HIGH EFFICIENCY TRANSONIC MIXED-FLOW COMPRESSOR METHOD AND APPARATUS

Abstract: A method and apparatus for compression of elastic fluids such as air in a rotary continuous mixed flow transonic process particularly of interest in the turbomachinery field. The com­ pressor comprises a housing (12) defining an annular wall (50) in which a rotor (22) is journaled, thereby defining an annular flowpath (24). The annular wall (50) has a cylindrical inlet (52a), a part-conical outlet (52b) and a smooth transitional intermediate portion (52c). The rotor (22) has blades (42) which closely conform to the inner surface of the annular wall (50).

Description: The present invention relates to a compression or pressurisation method and apparatus in particular, those of rotary continuous-flow type for use with elastic fluids such as air. The present invention is particularly 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 types. Consequently, the present invention is related in a general way to apparatus commonly grouped under the genus of mixed-flow axial-centrifugal type. The present invention is particularly applicable to a combustion turbine engine. 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 recognised objective in the field of turbomachinery design, and particularly in the jet propulsion field. The conventional way to achieve a reduction in the number of 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 copressor stages tend to be lower in cost and provide a higher static pressure ratio. They also tend to offer superior resistance to damage from ingestion of foreign objects "foreign object damage" or FOD) and superior tolerance to distortion or nonuniformity of the inlet air flow distribution. However, centrifugal compressors are in general slightly less efficient and have a larger outer diameter than comparable axial flow compressors. 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. This increased envelope size is attributable primarily to the substantial radius change in the rotor of the centrifugal compressor stage. The 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 has necessitated a stationary diffuser disposed annularly around the compressor rotor. It is this diffuser structure primarily which has caused 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 appreciated for many years. Such a rotor construction would allow the diffuser structure to be located axially of the rotor rather than radially outside and would result in a decreased overall outer diameter. Such compressors are shown in United States Patents 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 flows radially outwardly towards the axial direction within the rotor itself as taught by these patents resulting in such large aerodynamic losses that these designs are unattractive by contemporary performance standards. An alternative proposal has been to design a compressor rotor according to centrifugal-flow teachings, but with the air flow through the rotor turning only partially towards the true radial direction despite enjoying a significant increase of radial dimension in tranversing the rotor. The flow from such a mixed axial-centrifugal rotor is then received by modified channel or pipe diffuser which initially turns the flow from axially and radially outwards 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. United States Patent 2,609,141, of G. Aue proposes a mixed-flow compressor of this type in which it is proposed that the modified channel pipe diffuser may relieve only the radially outwards, or both the radially outward and tangential components of air flow velocity leaving the rotor. However, practical experience has again shown that the radially outward component of the air flow leaving this kind of 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. The 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 design a compressor with what is essentially an axial-flow rotor having an increase in its 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 increased radial dimension from inlet to outlet. Such a compressor is proposed in United States Patent 2,806,645, to E.A. Stalker. Again, practical experience suggests that such a compressor may be theoretically unsound and offers a performance far short of contemporary standards. In view of the above, the objects broadly stated for a compressor according to this invention are to achieve a compressor envelope or outer diameter the same as, or only slightly larger than, that offered by the best conventional axial-flow compressor technology; to achieve a static pressure ratio, cost, inlet distortion tolerance and resistance to FOD substantially the same as that offered by the best conventional centrifugal flow compressor technology; and to achieve a compressor efficiency at least equal to that of the conventional centrifugal compressors, and preferably approaching the efficiency of conventional axial-flow compressors. According to the invention there is provided compressor apparatus comprising a housing and a rotor, the housing comprising an axially and circumferentially extending annular wall having an inlet and an outlet, the rotor being journaled within the housing and having a hub with a plurality of blades, the hub and annular wall together defining an annular axially extending passage, characterised in that the annular wall includes: at the inlet end, a portion which is at least up to the leading edge of the stator vane of right circular cylindrical shape in section; adjacent the rotor an outlet portion which is of right circular conical shape in section and which diverges downstream relative to the inlet, the portion diverging downstream with a divergence angle of 45 o or less and extending upstream of the virtual intersection of the trailing edges of the blades; and a smooth transitional intermediate portion between the inlet portion and conical outlet portions; and further characterised in that the hub is cone shaped and extends between said inlet and said outlet portion. Such apparatus may be considered mixed-flow transonic compressor apparatus suitable for compressible fluids. Preferably, the blades extend axially, circumferentially and radially from the hub to a position just short of the annular wall , and conform closely to the shape of the annular wall at axial positions throughout the inlet portion, the intermediate portion and the outlet portion. Preferably the hub and the annular wall define a flow path for receiving a flow stream of compressible fluid having a first vector sum of meridional velocity and tangential relative velocity of at least Mach 1.2 with respect to a selected reference adjacent the inlet portion, and for diffusion the flow stream at a second supersonic relative velocity less than the first relative velocity while limiting deviation of radially outer local relative velocity vectors to no more than 10 o with respect to the first relative velocity vector. The hub preferably defines a curvilinear cone-shaped outer surface, which is axially congruent with the housing at the inlet and outlet. The blades may extend circumferentially from the inlet end to the outlet end in a direction opposite to the rotational direction of the rotor and further comprise determined compressible fluid stream line shapes stacked substantially radially outwards from the hub outer surface towards the housing. Preferably, each blade defines a radially extending leading edge, a radially extending trailing edge, and a radially, axially and circumferentially extending radially outer tip edge comforming in shape to the annular wall at the inlet portion, the intermediate portion, and the outlet portion, the tip edges considered in axial and radial aspect defining a tip edge meridional dimension for the blades, the leading edges defining a virtual intersection with the annular wall at the inlet portion and the trailing edges defining a virtual intersection with the annular wall at the outlet portion. Preferably, the inlet portion of right circular cylindrical section extends upstream of the virtual intersection of the leading edges for a distance of from about 10% to about 20% of the tip edge meridional dimension, and the inlet portion of right circular cylindrical sectin extends downstream of the virtual intersection of the leading edges for a distance of from 10% to 30% of the tip edge meridional dimension. Preferably the circular section inlet portion and right circular conical section oulet portion of the annular wall are in axial section angularly disposed relative to one another so as to define an acute angle between them in the range from about 5 o to about 45 o preferably 22 o . The leading edges may be swept downstream radially outwardly with respect to a radially extending line from the axis to define a leading edge sweep angle. The sweep angle may be in the range of substantially 0 o to about 15 o , preferably substantially 7 o . Preferably, the conical section outlet portion of the annular wall extends upstream of the virtual intersection of the trailing edges for a distance of from about 10% to about 30% of the tip edge meridional dimension, and the conical section outlet portion of the annular wall extends downstream of the virtual intersection of the trailing edges for a distance of from about 5% to about 15% of the tip edge meridional dimension. The trailing edges may be swept upstream radially outwardly with respect to a radially extending line from the axis to define a trailing edge sweep angle. This sweep angle may be in the range of substantially 0 o to about 35 o , preferably substantially 23 o . Preferably, the leading edges define a diameter RBi at their intersection with the hub and a diameter RBo at their intersection with the tip edges, and the trailing edges define a diameter REi at their intersection with the hub and a diameter REo at their intersection with the tip edges; the ratio of REi to RBi being in the range from about 1.5 to about 3.5 and the ratio of REo to RBo is in the range from about 1.05 to about 1.76. Preferably, the ratio of REi to RBi is substantially 2.75 and the ratio of REo to RBo is substantially 1.17. Preferably the blades define a quantity termed, average meridional blade length (AMBL), which is the length in axial and radial aspect of a line along a blade from the leading edge to the trailing edge and defined by points on the blade radially midway between the hub and the tip edge, the ratio (AR) of (RBo-RBi) + (REo-REi) to AMBL lying in the range from about 0.75 to about 1.30. Preferably, the ratio AR is substantially 1.12. Preferably, the blades have a height dimension at their trailing edge defined as REo-REi, and the compressor apparatus further includes a second axially and circumferentially extending annular wall disposed radially inwardly of the first annular wall and immediately downstream of the rotor, the two annular walls being radially spaced apart to define an axially extending annular flow path downstream of the rotor, the inner wall defining a first radially outwardly convex annular surface portion bounding the flow path and being arcuate of radius R in axial section, the ratio of the trailing edge blade height to the radius R lying in the range from about 1.0 to about 4.0. Preferably the ratio of trailing edge blade height to radius R is substantially 2.0. Preferably, the first annular wall defines a radially inwardly concave annular surface portion bounding the flow path downstream of the right circular conical outlet portion, the concave annular surface portion of the first annular wall cooperating with the convex annular surface portion of the second annular wall to define the flow path downstream of the rotor the flow path extending radially outwardly and axially with a substantially constant transverse sectional fluid flow area. Preferably, the apparatus includes annular diffuser means located at or near the rotor at the outlet end which receives, via the flow path, a flow of compressible fluid having a first determined subsonic relative velocity vector sum of tangential and meridional velocity, having a respective significant radially outward component, and discharges a flow of compressible fluid having a second determined relative velocity vector having radially outward and tagential components of substantially zero (substantially pure axial flow). Preferably, the diffuser means includes successive downstream first and second axially spaced apart annular arrays of circumferentially spaced apart diffuser vanes, in which the first diffuser vanes are arranged to receive from the outlet portion, a flow of compressible fluid having the first determined relative velocity vector, and to diffuse the fluid flow to discharge into an axially extending interdiffuser space, a flow of the fluid having a second determined relative velocity vector less than the first relative velocity vector and having a radially outward component of substantially zero, and in which the second diffuser vanes are arranged to receive from the interdiffuser space the flow of fluid having the second determined relative velocity vector and to diffuse the latter to discharge a flow of the fluid having a third determined relative velocity vector less than the second relative velocity vector and having both tangential and radially outward components of substantially zero. The annular wall may define a second right circular cylindrical portion radially outwardly bounding the flow path downstream of the concave annular surface portion, and a first annular array of plural radially extending diffuser vanes may extend between the convex annular surface portion and annular wall, the diffuser vanes each having a leading edge with respect to the fluid flow. The leading edges of the first array of diffuser vanes may intersect the annular wall downstream of the concave annular surface portion of the annular wall. The leading edges of the first annular array of diffuser vanes may be swept downstream radially outwardly with respect to a radial line from the axis. The other annular wall may define a third right circular cylindrical portion spaced radially inwardly of the second right circular cylindrical portion and may radially inwardly bound the flow path downstream of the convex annular surface portion and a second annular array of radially extending diffuser vanes may extend between the second and third right circular cylindrical wall portions. The first annular array of diffuser vanes may each include a trailing edge spaced axially downstream of the leading edge, the leading edges and trailing edges cooperating to define a chord dimension for the first annular array of diffuser vanes, the second annular array of diffuser vanes being spaced axially downstream of the first array by substantially one-half of the chord dimension. According to a preferred aspect of the invention, therefore, there may be provided a first axially ending radially outer annular wall having first to fifth portions arranged sequentially downstream with respect to a fluid flow through the compressor, the first wall portion defining a compressor inlet and inwardly being of right circular cylindrical shape in transverse section, the second wall portion being radially inwardly convex to transition between the first wall portion and the third wall portion, the third wall portion being inwardly of right circular conical shape and diverging downstream in transverse section, the fourth wall portion being radially inwardly concave to transition between the third wall portion and the fifth wall portion and defining an annular fluid flow path in cooperation with a second radially inner annular wall, the fifth wall portion being inwardly of right circular cylindrical shape in transverse section; the second radially inner annular wall being disposed downstream of the inlet in substantial radial juxtaposition with the fourth and fifth portions of the radially outer wall to define a rotor chamber in juxtaposition with the first to third wall portions, the second wall having portions thereof designated sixth and seventh in sequential downstream axial arrangement, the sixth wall portion being radially outwardly convex and arcuate in axial section to define a radius R and further cooperating with the fourth wall portion to define the annular flow path, the annular flow path extending axially and radially outwardly and being of substantially constant transverse fluid flow area, the seventh wall portion being outwardly of right circular cylindrical shape in transverse section and cooperating with the fifth wall portion to define a downstream extension of the annular flow path also having a substantially constant transverse fluid flow area; a rotor member rotationally disposed with the rotor chamber and including a substantially conical hub portion having an inlet end diameter and an outlet end of relatively larger diameter axially adjacent to and substantially matching in diameter with the sixth wall portion, a plurality of circumferentially spaced apart axially, and circumferentially extending compressor blades extending radially outwardly from the hub toward but short of the first annular wall to terminate in radially outer blade tip edges in shape matching movable relation with the first to third wall portions, said plurality of blades each defining a radially extending leading edge defining a virtual intersection with the first wall portion and a radially extending trailing edge defining a virtual intersection with the third wall portion; a first plurality of circumferentially spaced apart axially extending diffuser vanes extending radially between and intersecting with the fifth and sixth wall portions each of the first plurality of diffuser vanes defining a leading edge disposed at its radially outer end downstream of the fourth wall portion the first plurality of diffuser vanes also each having a trailing edge defining a chord dimension therefore; a second plurality of circumferentially spaced apart axially extending diffuser vanes spaced downstream of the first plurility by substantially one-half of the chord dimension and extending radially between the fifth and seventh wall portions. According to another aspect of the present invention, a transonic mixed flow compressor is presented comprising a housing defining an inlet portion an outlet portion, and an axially extending flow path extending therebetween for flow of the elastic fluid; a rotor journaled in the flow path for rotation about the axis and having a respective inlet end and oulet end, the housing and rotor defining cooperating means for defining an annular stream tube extending axially from the inlet towards the outlet in the flow path and diverging downstream radially outwardly to define at a radially outer boundary thereof upstream of the rotor outlet end substantially a right circular conical section. According to another aspect of the invention there is provided a method of pressurising an elastic fluid comprising the steps of : forming a tubular stream of the fluid having a radially inner diameter, a radially outer diameter, and a first relative velocity vector sum of meridional and tangential velocities of at least Mach 1.2 at the radially outer diameter; diffusing the fluid to a second supersonic relative velocity less than the first relative velocity while limiting deviation of radially outer local relative velocity vectors to no more than 10 o with respect to the first relative velocity vector; passing the fluid through a normal shock to a third relative velocity of less than Mach 1; and further diffusing the fluid stream while increasing downstream both the radially inner and radially outer diameters thereof to impart a significant radially outward component of meridional velocity thereto. A method of compressing elastic fluid is also encompassed by the present invention comprising the steps of forming a rotational annulus of axially flowing fluid fluid having an inner diameter, an outer diameter, and a first relative velocity vector sum of meridional and tangential velocities of less than Mach 1; diffusing the flowing fluid to a second relative velocity less than the first relative velocity while increasing progressively downstream the outer diameter and increasing the radially outward component of the meridional velocity; holding the increase of the outer diameter to a constant axial rate while further diffusing the flowing fluid to a third relative velocity less than that of the second relative velocity while decreasing the radially outward component of the meridional velocity to a value less than that of the second relative velocity. The invention may therefore extend to a method of pressurising elastic fluid to increase its pressure level from a first pressure level to a higher second pressure level, the method comprising the steps of: forming a first tubular stream of the fluid at the first pressure level, having a longitudinal axis, a first radially inner diameter, a first radially outer diameter, and a first relative velocity vector sum of meridional velocity and tangential velocity of at least Mach 1.2 with respect to a selected reference at the radially outer diameter; subdividing the flow stream into a plurality of axially extending and circumferentially spaced apart substreams; diffusing each of the substreams to a second supersonic relative velocity less than the first relative velocity while limiting deviation of radially outer local relative velocity vectors to no more than 10 o with respect to the first relative velocity vector; passing each of the substreams through a respective normal shock to a third relative velocity of less than Mach 1; increasing progressively downstream for the normal shock both the radially inner and radially outer diameters of the flow stream as an aggregate of the substreams respectively above the first diameters while further diffusing the relative velocity of each of the substreams to a fourth subsonic relative velocity less than the third relative velocity and increasing the radially outward component of local meridional velocity; limiting the downstream increase of the outer aggregate diameter to a constant axial rate while decreasing the radially outward component of local meridional velocity and effecting further diffusion of each substream to a fifth subsonic relative velocity less than the fourth relative velocity; and reuniting the substreams into a respective second tubular stream of the fluid having a sixth subsonic absolute velocity vector sum of meridional and tangential velocities. Such a method may include the further steps of subdividing the second tubular stream into a second plurality of axially extending and circumferentially spaced apart substreams each having substantially the sixth absolute velocity vector sum of tangential and meridional velocities, the latter of which includes a radially outward component of a seventh value; diffusing each of the second substreams to an eighth absolute velocity vector sum of tangential and meridional velocities having a tangential component less than that of the sixth vector sum and a radial component of substantially zero; and reuniting the second plurality of substreams into a respective third tubular stream of the fluid having the eighth subsonic absolute velocity vector sum or meridional and tengential velocities. The method may also include the further steps of : subdividing the third tubular stream into a third plurality of axially extending and circumferentially spaced apart substreams each having substantially the eight absolute velocity vector sum or meridional and tangential velocities; diffusing each of the third substreams to a ninth absolute velocity vector sum of meridional and tangential velocities having a tangential component of substantially zero while maintaining a radial velocity component of substantially zero; and reuniting the third plurality of substreams into a respective fourth tubular stream of the fluid having the ninth subsonic absolute velocity vector sum of meridional and tangential velocities having both radial and tangential components of substantially zero (substantially pure axial flow). The invention may also extend to a method of diffusing an annulus of flowing elastic fluid having a first absolute velocity and respective first tangential and radially outward components of velocity, the method comprising the steps of : subdividing the annulus into a plurality of radially outwardly and axially extending circumferentially spaced apart substreams each having substantially the first tangential and radially outward components of velocity; diffusing each of the plurality of substreams to a second absolute velocity less than the first absolute velocity and having a respective axial velocity, a respective tangential velocity significantly less than the first tangential velocity, and a respective radially outward velocity component in the range of substantially zero to a negative value; and reuniting the plurality of substreams into a second annuls of flowing fluid having the second absolute velocity, the second axial velocity component, the second tangential component of velocity, and the second radially outward component of velocity. This latter method may include the further steps of subdividing the second annulus into a second plurality of axially extending and circumferentially spaced apart substreams each having substantially the second absolute velocity, the second axial velocity component, the second tangential velocity component, and the second radially outward velocity component; diffusing each of the substreams to a third absolute velocity less than the second absolute velocity, and having a respective axial velocity component, a respective tangential velocity component of substantially zero, and a respective radial velocity component of substantially zero; and reuniting the third plurlity of substreams into a third annulus of flowing fluid having the third absolute velocity including the third axial velocity component, the third tangential velocity component of substantially zero, and teh third radial velocity component of substantially zero (substantially pure axial flow). This invention also encompasses a jet propulsion engine incorporating a compressor method and/or apparatus as described above. It will be appreciated that the present invention may substantially satisfy each of the obectives outlined above and by doing so may provide the highly desirable advantages resulting therefrom. Additionally, the applicants have found that the compressor provided, because its diffuser structure presents a diffuser flow path defined between coannular right circular cylindrical wall sections, affords a structure of greater strength for a particular weight than either conventional centrifugal or mixed-flow diffuser structures. The invention may be carried into practice in various ways and some embodiments will now be described by way of example with reference to the accompanying drawings, in which:- Figure 1 is a schematic longitudinal partially cross sectional view through a jet propulsion turbo fan engine according to the invention; Figure 2 is an enlarged fragmentary axial view taken at line 2-2 of Figure 1; Figure 3 is an enlarged fragmentary view of the encircled portion designated 3 in Figure 1, partially in cross section and having portions of the structure omitted for clarity of illustration; Figure 4 is similar to Figure 3, but to a larger scale, with details of construction omitted to present more clearly the geometric aspects of the invention; and; Figure 5 is a fragmentary view taken parallel with line 5-5 at the radially outer tip of the compressor rotor of Figure 3 with the perspective being radially inwards. Figure 1 schematically depicts a turbofan jet propulsion engine 10 which includes an elongate housing 12. The housing 12 defines an inlet opening 14 through which ambient air is drawn, and an outlet opening 16 through which a jet of heated air and combustion products is expelled to the atmosphere. Journaled within the housing 12 is a shaft 18 which is driven by a turbine section 20 of the engine 10. At its forward end the shaft 18 carries a mixed-flow compressor rotor 22 which draws ambient air through the inlet opening 14 and pressurises the inducted air for use by the remainder of engine 10. Immediately downstream of the rotor 22, the housing 12 defines an annular flow path 24 in which a diffuser structure generally referenced with numeral 26 is located, and which in combination with the rotor 22 comprises the first compressor stage of the engine 10. Downstream of the diffuser 26, the flowpath 24 is bifurcated into an outer annular flowpath passage 28, and an inner annular core engine flowpath 30. The flowpath 28 communicates directly downstream with a tailpipe portion 32 of the engine 10 and the tailpipe portion communicates with the outlet opening 16. Accordingly, it will be appreciated that the compressor rotor 22 serves also in the capacity of a fan with respect to the turbofan nature of the engine 10. The core engine flowpath 30 proceeds downstream through a two-stage axial flow compressor section 34, having two axially spaced apart blade wheels drivingly carried by the shaft 18. The flowpath 30 subsequently extends through an annular combustor 36, and through the turbine section 20. The turbine section 20 also communicates with the tailpipe portion 32 and with the outlet opening 16. Turning now to Figure 2, this illustrates that the compressor rotor 22 includes a hub portion 38, which reference to Figures 1 and 3 will show defines an outer surface 40 of elongate conical shape. Disposed upon the hub 38 and extending radially outwardly thereon is a plurality of axially and circumferentially extending blades 42. According to the preferred embodiment of the invention as shown, there are seventeen blades 42 which are equiangularly circumferentially spaced apart. Each blade 42 defines a radially extending leading edge 44, a radially extending trailing edge 46, and a radially outer axially and radially extending tip edge 48. As shown most clearly in Figure 3, it will be seen that the blades 42 extend radially outwardly towards a wall portion 50 of the housing 12 and terminate in the radially outer tip edges 48 which are spaced slightly radially inwardly of the radially inner surface 52 of the wall portion 50 and have a shape conforming to that of the inner surface 52. The wall portion 50 extends continuously axially and circumerentially from the inlet opening 16 downstream past the compressor rotor 22, the flow path 24, and the diffuser section 26. Beginning at inlet opening 16 and continuing downstream (to the right in Figure 3) at a selected distance therefrom, the wall portion 50 defines a radially inner surface subsection 52 a in the shape of a right circular cylindrical surface. The right circular cylindrical surface portion 52 a of wall 50 extends downstream beyond the leading edges 44 of the blades 42. The wall portion 50 adjacent the trailing edges 46 and extending certain distances both upstream and downstream of the virtual intersection of the trailing edges with the wall surface 52 defines a radially inner surface subsection 52 b in the shape of a truncated right circular conical surface. Between the right circular cylindrical subsection 52 a of the wall 50 and the right circular conical subsection 52 b , the wall portion 50 defines an axially curvilinear radially inner transition surface subsection 52 c which is radially inwardly convex. In other words, between the leading edges 44 and the trailing edges 46 of the blades 42, the wall 50 defines a subsection 52 c which is an axially curvilinear transition surface of revolution, and which avoids a defined cusp between the cylindrical and conical subsection 52 a and 52 b . It is significant that the curvilinear transition subsection 52 c does not extend to the trailing edges 46, and in fact joins the subsection 52 b some distance upstream of these trailing edges. As shown in Figures 3 and 4, upstream of the virtual intersection of the leading edges 44 with the wall 50 (figure 4, point B) and extending downstream thereto, the right circular cylindrical surface 52 a has an axial dimension of from about 10% to about 20% of the meridional dimension of the blades 42 at their tip edges 48 (Figure 4, A-B dimension). Similarly, extending downstream from the virtual intersection of the leading edges 44 with the wall 50 (Figure 4, point B), the right circular cylindrical surface 52 a has an axial dimension of from about 10% to about 30% of the meridional dimension of the blades 42 at their tip edges 48 (Figure 4, B-C dimension). Adjacent the virtual intersection of the trailing edges 46 with the wall 50 (Figure 4, point E), the right conical surface portion 52 b extends both upstream and downstream. The downstream meridional extension of the surface 52 b is from about 5% to about 15% of the meridional length of the blades 42 at their tip edges 48. The upstream meridional extension of the surface 52 b from the point E is from about 10% to about 30% of the meridional dimension of the blades 42 at their edges 48 (Figure 4, D-E dimension). Consequently, the transition surface subsection 52 c defines from about 40% to about 80% of the meridional dimension of the blades 42 at their edges 48. It will be seen that in axial cross section, the flowpath coextensive with the rotor 22 is radially outwardly bounded by the surfaces 52 a and 52 b defining two relatively inclined axially extending straight line segments. The straight line segments of the surfaces 52 a and 52 b are jointed by a continuous, smooth, nonlinear curved surface section 52 c which is tangential to both of the straight line surface sections. Preferably, the surfaces 52 a and 52 b define an acute angle (referenced with numeral 54) of about 22 o with respect to one another. However, the angle 54 may be from about 5 o to about 45 o . It will also be seen that the leading edges 44 of the blades 42 are swept downstream radially outwardly with respect to a radially extending line 56 perpendicular to the rotational axis of the rotor 22. Preferably, the leading edges 44 define an acute angle 58 of about 7 o . However, the angle 58 may be from about 0 o to about 15 o . Similarly, the trailing edges 46 are swept upstream radially outwardly with respect to a radially extending line 60 perpendicular to the rotational axis of the rotor 22. Preferably, the trailing edges 46 define an acute angle 62 of about 23 o . The angle 62 may, however, be from 0 o to about 35 o . Further, with respect to the hub 38 and blades 42 thereon, it will be seen from Figure 4 that a radius dimension RBi is defined at the intersection of the leading edge 44 with the outer surface of the hub 38. Similarly, at the intersection of the trailing edge 46 with the surface 40, a radius dimension REi is defined. In the illustrated embodiment the ratio of REi to RBi is about 2.75. However, this ratio may permissibly vary between about 1.5 and 3.5. The present applicants have discovered that, in combination with the other salient features herein described, a relatively small ratio of outer radius change of the rotor 22 from leading edge to trailing edge of blades 42 should be employed. In other words, at the virtual intersection of the leading edge 44 with the surface 52, a radius dimension RBo is defined. At the virtual intersection of the trailing edge 46 with the surface 52, a radius dimension REo is similarly defined. The ratio of REo to RBo is preferably 1.17, however, this ratio may very from about 1.05 to about 1.76. As will be seen, this relatively low ratio of radius increase from inlet to outlet of the rotor 22 contributes to the relatively small overall diameter which is possible in a compressor according to the invention, in comparison to its inlet diameter. A further geometric aspect of the rotor 22 which is considered of importance by the applicants is a dimensionless ratio termed the Aspect Ratio (AR), defined below:- The average meridional blade length of blades 42 is depicted on Figure 4 as line 64, which is generated by those points on the blade lying radially midway between the surface 40 and the tip edge 48. Preferably, the ratio AR is 1.12. This ratio may, however, vary between about 0.75 and 1.30. Downstream of the trailing edges 46, the housing 12 defines annular fluid flow path 24 between the radially outer wall 50 and an annular radially inner wall 65 which is spaced radially inwardly of the wall 50 and has a radially outwardly disposed surface portion 66 a . The surface 66 is a curvilinear surface of revolution having a radius 68 originating from a centre point 70. The radius 68 is related to the height of the blades 42 at the trailing edge 46 in that the radial distance along the trailing edge 46 from its intersection with the surface 40 of hub 38 to its virtual intersection with surface 52 is considered the blade height at the trailing edges of the blades 46. The ratio of the radius 68 to the blade height at the trailing edge 46 is preferably 2.0, however, this ratio may vary from about 1.0 to about 4.0. Immediately downstream of the trailing edges 46, the flow path 24 is circumferentially continuous between the walls 50 and 65. The radially outer wall 50 defines a surface portion 52 d which is a curvilinear surface of revolution tangent at its upstream end with the right circular conical surface portion 52 c . The radius of the wall surface 52 d is matched so that of the surface portion 66 a so that a flow path portion 24 a is defined which is of substantially constant area despite the radius change in the flow path with respect to the rotational axis of the rotor 22. At its downstream end, the surface portion 52 d is also tangent with a right circular cylindrical surface portion 52 e of the wall 50. Similarly the wall 65 defines a radially outwardly disposed right circular cylindrical surface portion 66 b which at its upstream end is tangent with the surface portion 66 a of the wall 65. It will be seen from Figures 3 and 4, that an annular array of radially extending and circumferentially spaced apart diffuser vanes 72 extend between the walls 50 and 65 in the region of the surface portions 52 e and 66 a thereof, respectively. The vanes 72 each define a leading edge 74, and a trailing edge 76. While it will be noted that at their radially inner ends, the vanes 72 are relatively close to the trailing edge 46 of the compressor blades 42 and intersect the curvilinear surface portion 66 a , the radially outer ends of the vanes 72 intersect the cylindrical surface portion 52 e . Thus, the diffuser vanes are swept downstream radially outwardly to intersect the radially outer wall 50 at the cylindrical portion 52 e . It is significant that the leading edge 74 of each vane 72 intersects the wall 50 down stream of the curvilinear surface portion 52 d . That is, the vanes 72 at their radially outer ends intersect the right circular cylindrical portion of the wall 50 at the surface 52 e . It will be noted that the vanes 72 are swept downstream radially outwardly with respect to a radial perpendicular from the rotational axis of the rotor 22. The physical sweep angle is in the range of from zero degrees to twenty-five degrees. However, the vanes 72 are swept to an even greater aerodynamic degree with respect to the air flow from the rotor 22. This is due to the combination of axial and radially outward velocity components of this air flow, as will be explained more fully below. Downstream of the trailing edges 76 of the diffuser vanes 72, the flow path 24 is once again circumferentially continuous to define a vane-wake dissipation area 24 b . Since the surfaces 52 e and 66 b are right circular cylindrical surface sections, it will be appreciated that the flow path portion 24 b is of constant cross sectional flow area. As a result, substantially no flow diffusion occurs within the portion 24 b . Downstream of the diffuser vanes 72, and spaced axially therefrom by approximately one-half a cord dimension of the latter, the diffuser 26 also includes a second annular array of radially extendin+g and circumferentially spaced apart diffuser vanes 78. The diffuser vanes 78 include leading edges 80 and trailing edges 82 which are swept with respect to a radial perpendicular from the rotational axis of rotor 22. The geometric sweep of the vanes 72 closely approximates to their aerodynamic sweep angle because the air flow traversing these diffuser vanes has little or no radial velocity component. Having described the structure of the compressor stage, its operation according to aerodynamic theory will now be discussed. During operation of the engine 10, the turbine section 20 drives the shaft 18 at a high speed of rotation, and the shaft 18 drives the compressor rotor 22. It will be seen from Figures 3 and 4, that the wall 50 defining the surface portion 52 a of the housing 12 cooperates with the axially extending conical nose portion 38 a of the hub 38 to define an axially extending annular passage 24'. In response to rotation of the rotor 22, ambient air is drawn through the passage 24', as shown by arrow 84. Consequently, it will be seen that the passage 24' defines an axially extending annular flow stream of axially flowing air 84. Upon encountering the leading edges 44 of the blades 42, the annular flow stream of air 84 is subdivided into substreams which are circumferentially spaced apart by the blades 42. As Figure 5 shows, adjacent the tip edge 48, the leading edges 44 have a tangential velocity vector (Vt) referenced with numeral 86. Consequently, air flow 84 approaching the blades 42 has a negative relative tangential velocity vector of -Vt, which is referenced by numeral 88. The air flow 84 adjacent the tip edges 48 also has a meridional component of relative velocity represented by vector (Vm) and referenced with numeral 90. Meridional velocity as used herein is the vector sum of axial and radial airflow velocity components. Consequently, the air flow 84 has a relative velocity vector sum of Vt and Vm which is represented by vector Vr, and referenced with number 92. The meridional relative velocity vector Vm includes also any radial relative velocity component (VR) between the blades 42 and the air flow 84. However, adjacent the intersection of the tip edges 48 and the leading edges 44, the outward radial velocity of the air flow 84 must be zero, or substantially zero because the wall surface 52 a is of right circular cylindrical shape. The relative velocity vector 92 (Vr) has a magnitude of at least Mach 1.2, and is preferably in the range of Mach 1.3 to 1.5, or higher. As a consequence of the high relative velocity of the vector Vr, the leading edge 44 of each blade 42 is believed to originate a Mach wave 94. Also, the surface of the blades 42 extending downstream of the leading edge 44 and facing in a tangential direction opposite to the rotational direction of the rotor 22 is shaped according to known aerodynamic principles to originate multiple additional Mach waves, which, as depicted, are two in number and are referenced with numerals 96, 98. It will be understood that the number of additional Mach waves may be other than two. However, it is believed to be important that at the selected operating condition for the compressor stage, one of the Mach waves (wave 98 as depicted) encounters the next circumferentially adjacent blade 42 opposite to the direction of rotation of the rotor 22 at or downstream of its leading edge 44. As a result, the wave 98 becomes a captured Mach wave. At or adjacent the leading edge 44 which has captured the Mach wave 98, an oblique shock wave 100 is believed to originate and to extend toward the next circumferentially adjacent blade 42 in the direction of rotation of the rotor 22. Subsequently, a normal shock 102 is believed to be formed downstream of oblique shock 100. Each of the Mach waves 94, 96, 98, the oblique shock 100 and the normal shock 102 is believed to effect a diffusion or slowing of the flow of air 84 relative to blades 42, and a concomitant increase of total pressure of the airflow. As an aid to the reader, notations have been placed upon Figure 5 which are generally indicative of the relative velocity of the airflow field at the location of the particular notation. To summarise therefore, the airflow 84 approaches the rotor 22 as an axially flowing annular stream tube having a relative velocity of at least Mach 1.2. (Figure 5, M»1) represented by the vector 92 (Vr). Upon encountering the blades 42, the airflow is weakly diffused through successive Mach waves 94-98, the last of which is a captured Mach wave preceding a stronger oblique shock 100. Upstream of the oblique shock 100, the air flow has a relative velocity vector 104 which is supersonic (Figure 5, M»1), but less in magnitude than velocity vector 92. Downstream of the oblique shock 100, the airflow has a relative velocity which is greater than Mach 1, (Figure 5, M>1), but less than the velocity upstream of the shock 100. Immediately downstream of the normal shock 102, the airflow has a relative velocity less than Mach 1 (Figure 5, M<1), and a direction substantially perpendicular to the shock 102, as is represented by vector 106. It is significant that the vector 106 has a direction which deviates by no more than 10 o with respect to the vector 92. Large turning angles in the presence of diffusion of supersonic flow are extremely difficult to obtain without very high aerodynamic losses. The Applicants have discovered that a deviation of 10 o or less will allow diffusion of the supersonic airflow preceding a normal shock with high efficiency. The Applicants believe that a limited supersonic flow turning in range of about 10 o or less enhances the probability of achieving a shock structure having only a single normal shock, and monotonically decelerating flow to a velocity of less than Mach 1. Returning to Figure 3, it will be seen that the hub 38 increases in radial dimension somewhat uniformly in the axial direction. However, the tangential velocity of points on the blades 42 increases with radius and is the product of rotor angular velocity and radius. Consequently, the relative velocity adjacent the hub 38 is much less than the level of Mach 1.2 or higher, which is experienced at the tip edges 48 adjacent the leading edges 44. As a result, flow deviation adjacent the intersections of the leading edges 44 with the hub 38 may permissibly exceed 10 o . Nevertheless, the applicants believe that adjacent the intersection of the tip edges 48 and the leading edge 44 where the relative velocity is Mach 1.2 or higher, the flow turning cannot be allowed to exceed 10 o without incurring undesirable aerodynamic losses. It follows that the cylindrical wall subsection 50 a is of importance in the present invention because such a subsection limits the radial component of local meridional velocity to substantially zero. Accordingly, the tangential velocity component Vt and rotational speed of the rotor 22 is easily determined so that Vr at the intersection of the leading edges 44 with the tip edges 48 is Mach 1.2 or higher. While the cylindrical wall subsection 50 a is considered an important and desirable feature of a transonic mixed flow compressor according to the invention, deviation from cylindrical shape is permissible so long as the resulting radial component of Vm (vector 90) is taken into consideration. It will further be recalled from Figure 3, that the wall subsection 52 c curves outwardly to increase its radial dimension downstream. This outward flare of the wall subsection 52 c blending with the upstream end of the right circular conical wall subsection 52 b is accompanied by an increase in radially outward component (VR) of the relative meridional velocity and a further diffusion to a relative velocity considerably below Mach 1. This increase in the radially outward component of the meridional velocity with increasing radius is expected and is the advantageous effect of centrifugal compressors which conventional mixed flow compressors have attempted to use. However, the radially outward velocity component continues to increase with the radial dimension as air flows along the rotor in conventional centrifugal and mixed-flow compressors to such an extent that conventional downstream flow turning losses, diffuser structure difficulties, and excessive overall outer diameter limitations have always persisted in the past. Surprisingly, the applicants have discovered that in a compressor according to the invention, even though the annular flow stream is diverging downstream and increasing in its radial dimension, if the radially outer flow stream boundary is limited to increasing its radial dimension as a substantially constant linear function of the axial dimension, then the growth of the radially outward velocity component will cease or reverse itself within the axial confines of the compressor rotor. In other words, the meridional velocity vector (summation of axial and radial velocity components) which begins to swing radially outwardly with increasing radial dimension of the annular flow stream downstream of the wall subsection 52 a (arrow 108, viewing Figure 3) actually stops its outward swing and begins to swing back towards the true axial direction (arrow 110, viewing Figure 3) upstream of the trailing edges 46 of the rotor 22 despite the increasing radial dimension of the annular flow stream. In view of the above, the importance of the right circular conical wall subsection 52 b is easily appreciated this wall subsection 52 b forms a critically important radially outer boundary of the annular flow stream traversing the compressor rotor 22. Preferably, the right circular conical subsection 52 b extends upstream of the trailing edges 46 a distance of about 10% to 30% of the meridional dimension of the tip edges 48. However, a relatively shorter segment of conical wall section may be employed within the scope of the present invention provided that the conic section extends upstream of the trailing edges of the rotor blades sufficiently to limit the growth of, to hold constant, or to effect a decrease in the radially outward component of the meridional airflow velocity. The decrease in the radially outward airflow velocity on the rotor 22 described above makes it possible to use a diffuser structure 26 which is believed to be more akin to axial-flow technology and appears to be quite different to any conventional mixed-flow compressor stage known to the Applicants. Considering the diffuser 26, the wall 50 radially outwardly of the flowpath subsection 24 a is axially curvilinear to effect a limited turning of the air flow from the rotor 22 which has tangential, axial and radially outward absolute velocity components. Turning the airflow in response to the curvature of the wall 50 effects a reduction in the radially outward velocity component. Additionally, the vaneless space 24 a is believed to effect an accommodation of local flow aberrations so that the flow is more fully homogenised before encountering the diffuser vanes 72. The diffuser vanes 72 each have a leading edges 74 which is apparently swept in a physical sense, i.e. swept relative to a perpendicular from the rotational axis of rotor 22. The vanes 72 are swept with their radially outer leading edge farther downstream than their radially inner leading edge. What is not immediately apparent from viewing Figure 3 is that the vanes 72 have an aerodynamic sweep angle exceeding their physical sweep angle. This is because the air flow in the diffuser space 24 a has a substantial radially outward velocity component so that the airflow may be represented by a radially outwardly angulated vector like the arrows 108 and 110, viewing Figure 3. It will be appreciated that because the air flow is angled radially outwardly, the aerodynamic sweep angle of the vanes 72 exceeds their physical sweep angle according to the angulation of the air flow in the space 24 a relative to the axis of the rotor 22. An additional aspect of the diffuser vane 72 which may not be immediately apparent from the structure alone is that while the vanes extend substantially radially between the walls 50 and 65, they are leaned tangentially in an aerodynamic sense. In order to understand this lean it must be recalled that the airflow in the diffuser space 24 a has a significant tangential component of velocity. Furthermore, this tangential velocity component increases radially outwardly from the wall 65 towards the wall 50. Consequently, even though the vanes 72 extend substantially radially between the walls 50, 65 they are aerodynamically leaned radially outwardly in a direction opposite to the direction of rotation of the rotor 22, and opposite to the direction of tangential velocity of the airflow in the diffuser space 24 a . In other words, the radially outer end of each vane 72 is aerodynamically displaced tangentially relative to the radially inner end of the vane in a direction opposite to the direction of rotation of the rotor 22. Thus, an air flow element leaving rotor 22 near the tip edge 48 will, because of its higher tangential and axial velocity, encounter the leading edges 74 of a diffuser vane 72 in a shorter time than an airflow element leaving the rotor near the hub 38. Downstream of the array of diffuser vanes 72, the diffuser section 26 includes another vaneless space 24b which extends axially between the walls 50 and 65. In a way similar to to the vaneless space 24 a , the space 24 b is believed to allow "adjustment" or accommodation of local flow aberrations so that the airflow is more fully homogenised before encountering the diffuser vanes 78. The diffuser vanes 78 extend radially between the walls 50, 65 and diffuse the airflow from the diffuser space 24 b to a lower absolute velocity and an increased static pressure while reducing the tangential velocity component to substantially zero. Furthermore, the radial component of the airflow velocity which was reduced to substantially zero by the diffuser vanes 72 is maintained at the substantially zero level by the vanes 78. Consequently, the diffuser vanes 78 deliver to the flow path subsection 24 c an airflow having a substantially pure axial flow. The shape of the blades 42 on the rotor 22 is determined in accordance with conventional airflow streamline productive techniques. In other words, with a view to the inventive compressor methods herein described, the airflow streamline directions at points in the annular flow stream are predicted based upon known aerodynamic principles. Appropriate blade surface segments are then established based upon accepted blade pressure loading, airflow velocity, and diffusion parameters. The resulting blade surface segments are stacked radially outwardly from the hub 38 and blended to determine the resulting shape of the blades 42. Accordingly, the precise shape, thickness, curvature and other characteristics of the blades 42 will vary in dependence upon the design objectives of each particular compressor stage. An actual reduction to practice the present invention has been made substantially according to the preferred embodiment herein described. This included a rotor having 17 blades, an overall axial dimension of 4 inches (101.6mm) excluding the rotor nose portion, a hub/tip diameter ratio of 0.35 at the blade leading edges, and an operating speed of 33,216 R.P.M. With a compressor stage static pressure ratio of 3.0 and a corrected through flow of 20 pounds (9.1kg) of air per second, the rotor adiabatic efficiency was 89.5%, and the stage efficiency was 85.3%, with an airflow velocity of Mach 0.54 downstream of the stator vanes 78. The diffuser vanes 72 and 78 each numbered 44 vanes equiangularly disposed downstream of the rotor 22. The overall compressor stage length was 10.6 inches (269.2mm) with a rotor outer diameter of 12.8 inches (325.1mm). As can be seen from Figure 3, the overall outer diameter of the compressor stage at about 105% of the rotor outer diameter was only slightly larger than the outer diameter of the compressor rotor. Accordingly, it will be seen in view of the above that the present invention can provide a transonic mixed-flow compressor which achieves a static pressure ratio favourably comparable to a centrifugal compressor while achieving an efficiency approaching the best contemporary axial-flow compressor technology. Testing has verified the tolerance to inlet airflow distortion of a compressor according to the invention. Resistance to damage from ingestion of foreign objects is believed to be substantially the same for a compressor according to the invention as for a centrifugal compressor. Furthermore, it is believed that a compressor according to the present invention will have a cost comparable to contemporary centrifugal compressors. Finally, the outer diameter or envelope of a compressor stage according to the present invention is very favourably comparable to conventional axial flow compressors.

Claims: Compressor apparatus comprising a housing (12) and a rotor (22), the housing (12) comprising an axially and circumferentially extending annular wall (50) having an inlet (14) and an outlet (16), the rotor (22) being journaled within the housing (12) and having a hub (38) with a plurality of blades (42), the hub (38) and annular wall (50) together defining an annular axially extending passage (24), characterised in that the annular wall (50) includes: at the inlet end, a portion (52a) which is at least up to the leading edge of the stator vane of right circular cylindrical shape in section; adjacent the rotor (22) an outlet portion (52b) which is of right circular conical shape in section and which diverges downstream relative to the inlet, the portion (52b) diverging downstream with a divergence angle (54) of 45 o or less and extending upstream of the virtual intersection (E) of the trailing edges (46) of the blades (42); and a smooth transitional intermediate portion (52c) between the inlet portion and conical outlet portions; and further characterised in that the hub (38) is cone shaped and extends between said inlet and said outlet portion (52b). Apparatus as claimed in Claim 1 characterised in that the blades (42) extend axially, circumferentially and radially from the hub (38) to a position just short of the annular wall (50), and conform closely to the shape of the annular wall (50) at axial positions throughout the inlet portion (52 a ), the intermediate portion (52 c ) and the outlet portion (52 b ). Apparatus as claimed in Claim 1 or Claim 2, characterised in that the blades (42), the hub (38) and the annular wall (50) define a flow path for receiving a flow stream of compressible fluid having a first vector sum of meridional velocity and tangential relative velocity of at least Mach 1.2 with respect to a selected reference adjacent the inlet portion (52 a ), and for diffusing the flow stream at a second supersonic relative velocity less than the first relative velocity while limiting deviation of radially outer local relative velocity vectors to no more than 10 o with respect to the first relative velocity vector. Apparatus as claimed in any preceding claim, characterised in that each blade (42) defines a radially extending leading edge (44), a radially extending trailing edge (46), and a radially, axially, and circumferentially extending radially outer tip edge (48) conforming in shape to the annular wall (50) at the inlet portion, the intermediate portion, and the outlet portion, the tip edges considered in axial and radial aspect defining a tip edge meridional dimension for the blades, the leading edges defining a virtual intersection (B) with the annular wall (50) at the inlet portion, and the trailing edges defining a virtual intersection (E) with the annular wall (50) at the outlet portion. Apparatus as claimed in Claim 4, characterised in that the inlet portion (52 a ) of right circular cylindrical section extends upstream of the virtual intersection (B) of the leading edges (44) for a distance of from about 10% to about 20% of the tip edge meridional dimension, and the inlet portion (52 a ) of right circular cylindrical section extends downstream of the virtual intersection (B) of the leading edges (44) for a distance of from 10% to 30% of the tip edge meridional dimension. Apparatus as claimed in any preceding claim characterised in that the right circular section inlet portion (52 a ) and the right circular conical section outlet portion (52 b ) of the annular wall (50) are in axial section angularly disposed relative to one another so as to define an acute angle between them in the range from about 5 o to about 45 o Apparatus as claimed in any of Claims 4 to 6 characterised in that the conical section outlet portion (52 b ) of the annular wall (50) extends upstream of the virtual intersection (E) of the trailing edges (46) for a distance of from about 10% to about 30% of the tip edge meridional dimension, and the conical section outlet portion (52 b ) of the annular wall (50) extends downstream of the virtual intersection (E) of the trailing edges (46) for a distance of from about 5% to about 15% of the tip edge meridional dimension. Apparatus as claimed in any of Claims 4 to 7 characterised in that the leading edges (44) define a diameter RBi at their intersection with the hub (38) and a diameter RBo at their intersection with the tip edges 48 and the trailing edges 46 define a diameter REi at their intersection with the hub 38 and a diameter REo at their intersection with the tip edges (48); the ratio of REi to RBi being in the range from about 1.5 to about 3.5, and the ratio of REo to RBo is in the range of about 1.05 to about 1.76. Apparatus as claimed in Claim 8 characterised in that each of the blades define a quantity termed, average meridional blade length (AMBL), which is the length in axial and radial aspect of a line along a blade (42) from the leading edge (44) to the trailing edge (46) and defined by points on the blade (42) radially midway between the hub (38) and the tip edge (48), the ratio (AR) of (RBo-RBi) + (REo-REi) to AMBL lying in the range from about 0.75 to about 1.30. Apparatus as claimed in any of Claims 4 to 9 characterised in that the blades (42) have a height dimension at their trailing edge defined as REo-REi, and the compressor apparatus further includes a second axially and circumferentially extending annular wall (65) disposed radially inwardly of the first annular wall (50) and immediately downstream of the rotor (22), the two annular walls (50,65) being radially spaced apart to define an axially extending annular flow path downstream of the rotor (22), the inner wall (65) defining a first radially outwardly convex annular surface portion (66 a ) bounding the flow path and being arcuate of radius R in axial section, the ratio of the trailing edge blade height to the radius R lying in the range from about 1.0 to about 4.0. Apparatus as claimed in Claim 10 characterised in that the first annular wall (50) defines a radially inwardly concave annular surface portion (52 d ) bounding the flow path downstream of the right circular conical outlet portion (52b), the concave annular surface portion (52 d ) of the first annular wall (50) cooperating with the convex annular surface portion (66 a ) of the second annular wall (65) to define the flow path downstream of the rotor (22), the flow path extending radially outwardly and axially with a substantially constant transverse sectional fluid flow area. Apparatus as claimed in any preceding Claim characterised by annular diffuser means (26) located at or near the rotor (22) at the outlet end which receives, via the flow path, a flow of compressible fluid having a first determined subsonic relative velocity vector sum of tangential and meridional velocity, having a respective significant radially outward component, and discharges a flow of compressible fluid having a second determined relative velocity vector having radially outward and tangential components of substantially zero (substantially pure axial flow). Apparatus as claimed in Claim 12 characterised in that the diffuser means includes successive downstream first (72) and second (78) axially spaced apart annular arrays of circumferentially spaced apart diffuser vanes, in which the first diffuser vanes (72) are arranged to receive from the outlet portion a flow of compressible fluid having the first determined relative velocity vector, and to diffuse the fluid flow to discharge into an axially extending interdiffuser space, a flow of the fluid having a second determined relative velocity vector less than the first relative velocity vector and having a radially outward component of substantially zero, and in which the second diffuser vanes (78) are arranged to receive from the interdiffuser space the flow of fluid having the second determined relative velocity vector and to diffuse the latter to discharge a flow of the fluid having a third determined relative velocity vector less than the second relative velocity vector and having both tangential and radially ouward components of substantially zero. A method of pressurising an elastic fluid by means of a rotor (22) within a housing (12), comprising the steps of: forming a tubular stream of the fluid having a radially inner diameter, a radially outer diameter, and a first relative velocity vector sum of meridional and tangential velocities of at least Mach 1.2 at the radially outer diameter; diffusing the fluid to a second supersonic relative velocity less than the first relative velocity while limiting deviation of radially outer local relative velocity vectors to no more than 10 o with respect to the first relative velocity vector; passing the fluid through a normal shock to a third relative velocity of less than Mach 1; and further diffusing the fluid stream while increasing downstream both the radially inner and radially outer diameters thereof to impart a significant radially outward component of meridional velocity thereto.

Inventors: BUSH, DUANE B., DODGE, JOHN L., PECHUZAL, GEORGES A., RAVINDRANATH, AMBRISH
Applicants: ALLIED-SIGNAL INC. (A DELAWARE CORPORATION)
Application year: 1986
Main classification: F04D29
IPC codes: F04D17, F04D29, F04D21