Abstract: A guide vane within an annular inlet duct of a gas-turbine engine provides for generating swirl within an annular inlet duct so as to provide for reducing the rate of deceleration of the inlet air flow within the annular inlet duct while providing for diffusion of the meridional component of velocity thereof.

Abstract: A diffuser (10) comprises first (10.1) and second (10.2) annular portions bounded by forward (32) and aft (26) annular walls, wherein the first annular portion (10.1) is vaneless and radially within the second annular portion (10.2), and the second annular portion (10.2) is radially relatively compact and incorporates a plurality of vanes (48) with relatively high solidity, wherein the forward (32) or/and aft (26) annular walls is/are sloped so as to provide for meridional divergence within the second annular portion (10.2) of the diffuser (10), and the vanes (48) are shaped so as to substantially conform to the flow field within the second annular portion (10.2).

Abstract: A compressor assembly is disclosed. The compressor assembly may have a compressor housing. The compressor housing may have an inner wall. The compressor assembly may also have a compressor impeller disposed within the compressor housing. Further, the compressor assembly may have a bearing housing attached to the compressor housing. The bearing housing may have a body portion and a web extending outward from the body portion to a web end. The compressor assembly may also have a diffuser ring disposed between the inner wall and the web. The diffuser ring may have at least one vane. In addition, the compressor assembly may have a vaneless space extending between the compressor impeller and the vane. The vaneless space may be inclined at an angle relative to a plane disposed orthogonal to a rotational axis of the compressor assembly.

Abstract: A diffuser (10) comprises first (10.1) and second (10.2) annular portions bounded by forward (32) and aft (26) annular walls, wherein the first annular portion (10.1) is vaneless and radially within the second annular portion (10.2), and the second annular portion (10.2) is radially relatively compact and incorporates a plurality of vanes (48) with relatively high solidity, wherein the forward (32) or/and aft (26) annular walls is/are sloped so as to provide for meridional divergence within the second annular portion (10.2) of the diffuser (10), and the vanes (48) are shaped so as to substantially conform to the flow field within the second annular portion (10.2).

Abstract: A compressor assembly is disclosed. The compressor assembly may have a compressor housing. The compressor housing may have an inner wall. The compressor assembly may also have a compressor impeller disposed within the compressor housing. Further, the compressor assembly may have a bearing housing attached to the compressor housing. The bearing housing may have a body portion and a web extending outward from the body portion to a web end. The compressor assembly may also have a diffuser ring disposed between the inner wall and the web. The diffuser ring may have at least one vane. In addition, the compressor assembly may have a vaneless space extending between the compressor impeller and the vane. The vaneless space may be inclined at an angle relative to a plane disposed orthogonal to a rotational axis of the compressor assembly.

Abstract: A rocket engine comprises first and second rotary injectors for injecting respective fuel and oxidizer propellant components into a first combustion chamber, and the effluent therefrom drives a turbine that rotates the rotary injectors. The mixture within the first combustion chamber is preferably fuel-rich so as to reduce the associated combustion temperature, and the fuel-rich effluent mixes in a second combustion chamber with additional oxidizer injected by a third rotary injector so as to generate a high temperature effluent suitable for propulsion. The rotary injectors are adapted so as to isolate the low pressure propellant supply from the relatively high pressures in the respective combustion chambers.