The field of the present invention is compression or pressurization method and apparatus of rotary continuous-flow type for use with elastic fluids such as air.
More particularly, the present invention is concerned with turbomachinery compressor method and apparatus of a type having characteristics both of known axial-flow and known centrifugal-flow types, but differing quite remarkably in structure and method of operation from either of these known turbomachinery types. Consequently, the present invention is related in a general way to known turbomachinery compressor method and apparatus commonly grouped under the genus of mixed-flow axial-centrifugal type.
The present invention is also related to a combustion turbine engine employing turbomachinery compressor method and apparatus of the type described above.
The cost and reliability of modern combustion turbine engines are both strongly affected by the number of compression stages, blade rows, or acceleration/diffusion operations in the compressor sections of these engines. Accordingly, reducing the number of compressor stages has been a long-recognized objective in the field of turbomachinery design, and particularly in the jet propulsion field.
The conventional way to achieve a reduction in compressor stages has been to use one or more centrifugal-flow compressor stages in place of a greater number of axial-flow compressor stages. Centrifugal compressor stages in comparison with axial-flow compressor stages are recognized as offering a lower cost and higher static pressure ratio. They have also been recognized as offering superior resistance to damage from ingestion of foreign objects (hereinafter, foreign object damage, FOD) and superior tolerance to distortion or nonuniformity of inlet air flow distribution. However, centrifugal compressors are in general slightly less efficient and have a larger outer diameter than comparable axial flow compressor.
Balancing all these factors, early developments of jet engines for aircraft uses concentrated on axial-flow compressor stages and avoided centrifugal compressor designs primarily because of the adverse engine envelope or increased frontal area which would have resulted from the use of centrifugal compressor stages. Such increased envelope of centrifugal compressors is attributable primarily to the substantial radius change in the rotor of the centrifugal compressor stage. This radius change results in an outlet air flow having, in addition to a substantial tangential velocity component, a high radially outward velocity component. Conventionally, this high radially outward air flow velocity component dictated a stationary diffuser disposed annularly around and radially outwardly of the compressor rotor. It is this diffuser structure primarily which results in the comparatively large outer diameter of centrifugal compressors.
The theoretical possibility of structuring the rotor of a centrifugal compressor with an outlet portion turning the outlet flow toward an axial direction has been recognized in the pertinent art for many years. Such a rotor construction would allow the diffuser structure to be disposed axially of the rotor rather than radially outwardly thereof and would result in a decreased overall outer diameter. Such compressors are depicted by the U.S. Pat. Nos. 2,570,081; to B. Szczeniowski; and 2,648,492; 2,648,493; to E.A. Stalker. However, it has been learned from practical experience that substantial turning of a centrifugal compressor flow from radially outwardly toward the axial direction within the rotor itself as taught by these patents occasions such large aerodynamic losses that these designs are unattractive by contemporary performance standards.
Another alternative proposal has been to structure a compressor rotor according to centrifugal-flow teachings, but with the air flow through the rotor turning only partially toward the true radial direction despite enjoying a significant increase of radial dimension in traversing the rotor. The flow from such a mixed axial-centrifugal rotor is then received by a modified channel or pipe diffuser which initially turns the flow from axially and radially outward to, or past, the axial direction to flow axially, and perhaps radially inwardly, all substantially without diffusion. The diffuser also includes divergent pipe diffuser channels which extend a considerable distance in the downstream axial direction, and which thus contribute to an undesirably long axial dimension for such a compressor stage. U.S. Pat. No. 2,609,141, of G. Aue proposes a mixed-flow compressor of the above-described type wherein it is proposed the modified channel pipe diffuser may relieve only the radially outward, or both radially outward and tangential components of air flow velocity exiting from the rotor. However, practical experience has again shown that the radially outward component of air flow leaving such a proposed rotor is of sufficient magnitude that when the modified channel pipe diffuser is configured to relieve only this radially outward component, performance of the compressor is unacceptably low by contemporary standards. Configuration of the channel pipe diffuser to relieve both radial and tangential velocity components of air flow from the compressor rotor further increases the performance shortfall of such a compressor by current standards.
Yet another theoretical proposal has been to structure a compressor with what is essentially an axial-flow rotor having an increase of radial dimension from inlet to outlet, at least with respect to the mean radius of bulk flow through the rotor. In theory, such a compressor rotor enjoys, at least to some small degree, the advantages which centrifugal compressor rotors derive from their increase of radial dimension from inlet to outlet. Such a compressor is proposed by the U.S. Pat. No. 2,806,645, to E.A. Stalker. Again, practical experience has shown such a proposed compressor to be theoretically unsound and to offer performance far short of contemporary standards.
USPTO Backgrounds
USPTO Backgrounds