A rotary apparatus for inputting thermal energy into fluidic medium is provided, the apparatus comprises: a casing with at least one inlet and at least one outlet; a rotor comprising at least one row of rotor blades configured as impulse impeller blades arranged over a circumference of a rotor hub mounted onto a rotor shaft; at least one row of stationary nozzle guide vanes arranged upstream of the at least one row of the rotor blades, respectively; and at least one row of stationary diffuser vanes arranged downstream of the at least one row of the rotor blades, respectively. The apparatus is configured to impart an amount of thermal energy to a stream of fluidic medium directed along a flow path formed inside the casing between the inlet and the outlet by virtue of a series of energy transformations occurring when said stream of fluidic medium successively passes through the blade/vane rows formed by the nozzle guide vanes, the rotor blades and the diffuser vanes, respectively, wherein, in said apparatus, a space formed between an exit from the at least one row of diffuser vanes and an entrance to the at least one row of nozzle guide vanes in a direction of the flow path formed inside the casing between the inlet and the outlet is made variable to regulate the amount of thermal energy input to the stream of fluidic medium propagating through the apparatus. Related uses and a method for inputting thermal energy into a fluidic medium are further provided.

A diffuser assembly for a compressor is provided. The diffuser may include a first stationary wall and a second stationary wall coupled with one another and forming at least in part a volute configured to receive compressed process fluid from the diffuser. The second stationary wall may define a plurality of stationary wall grooves. The diffuser may also include a moveable wall defining a plurality of moveable wall grooves. The moveable wall may be disposed between the first stationary wall and the second stationary wall such that respective stationary wall grooves and moveable wall grooves form respective flow passages configured to receive the flow of process fluid exiting an impeller of the compressor. The diffuser assembly may also include an actuator assembly configured to displace the moveable wall to alter a cross-sectional area of each of the flow passages based on a flow rate of the process fluid.

A pneumatically actuated vacuum pump is disclosed. The pneumatically actuated vacuum pump includes a body. The body defines at least two converging motive sections each having an outlet end, at least two diverging discharge sections each having an inlet end, and at least one Venturi gap. The Venturi gap is located between the outlet ends of the at least two converging motive sections and the inlet ends of the at least two diverging discharge sections.

A pneumatically actuated vacuum pump is disclosed. The pneumatically actuated vacuum pump includes a body. The body defines at least two converging motive sections each having an outlet end, at least two diverging discharge sections each having an inlet end, and at least one Venturi gap. The Venturi gap is located between the outlet ends of the at least two converging motive sections and the inlet ends of the at least two diverging discharge sections.

Flutter-resistant transonic turbomachinery blades and methods for reducing transonic turbomachinery blade flutter are provided. The flutter-resistant transonic turbomachinery blade comprises a transonic turbomachinery blade that includes opposite pressure and suction surfaces extending longitudinally in span from a root to an opposite tip, and extending axially in chord between opposite leading and trailing edges. The flutter-resistant transonic turbomachinery blade includes a local positive camber in or proximate a predicted local region of supersonic flow over the transonic turbomachinery blade. The method comprises predicting a local region of supersonic flow over the transonic turbomachinery blade and inducing the local positive camber to the transonic turbomachinery blade in or proximate the predicted region of supersonic flow.

A mixed-flow compressor includes an impeller attached to a shaft and rotatable about a shaft axis. A vaneless diffuser is located axially downstream of the impeller and has a converging portion and a diverging portion. A vaned diffuser is located axially downstream of the vaneless diffuser.

A mixed-flow compressor includes an impeller attached to a shaft and rotatable about a shaft axis. A vaneless diffuser is located axially downstream of the impeller and has a converging portion and a diverging portion. A vaned diffuser is located axially downstream of the vaneless diffuser.

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.

An impeller (60) for use in a supercharger (42), which is cable of increasing the efficiency while suppressing the increase in size of the impeller (60), is driven in driving connection with a crankshaft (26) of a combustion engine (E) mounted on a motorcycle to supply an intake air (I) towards the combustion engine (E). When the supercharger (42) is driven at a maximum permissible engine speed of an inlet diameter (Ii), the impeller (60) is so set that the peripheral velocity at an inlet side tip end portion (112) of the impeller (60) may exceed the sonic velocity, and may be smaller than 1.3 times the sonic velocity.

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.