanes including: a first pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a first pitch, and a second pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a second pitch different than the first pitch.

Abstract: A method includes: obtaining a rotor having a hub and a plurality of blades protruding from the hub, the plurality of blades including first blades and second blades disposed in alternation around a central axis of the rotor, natural vibration frequencies of the first blades different from natural vibration frequencies of the second blades; determining that a difference between a first natural vibration frequency of a first blade of the first blades and a second natural vibration frequency of a second blade of the second blades is below a threshold; and modifying a shape of the first blade until the difference between the first natural vibration frequency and the second natural vibration frequency is at or above the threshold.

Abstract: Assemblies and methods for directing a flow of air through a compressor section of a gas turbine engine are disclosed. A method includes receiving the flow of air over a strut extending radially across a substantially annular gas path of the gas turbine engine, and at least partially confining and modifying a strut wake generated in the flow of air by the strut using one or more variable orientation guide vanes.

Abstract: Device and methods for guiding bleed air in a turbofan gas turbine engine are disclosed. The devices provided include louvers and baffles that guide bleed air toward a bypass duct of the turbofan engine. The louvers and baffles have a geometric configuration that promotes desirable flow conditions and reduced energy loss.

Abstract: Device and methods for guiding bleed air in a turbofan gas turbine engine are disclosed. The devices provided include louvers and baffles that guide bleed air toward a bypass duct of the turbofan engine. The louvers and baffles have a geometric configuration that promotes desirable flow conditions and reduced energy loss.

Abstract: Device and methods for guiding bleed air in a turbofan gas turbine engine are disclosed. The devices provided include louvers and baffles that guide bleed air toward a bypass duct of the turbofan engine. The louvers and baffles have a geometric configuration that promotes desirable flow conditions and reduced energy loss.

Abstract: A stator that has vanes extending between a first end and a second end along a span and from a leading edge to a trailing edge along a chord length, the vanes having a first end portion extending from the first end to about 30% of the span to a first location, a chord ratio of the chord length at the first end to the chord length at the first location greater than or equal to 1.1, a throat ratio of a width of a throat between two adjacent vanes at the first location to a width of the throat at the first end is greater than or equal to 1.3, a sweep angle difference between a maximum sweep angle of the leading edge along the first end portion and a minimum sweep angle of the leading edge along the first end portion is at least 15 degrees.

Abstract: Assemblies and methods for directing a flow of air through a compressor section of a gas turbine engine are disclosed. A method includes receiving the flow of air over a strut extending radially across a substantially annular gas path of the gas turbine engine, and at least partially confining and modifying a strut wake generated in the flow of air by the strut using one or more variable orientation guide vanes.

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 stator that has vanes extending between a first end and a second end along a span and from a leading edge to a trailing edge along a chord length, the vanes having a first end portion extending from the first end to about 30% of the span to a first location, a chord ratio of the chord length at the first end to the chord length at the first location greater than or equal to 1.1, a throat ratio of a width of a throat between two adjacent vanes at the first location to a width of the throat at the first end is greater than or equal to 1.3, a sweep angle difference between a maximum sweep angle of the leading edge along the first end portion and a minimum sweep angle of the leading edge along the first end portion is at least 15 degrees.

Abstract: There is disclosed a fan assembly including a fan rotor including a hub and fan blades. The fan blades have a leading edge and a trailing edge. A fan stator downstream of the fan rotor relative to a direction of an airflow through the fan assembly. The fan stator includes vanes extending between radially inner ends and radially outer ends. A flow recirculation circuit has an inlet downstream of radially inner ends of the vanes of the fan stator and an outlet upstream of radially inner ends of the vanes. A recirculation stator has a plurality of stationary guide vanes circumferentially distributed around the axis and located in the flow recirculation circuit between the inlet and the outlet A method of operating the fan assembly is also disclosed.

Abstract: There is disclosed a fan assembly including a fan rotor including a hub and fan blades. The fan blades have a leading edge and a trailing edge. A fan stator downstream of the fan rotor relative to a direction of an airflow through the fan assembly. The fan stator includes vanes extending between radially inner ends and radially outer ends. A flow recirculation circuit has an inlet downstream of radially inner ends of the vanes of the fan stator and an outlet upstream of radially inner ends of the vanes. A recirculation stator has a plurality of stationary guide vanes circumferentially distributed around the axis and located in the flow recirculation circuit between the inlet and the outlet A method of operating the fan assembly is also disclosed.