Described is a method for operating a compressor of a turbomachine, in which, when considered in the direction of a main flow, an, in particular, radially averaged degree of reaction has dropped in a front compressor area from a maximum to a minimum, is held constant or virtually constant across a central compressor area up into a rear compressor area, an, in particular, radially averaged degree of reaction being adjusted in the rear compressors area which is closer to the minimum than to the maximum, and a residual swirl of at least 47 is present in the middle section, and a compressor and a turbomachine.

A pumping system includes a motor and a drive shaft configured for rotation by the motor. The pumping system includes an upper pump positioned above the motor, an upper pump shaft and an upper directional coupling connected between the drive shaft and the upper pump shaft. The upper directional coupling is configured to lock the upper pump shaft to the drive shaft when the drive shaft is rotated in a first direction. The pumping system further includes a lower pump positioned below the motor, a lower pump shaft, and a lower directional coupling connected between the drive shaft and the lower pump shaft. The lower directional coupling is configured to lock the lower pump shaft to the drive shaft when the drive shaft is rotated in a second direction.

A refrigerant compressor according to an exemplary aspect of the present disclosure includes, among other things, a first compression stage arranged in a main refrigerant flow path. The first compression stage is a mixed compression stage having both axial and radial components. The compressor further includes a second compression stage arranged in the main refrigerant flow path downstream of the first compression stage. The second compression stage is a radial compression stage.

A fan system includes a rotor having plurality of blades and a ring airfoil, the plurality of blades being rotatably joined to a hub and the ring airfoil. The fan system may include a second contra-rotationally disposed rotor having a plurality of blades and a ring airfoil. The first and second ring airfoils having a cambered shape and an angle of attack between about 5 degrees and about 45 degrees, more preferably between about 5 degrees and about 30 degrees. Optionally, an outlet guide vane may be mounted rearward of the one or more rotors having a ring airfoil.

A compressor section for a gas turbine engine according to an example of the present disclosure includes, among other things, a low pressure compressor including a plurality of rotor blades arranged about an axis, a high pressure compressor, and a core flowpath passing through the low pressure compressor. The core flowpath at the low pressure compressor defines an inner diameter and an outer diameter relative to the axis. The outer diameter has a slope angle relative to the axis.

The invention concerns a transition duct for a multi-stage compressor of a gas turbine engine. Regions of the inner surface of the duct are provided with a predetermined and dissimilar surface roughness to optimise gas flow efficiency within the duct.

A compressor for a gas turbine engine is disclosed. The compressor includes a first compression stage mounted for rotation about a central axis that includes a plurality of first-stage blades. The compressor also includes a second compression stage mounted along the central axis aft of the first compression stage to receive air compressed by the first compression stage. The second compression stage includes a plurality of second-stage blades.

A controlled convergence compressor flowpath (10) configured to better distribute the limited flowpath (10) convergence within compressors (12) in turbine engines (14) is disclosed. The compressor (12) may have a flowpath (10) defined by circumferentially extending inner and outer boundaries (16, 18) that having portions in which the rate of convergence changes to better distribute fluid flow therethrough. The rate of convergence may increase at surfaces (20, 22) adjacent to roots (24) of airfoils (26) and decrease near airfoil tips (68) and in the axial gaps (28) between airfoil rows (30). In at least one embodiment, the compressor flowpath (10) between leading and trailing edges (44, 46) of a first compressor blade (42) may increase convergence moving downstream to a trailing edge (46) of the first compressor blade (42) due to increased convergence of the inner compressor surface (22). The compressor flowpath (10) between leading and trailing edges (32, 34) of a first compressor vane (36) immediately downstream from the first compressor blade (42) may increase convergence moving downstream due to increased convergence of the outer compressor surface (20).

A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor section including a propulsor that delivers flow to a core flowpath and a compressor section including first and second compressors. The core flowpath passes through the first compressor. The core flowpath in the first compressor has an outer diameter relative to the engine longitudinal axis. The outer diameter has a slope angle relative to the axis.

A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. 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.