The invention relates to a recirculation device for a gas of a process device, said recirculation device comprising a recirculation pump, wherein the recirculation pump is a side channel pump.

An apparatus for diagnosing and controlling the aerodynamic stability of a compressor and method there of are provided. The apparatus includes a measurement device (100), a signal processing device (200) and a control and execution device (300), wherein the measurement device (100) is configured to measure the pressure or velocity fluctuations of air flows in different positions inside a compressor in real time, and to transmit real-time measurement signals obtained from different positions to the signal processing device (200); the signal processing device (200) is configured to determine, according to the real-time measurement signals, a type and spatial distribution of instability precursor in the compressor, and to output corresponding control strategy signals to the control and execution device (300); and the control and execution device (300) executes, according to the received control strategy signals, corresponding control actions to regulate the stability of the compressor (S3).

This disclosure is related to a thin type counter-rotating axial air moving device. The ratio of the front hub diameter to the front blade diameter is about 0.3 to about 0.85. The front average pitch angle of the front blades is greater than about 46 degrees. The ratio of the rear hub diameter to the rear blade diameter is about 0.3 to about 0.85. The rear average pitch angle of the rear blades is less than about 38 degrees. The ratio of the total thickness to the greater one between the front blade diameter and the rear blade diameter is less than or equal to about 0.75.

A compressor section of a gas turbine engine includes an upstream portion and a downstream portion. The upstream portion includes at least one stage of stator vanes and at least one stage of blades configured to rotate about an axial centerline of the compressor section. The at least one stage of stator vanes and the at least one stage of blades are in an alternating arrangement along an axial direction of the gas turbine engine. The downstream portion is disposed immediately adjacent to and downstream along the axial direction from the upstream portion. The downstream portion includes a first set of rotating blade rows and a second set of rotating blade rows. The first and second sets of rotating blade rows are in an alternating arrangement along the axial direction of the gas turbine engine. The first and second sets of rotating blade rows are in a counter-rotating arrangement.

A rotor blade includes an airfoil having an airfoil shape. The airfoil shape has a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in one of Table I, Table II, Table III, Table IV, Table V, Table VI, Table VII, Table VIII, or Table IX. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances expressed in a unit of distance by multiplying the Cartesian coordinate values of X, Y and Z by a scaling factor of the airfoil in the unit of distance. The X and Y values, when connected by smooth continuing arcs, define airfoil profile sections at each Z value. The airfoil profile sections at Z values are joined smoothly with one another to form a complete airfoil shape.

A counter-rotating fan, comprising an impeller assembly and an air guide structure. The impeller assembly comprises a first stage impeller and a second stage impeller, of which the rotation directions are opposite. The pressure surfaces of first blades of the first stage impeller are configured to be opposite the suction surfaces of second blades of the second stage impeller, and from the blade root to the blade tip, each of the first blades and the second blades bends toward its own rotation direction. The air guide structure comprises a flow guide cover. The flow guide cover is provided at the center position of the air intake side of the first stage impeller, and the air intake side surface of the flow guide cover at least partially forms a flow guide surface, the flow guide surface extending along the axis of the first stage impeller in the direction away from the counter-rotating fan.

A transonic compressor of a turbomachine comprising exactly two or exactly three annular rows of rotor blades and respectively exactly one or exactly two annular rows of intermediate stator vanes interposed between two respective rows of rotor blades, wherein the flow velocity relative to the rotor downstream of the intermediate stator vanes is less than or equal to a Mach number of 0.9 over a radial portion of the blades extending over the radially inner 40% of the blades, and less than or equal to a Mach number of 1 over a radial portion extending over the radially inner 80% of the blades, and less than or equal to a Mach number of 1.05 over a radial portion extending over the radially outer 20% of the blades.

A heat exchanger (HEX) for cooling air in a gas turbine engine is provided. The HEX may comprise a central manifold comprising an inlet portion, a first outlet portion, and a second outlet portion. The HEX may further comprise a plurality of tubes coupled to the central manifold, the plurality of tubes comprising at least a first tube, a second tube, a third tube, and a fourth tube, a shroud at least partially encasing said plurality of tubes, and a cooling air flow path defined by at least one of the shroud, the plurality of tubes, and an outer surface of the central manifold, wherein the cooling air flow path is orthogonal to said plurality of tubes.

A variable guide vane (VGV) described herein includes an airfoil for interacting with a fluid inside a gas path of a gas turbine engine. The airfoil is mounted to a button and rotatable with the button about an axis. The button includes a platform surface defining part of the gas path adjacent the airfoil during use. The platform surface of the button includes a depression for receiving therein part of an adjacent VGV and providing clearance between adjacent VGVs at aggressive vane angles.

A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor including a circumferential array of blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area and a low pressure turbine section. The low pressure turbine section includes a maximum gas path radius, the blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades is equal to or greater than 0.35, and is less than 0.55.