The gas turbine compressor for an aircraft gas turbine engine includes a compressor rotor having a plurality of compressor blades circumferentially distributed around a hub. Each of the compressor blades has an airfoil extending radially outward from the hub to a blade tip. A circumferential row of the compressor blades includes two or more different blade types, at least one modified blade of the two or more different blade types having means for generating different shock patterns between adjacent ones of the two or more different blade types when the gas turbine compressor operates in supersonic flow regimes. The means for generating different shock patterns on the modified blade aerodynamically mistune the two or more different blade types.

The invention relates to a method (100) for manufacturing a turbomachine fan comprising a plurality of blades mounted on a disc extending along a longitudinal axis, said method comprising the following steps of: measuring (102) at least one structural parameter of each of the blades in the cold state, and, for each of the blades, estimating (103) at least one structural parameter in operation relating to a blade from the structural parameter(s) measured on said blade in the cold state, determining (104) an optimal sequence of the blades around the disc from the structural parameters estimated in operation for each of the blades, and mounting (105) the blades on the disc in the optimal sequence thus determined.

A compressor rotor blade for an axial-type compressor has a blade profile having a transonic section and a profile section which extends in the transonic section and has concave and convex suction-side regions on the suction side, the convex suction-side region arranged downstream of the concave suction-side region, and has convex and concave pressure-side regions on the pressure side, the concave pressure-side region arranged downstream of the convex pressure-side region. Curvature progressions on the pressure side and on the suction side are both applied in a continuous manner over a profile chord of the profile section, the positions of the minimum values of the curvature progressions deviate from each other by not more than 10% of the length of the profile chord, and the positions of the maximum values of the curvature progressions deviate from each other by not more than 10% of the length of the profile chord.

A component system of a gas turbine engine including: a first component having an outer surface; a second component having an outer surface, the second component and the first component being in a facing spaced relationship defining an air passageway therebetween; and a first recess located in at least one of the outer surface of the first component proximate the air passage and the outer surface of the first component proximate the air passage, wherein the first recess is located proximate a throat within the air passageway stretching between the first component and the second component.

A compressor rotor, such as a fan, for a gas turbine engine is described which includes alternating at least first and second blade types. The leading edge of the second blade types includes a leading edge tip cutback extending to the blade tip thereof. The leading edge tip cutback of the second blade type defines a chord length at the blade tip of the second blade types that is less than that of the first blades types. The first and second blade types generate different shock patterns when the fan or compressor rotor operates in supersonic flow regimes.

To provide a blade of a fan or compressor that is reduced in loss by enlarging a laminar flow region over a blade surface. A blade according to the present disclosure is divided into a subsonic region where the relative Mach number of the inlet air flow during rated operation of a turbofan engine is lower than 0.8 and a transonic region where the relative Mach number is equal to or higher than 0.8. Provided that a parameter () defined according to =(in)/(inex) is referred to as a blade surface angle change rate where denotes an angle formed by a tangent to the blade surface and the axial direction of the turbofan engine, in denotes the blade surface angle at the leading edge of the blade, and the ex denotes the blade surface angle at the trailing edge, in each of the subsonic region and the transonic region, the minimum value of the blade surface angle change rate on the pressure surface, an upper limit value of the blade surface angle change rate at a predetermined axial location along the chord on the pressure surface, and an upper limit value and a lower limit value of the blade surface angle change rate at a predetermined axial location along the chord on the suction surface are defined.

An airfoil section of a blade for an open rotor includes: a pressure side and a suction side, the pressure side and the suction side intersecting at a leading edge and a trailing edge, wherein a chord of the airfoil section is defined as a straight-line distance between the leading edge and the trailing edge; the airfoil section has a meanline defined midway between the pressure side and the suction side; and the meanline is shaped such that, in the presence of predetermined transonic or supersonic relative velocity conditions, maximum and minimum ideal Mach numbers on the suction side will lie within a 0.08 band, between 25% and 80% percent of the chord.

A compressor rotor blade for an axial-type compressor has a blade profile having a transonic section and a profile section which extends in the transonic section and has concave and convex suction-side regions on the suction side, the convex suction-side region arranged downstream of the concave suction-side region, and has convex and concave pressure-side regions on the pressure side, the concave pressure-side region arranged downstream of the convex pressure-side region. Curvature progressions on the pressure side and on the suction side are both applied in a continuous manner over a profile chord of the profile section, the positions of the minimum values of the curvature progressions deviate from each other by not more than 10% of the length of the profile chord, and the positions of the maximum values of the curvature progressions deviate from each other by not more than 10% of the length of the profile chord.

An airfoil section of a blade for an open rotor includes: a pressure side and a suction side, the pressure side and the suction side intersecting at a leading edge and a trailing edge, wherein a chord of the airfoil section is defined as a straight-line distance between the leading edge and the trailing edge; the airfoil section has a meanline defined midway between the pressure side and the suction side; and the meanline is shaped such that, in the presence of predetermined transonic or supersonic relative velocity conditions, maximum and minimum ideal Mach numbers on the suction side will lie within a 0.08 band, between 25% and 80% percent of the chord.

The invention relates to a method (100) for manufacturing a turbomachine fan comprising a plurality of blades mounted on a disc extending along a longitudinal axis, said method comprising the following steps of: measuring (102) at least one structural parameter of each of the blades in the cold state, and, for each of the blades, estimating (103) at least one structural parameter in operation relating to a blade from the structural parameter(s) measured on said blade in the cold state, determining (104) an optimal sequence of the blades around the disc from the structural parameters estimated in operation for each of the blades, and mounting (105) the blades on the disc in the optimal sequence thus determined.