A stator vane assembly for a compressor of a gas turbine, in particular of an aircraft engine, including a plurality of stator vanes whose airfoil sections form a stagger angle with an axis of rotation of the compressor, which stagger angle varies along a duct height of the stator vane assembly. Along the duct height from the inside to the outside, the stagger angle increases to a local maximum in a second section adjoining a first, radially innermost section, and decreases to an outer local minimum in a third section adjoining this second section and, along the duct height from the inside to the outside, the stagger angle decreases from the initial value to an inner local minimum in the first, radially innermost section and/or increases from the outer local minimum to a final value in a fourth, radially outermost section adjoining the third section.

A system includes an airfoil disposed inside an engine assembly that includes a pressure side and a suction side that are coupled together at a leading edge and a trailing edge. The airfoil extends a radial length away from a central axis of the engine assembly between a hub end and a tip end. The airfoil includes a sweep feature disposed at the leading edge that is shaped to alter the air inside the engine assembly. Altering the air inside the engine assembly reduces a surface unsteady pressure level on the airfoil. The system includes a fan frame assembly comprising an inner and outer surface. The hub end is coupled with the inner surface and the tip end is coupled with the outer surface. The airfoil is integrated with the fan frame assembly such that the airfoil increases a structural load supporting capability of the fan frame assembly.

The present disclosure relates to a support frame for a ventilation device for cooling a fluid which passes through a cooling circuit of a motor vehicle. The support frame comprises an opening, which is designed to receive a propeller, and a central support, which is positioned in a center of the opening. The central support is formed to receive a motor which activates the propeller, so as to generate a ventilation flow. The central support is attached through the opening to the support frame by retention arms. The retention arms extend according to a curved form, between the central support and a periphery of the opening.

The present invention relates to a blade row group arrangeable in a main flow path of a fluid-flow machine and including N adjacent member blade rows firmly arranged relative to one another in both the meridional direction and the circumferential direction, with the number N of the member blade rows being greater than/equal to 2 and (i) designating the running index with values between 1 and N, with a front member blade row (i) as well as a rear member blade row (i+1) being provided, and with the blade row group having two main flow path boundaries. It is provided that there is a gap between one blade end of at least one blade of at least one of the member blade rows and at least one of the two main flow path boundaries.

The present disclosure provides an axial flow compressor comprising a high pressure compressor section having a core flow path, an aft stage and a forward stage; a diffuser in fluid communication with the core flow path and coupled to the aft stage; a plenum coupled to the diffuser; a pre-swirl nozzle coupled to the plenum, an exit of the pre swirl nozzle being directed at an aft stage rotor disk and configured to impart a swirl to a cooling fluid. The axial flow compressor further may further comprise an aft stage rotor cavity defined by a portion of the aft stage rotor disk and having an aft stage axial overlap seal, wherein a portion of the cooling fluid returns to the core flow path though the aft stage labyrinth seal. The present disclosure provides a method of high pressure compressor aft stage cooling.

An aircraft gas turbine engine nacelle comprises a thrust reversal arrangement. The thrust reversal arrangement comprises at least first and second circumferentially spaced fixed thrust reverser cascade boxes each comprising a plurality of thrust reverser vanes configured to direct air forwardly and circumferentially and at least one inter-leaved translating circumferential turning vane configured to direct air in a direction having a circumferential component. The circumferential turning vane is moveable from a stowed position provided between the first and second circumferentially spaced thrust reverser cascade boxes, and a deployed position axially rearwardly of the thrust reverser cascade boxes.

According to one aspect of the present disclosure, a gas turbine engine is disclosed that includes an engine section comprising a plurality of stages of variable vanes, and also includes first and second synchronizing rings (sync-rings). Movement of the first sync-ring adjusts vane angles of a first one of the stages of variable vanes, and movement of the second sync-ring adjusts vane angles of a second one of the stages of variable vanes. At least one sensor is configured to measure a condition of the gas turbine engine. A controller is configured to move the first sync-ring independently of the second sync-ring based on data from the at least one sensor.

A guide vane segment for a turbomachine stage that has an inner ring segment (10) and a plurality of guide vanes (21-25) that are configured on the inner ring segment; at and/or in at least one detuning region (A; B) of the inner ring segment, that extends circumferentially in a portion of the inner ring segment that, from one end face (11; 12) of the inner ring segment to an inner guide vane (22; 24) that is circumferentially adjacent to an outermost guide vane (21; 25) adjacent to the end face and/or extends at most over an outermost third of a length (L) of the inner ring segment bounded by an end face (11; 12) of the inner ring segment, at least one cavity (112) is configured which contains at least one impulse element (100) with clearance of motion for providing impact contacts.

Provided are a blower, capable of suppressing noise occurring in a stator while significantly improving blowing efficiency, and an outdoor unit using the same. The present disclosure comprises: a bell mouth part spaced apart at a predetermined distance in the radial direction with respect to an outer circumferential end of a propeller fan; and a diffuser part installed on the downstream side of the bell mouth part, and having a flow path area which is enlarged from the upstream side toward the downstream side with a larger magnification rate than the magnification rate of the flow path area in the downstream end of the bell mouth part; and a stator part having a plurality of stators, wherein the stator part is arranged within the diffuser part.

A fan module for a gas turbine engine is disclosed herein. The fan module includes a fan, a plurality of outlet guide vanes, and a fluid cooling system. The fan is adapted to rotate about a central axis to pass air at least in part aftward along the central axis and around an engine core of the gas turbine engine. The outlet guide vanes are spaced aft of the fan along the central axis and configured to receive the air passed aftward along the central axis by the fan. The fluid cooling system is configured to transfer heat from a fluid to the air from the fan to cool the fluid.