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).

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.

The present invention relates to a compressor stage for a gas turbine, in particular, an aircraft engine, having a row of rotating blades (3) and a row of guide vanes (4), which is adjacent downstream, wherein the choke point and the aspect ratio AR.sub.ax, which is defined by the quotient between average channel height (h) and average chord length (l.sub.ax), satisfy the condition

>1.33AR.sub.ax+5.16.

A flow path surface of a gas turbine engine at the location of a bladed component is disclosed in which the flow path surface includes a cylindrical upstream side and a conical downstream side. The bladed component is located at the intersection of the cylindrical upstream side and the conical downstream side. The cylindrical upstream side can extend from a leading edge of the bladed component, or a point upstream of it, to a location between the leading edge and trailing edge of the component. The conical downstream side can extend past the trailing edge of the bladed component. The bladed component can be a fan blade or a compressor blade.

Blowers, e.g., for use in outdoor applications, are provided. A blower includes an air inlet; an air outlet; and a fan assembly disposed between the air inlet and the air outlet. The fan assembly has a multi-stage fan. The multi-stage fan includes a plurality of axial fans mounted on a drive shaft and a plurality of stator fans arranged in series with the axial fans, wherein each axial fan comprises a plurality of blades each having a blade tip. Each stator fan includes a stackable housing segment. Each of the stackable housing segments are coupled together in series to form the blower housing.