Abstract:
A gas compressor based on the use of a driven rotor having an axially oriented compression ramp traveling at a local supersonic inlet velocity (based on the combination of inlet gas velocity and tangential speed of the ramp) which forms a supersonic shockwave axially, between adjacent strakes. In using this method to compress inlet gas, the supersonic compressor efficiently achieves high compression ratios while utilizing a compact, stabilized gasdynamic flow path. Operated at supersonic speeds, the inlet stabilizes an oblique/normal shock system in the gasdyanamic flow path formed between the gas compression ramp on a strake, the shock capture lip on the adjacent strake, and captures the resultant pressure within the stationary external housing while providing a diffuser downstream of the compression ramp.

Description:
RELATED PATENT APPLICATIONS 
     This application is a Continuation-In-Part of prior U.S. patent application Ser. No. 10/355,784 filed Jan. 29, 2003, now abandoned entitled SUPERSONIC COMPRESSOR, (assigned of record on Mar. 16, 2004 and Mar. 29, 2004 and recorded on Apr. 19, 2004 at Reel/Frame 015229/0879 to Ramgen Power Systems, Inc. of Bellevue, Wash.), which utility application claimed priority from prior U.S. Provisional Patent Application Ser. No. 60/352,943, filed on Jan. 29, 2002, the disclosures of which are incorporated herein in their entirety by this reference, including the specification, drawings, and claims of each application. 
    
    
     STATEMENT OF GOVERNMENT INTEREST 
     This invention was made with United States Government support under Contract No. DE-FC026-00NT40915 awarded by the United States Department of Energy. The U.S. Government has certain rights in the invention. 
    
    
     COPYRIGHT RIGHTS IN THE DRAWING 
     A portion of the disclosure of this patent document contains material that is subject to copyright protection. The applicant no objection to the facsimile reproduction by anyone of the patent document or the patent disclosure, as it appears in the Patent and Trademark Office patent file or records, but otherwise reserves all copyright rights whatsoever. 
     TECHNICAL FIELD 
     This invention relates to the field of fluid compression. More particularly, the invention relates to a high efficiency, novel gas compressor in which saving of power as well as improved compression performance and durability are attained by the use of supersonic shock compression of the process gas. Compressors of that character are particularly useful for compression of air, refrigerants, steam, and hydrocarbons, or other gases, particularly heavy gases. 
     BACKGROUND 
     In axial flow compressors, as employed in conventional gas turbine engines, the Mach number of the flow relative to the individual rotor and/or stator blades is typically in the subsonic or transonic flow regime. Blade tip Mach numbers from about 0.5 Mach to about 0.7 Mach are common. Rotor/stator operation at Mach numbers in this range results in high lift to drag levels, with minimal shock losses for the rotor and stator blades. However, one of the disadvantages of such a design is that the pressure ratio that can be achieved across any given rotor/stator stage is typically limited to about 1.5:1. Yet, simple gas turbine systems that must achieve high cycle efficiency levels require overall compression ratios in excess of about 10:1. Further, systems with compression ratios up to about 25:1 have been demonstrated for applications with demanding performance requirements. Consequently, when there is a demand for high compression ratios, compressors with many stages of compression are provided. Unfortunately, such multi-staged compressors are relatively heavy, complex, and thus are expensive. As a result, there has been a continuing interest in the compressor design field to explore higher rotor/stator loadings, in order to deliver high compression ratios with fewer stages. 
     Further, the limitations imposed by low stage pressure ratios in subsonic compressors have stimulated the study, design, development, and testing of transonic and supersonic flow velocities in the rotor blades. Such a design approach provides a much greater amount of kinetic energy to the gas at each stage. With supersonic compressors of axial or mixed flow (i.e. using combinations of axial and radial flow), past designs have shown the potential to attain stage pressure ratios as high as 4 or even 6, with attendant adiabatic efficiencies of 75% to 80%. In such designs, it follows that high air handling ability is obtained with minimal frontal area, which is particularly important for flight applications. Furthermore, high pressure ratios per stage means that fewer stages are required, with resultant saving in compressor weight and expense. 
     High compression ratios per stage have been provided by several types of designs. In supersonic designs, shock waves may be handled in the stator, or in the rotor, or both. No matter where the shocks occur, the basic design criteria is to minimize the pressure loss and to insure flow stability over as wide of an operating range as possible. One type of prior art blade configuration that accomplishes shock compression within the rotor flow is illustrated in  FIG. 1 . There, a rotor operating at an inlet Mach number of 1.8 provides a first rotor blade R 1  which generates a first shock S 1  which is captured and reflected in the form of a second, oblique shock S 2  by the adjacent rotor blade R 2 . Resultant downstream flow decelerates behind the third shock S 3  to a Mach number of 0.75, and, after expansion to a Mach number of 0.5. For achievable supersonic blade tip Mach numbers, considerable static pressure rise is obtained in the rotor itself. Yet, one of the disadvantages of such a rotor blade design has been the loss incurred in the subsonic diffusion process, due to separation occurring from interaction between the normal shock S 3  and the boundary layer. 
     As an initial step in an attempt to overcome the shortcomings inherent in presently available compressor designs, we have evaluated the performance of various supersonic flight inlets. Most manned aircraft and many missiles rely on some form of air-breathing propulsion for sustained flight within the earth&#39;s atmosphere. Air breathing engines require an inlet to diffuse air from the free-stream velocity to a lower velocity that is acceptable for further processing by other engine components. The inlet components are designed to capture the exact amount of air required, and to accomplish diffusion with a minimum of total pressure loss. Importantly, such inlets also must deliver the air to the subsequent components within acceptable levels of flow distortion. Such inlets must also be configured to contribute the lowest amount of external drag possible for a given application. Because a wide range of supersonic air-breathing propulsion systems have been designed, developed, and tested over the last 60 years, the optimization of supersonic inlet configurations for various flight applications has received a great deal of attention. As a result, many techniques are known in that art for maximizing the performance of supersonic inlets. In fact, performance levels of such inlets have been well established over a wide range of operational flight Mach numbers. 
     Attention is directed to  FIG. 2 , wherein the supersonic inlet performance (as represented by the total pressure recovery) is shown as a function of flight Mach number, for a wide range of supersonic inlet systems. The various inlet performance data points shown on this plot represent actual test data from supersonic inlet systems of many different configurations that have been designed to operate over a wide flight Mach number range, while exposed to a range of angle of attack attitudes and yaw angles. Examples are provided for specific designs by Pratt &amp; Whitney (P&amp;W), United Technologies (UTRC), the NASA Hypersonic Research Engine (HRE), the US Army Transportation Material Command (TMC), and the National Aeronautics and Space Administration (NASA) TMX 413 design. Additionally, the nominal inlet performance requirement set forth in United States Military Specification MIL-E-5007 is illustrated. In short, the peak performance levels of each of the noted systems have been compromised to achieve the robust operability and stability requirements of a true flight system that must complete a mission wherein a wide range of inlet flow conditions are encountered. 
     For categorizing anticipated performance characteristics, and for evaluation and comparison purposes, supersonic flight inlet designs are often defined into three broad groups. These groups are (a) normal shock inlets, (b) external compression inlets, and (c) mixed compression inlets. These three different groups are depicted in  FIG. 3 , along with a relative representation of performance levels for each group as a function of design Mach number. As can be appreciated by reference to  FIG. 3 , each of these three types of inlets has advantages and disadvantages. The normal shock inlet exhibits excellent pressure recovery at relatively low Mach number, but recovery drops markedly as design Mach number increases. The external compression inlet shows good pressure recovery in the Mach 2 range, but also drops off markedly as the operating Mach number increases above this design range. The mixed compression inlet provides acceptable pressure recovery over a relatively broad range of design Mach numbers, but again, efficiency drops off markedly as the operating Mach number reaches 4 or more. 
     Referring again to  FIG. 2 , performance data is included for all three types of inlet that were depicted in  FIG. 3 , as can generally be appreciated by the range provided for flight Mach numbers across which the various designs operate. Importantly, it can be appreciated that a properly designed supersonic inlet that is optimized for operation at a single Mach number (or a small operating Mach number range), with minimal variability in angle or attack or in yaw exposure, could achieve performance levels somewhat greater than the performance levels indicated in  FIG. 2 . 
     Further, in  FIG. 4 , nominal inlet performance requirements as delineated in Military Specification MIL-E-5007 are provided. First, this figure provides a curve that corresponds to the mean line of the flight inlet performance, expressed as static pressure ratio, derived from the curve for the MIL-E-5007 inlet depicted in  FIG. 2 . Second, based on the first derived curve,  FIG. 4  provides a curve that corresponds to the mean line of flight inlet performance expressed as adiabatic (isentropic) compression efficiency as a function of flight Mach number. For example, at a flight Mach number of 2.4, the MIL-E-5007 specification inlet will provide compressor performance at 94% adiabatic efficiency. 
     Total pressure recovery is commonly used by flight inlet designers to evaluate the performance of such systems. The total pressure recovery is the ratio of the total pressure of the flow leaving the inlet to the free stream total pressure level. The flight Mach number can also be thought of as quantifying the magnitude of the total pressure available to an inlet. Thus, it is possible for any given flight Mach number and total pressure recovery level to calculate a corresponding system pressure ratio. The pressure ratio is a critical factor for all compression applications. 
     In  FIG. 5 , the compression efficiency for the flight inlet data shown in  FIGS. 2 and 4  has been averaged, and reduced to a single functional line where adiabatic compression efficiency is illustrated as a function of inlet static pressure ratio. Thus,  FIG. 5  represents the basic efficiency characteristics of a wide range of flight inlets as a function of static pressure ratio. Importantly, this comparison allows the generalized efficiency of flight inlets to be directly compared to the performance of (a) centrifugal compressors, and (b) axial flow compressors, in terms of adiabatic (isentropic) compression efficiency. To facilitate this comparison, the adiabatic (isentropic) compression efficiency of a number of selected axial flow industrial gas turbine compressors are shown in this  FIG. 5 . It can readily be appreciated from  FIG. 5  that supersonic flight inlet designs have the potential to operate at significantly greater efficiencies than heretofore known centrifugal compressors or conventional axial flow turbo-compressors. 
     The isentropic compressor efficiency (sometimes referred to as the adiabatic compressor efficiency) is a performance parameter commonly used by compressor designers. This is based on a theoretical frictionless adiabatic compression process, and thus also is known as isentropic, or having a constant entropy. Although it is evident that compression is not frictionless, and that friction results in heating of the metal parts of the compressor, the assumption that adiabatic compression takes place may be used in computing the theoretical power requirements for a particular compression requirement. The isentropic compressor efficiency is defined as the ratio of isentropic work of compression to the actual work of compression. Equation 1 shows the definition for the isentropic compressor efficiency:
 
ηcomp=( h   t2i   −h   t1 )/( h   t2a   −h   t1 )
 
     Although a variety of supersonic gas compressor and compressor diffuser designs have been heretofore proposed, in so far as we are aware, none have been widely utilized for primary compressor service, whether for common applications such as air or steam, or heavier gases such as certain refrigerants, or heavy chemical intermediates such as uranium hexafluoride. Undoubtedly, improvements in compression efficiency, which would be especially advantageous in order to reduce energy costs for a particular compression service, would be desirable. In various attempts to achieve such improvements, many different methods and structures have been tried, either experimentally or commercially. Some of such attempts have included the use of various shock patterns, such as adapted from conventional centrifugal compressor wheels, or incorporating various multiple rotor configurations. The challenge, however, has been in the selection of methods and structures that assure adequate performance (including such acceptability of attributes such as starting performance, overall efficiency, and acoustic stability) while reducing capital and operating costs. It would be especially desirable for compressors operating under such conditions to have an inlet and diffuser configuration that would be resistant to small changes about a design point with respect to external flow field dynamics and shock perturbations. 
     Consequently, it would be desirable to provide a reliable supersonic gas compressor, specifically including a compressor and diffuser chamber structure that enables the compressor to maintain high isentropic efficiency with a minimum of structure, while operating at a high pressure ratio. Therefore, a continuing demand exists for simple, highly efficient and inexpensive gas compressors as may be useful in a wide variety of gas compression applications. This is because many gas compression applications could substantially benefit from incorporating a compressor that offers a significant efficiency improvement over currently utilized designs. In view of ever increasing energy costs, particularly for both electricity and for natural gas, it would be desirable to attain significant cost reduction in utility expense for gas compression. Importantly, it would be quite advantageous to provide a novel compressor which provides improvements (1) with respect to operating energy costs, (2) with respect to reduced first cost for the equipment, and (3) with respect to reduced maintenance costs. Fundamentally, particularly from the point of view of reducing long term energy costs, this would be most effectively accomplished by attaining gas compression at a higher overall compression efficiency than is currently known or practiced industrially. Thus, the important advantages of a new gas compressor design providing the desirable feature of improved efficiency can be readily appreciated. 
     SUMMARY 
     We have now invented a gas compressor based on the use of a driven rotor having a substantially axial compression ramp traveling at a local supersonic inlet velocity (based on the combination of inlet gas velocity and tangential speed of the ramp) which compresses inlet gas by use of supersonic shock wave located substantially axially between an upstream strake and a downstream strake, while containing the compressed gas via a stationary sidewall housing. In using this method to compress inlet gas, the supersonic compressor efficiently achieves high compression ratios while utilizing a compact, stabilized gasdynamic flow path. Operated at supersonic speeds, the inlet stabilizes an oblique/normal shock system in the gasdynamic flow path formed between the upstream strake and the downstream strake, while retaining the compressed gas against a stationary external housing. And, to provide for ease of start, and to improve operating efficiency, an inlet bleed air system is provided to remove boundary layer air downward through selected rim segments and out through rotor internals. 
     The structural and functional elements incorporated into this novel compressor design overcomes significant and serious problems which have plagued earlier attempts at supersonic compression of gases in industrial applications. First, at the design Mach numbers at which my device can be engineered to operate may be in the range from about Mach 1.5 or slightly lower to about Mach 4.0, the design minimizes aerodynamic drag. This is accomplished by both careful design of the shock geometry, as related to the upstream and downstream strakes and rotating compression ramp, as well as by effective use of a boundary layer control and drag reduction techniques. Thus, the design minimizes parasitic losses to the compression cycle due to the flow field distortion resulting from boundary layers and shock boundary layer interactions. This is important commercially because it enables a gas compressor to avoid large parasitic losses that undesirably consume energy and reduce overall plant efficiency. 
     Also, more fundamentally, this compressor design can develop high compression ratios with very few aerodynamic leading edges. The individual leading edges of the thousands of rotor and stator blades in a conventional high pressure ratio compressor, especially as utilized in the gas turbine industry, contribute to the vast majority of the viscous drag loss of such systems. However, in that the design of the novel gas compressor disclosed herein utilizes, in one embodiment, only a handful of individual aerodynamic leading edges that are subjected to stagnation pressure, viscous losses are significantly reduced, compared to conventional gas compression units heretofore known or utilized. As a result, the novel compressor disclosed and claimed herein has the potential to be much more efficient than a conventional gas turbine compressor, when compared at competing compression ratios. 
     Second, the selection of materials and the mechanical design of rotating components avoids the use of excessive quantities or weights of materials (a vast improvement over large rotating mass bladed centrifugal compressor designs). Yet, the design provides the necessary strength, particularly tensile strength where needed in the rotor, commensurate with the centrifugal forces acting on the extremely high speed rotating components. 
     Third, the design provides for effective mechanical separation of the low pressure incoming gas from the exiting high pressure gases, while allowing gas compression operation along a circumferential pathway. The novel design enables the use of lightweight components in the gas compression pathway. 
     To solve the above mentioned problems, we have now developed compressor design(s) which overcome the problems inherent in the heretofore known apparatus and methods known to us which have been proposed for the application of supersonic gas compression in industrial applications. Of primary importance, we have now developed a low drag rotor which has an upstream strake and a downstream strake, and one or more gas compression ramps mounted on at least one of the upstream strake and/or the downstream strake. A number N of peripherally, preferably partially helically extending strakes S partition the entering gas flow sequentially to the inlet to a first one of the one or more strake mounted gas compression ramps, and then to a second one of the one or more strake mounted gas compression ramps, and so on to an Nth one of the one or more strake mounted gas compression ramps. Each of the strakes S has an upstream or inlet side and a downstream or outlet side. For rotor balance and gas compression efficiency purposes, in one embodiment the one gas compression ramp is provided for each downstream strake. In another embodiment, one gas compression ramp is provided for each upstream strake. In yet another embodiment, one gas compression ramp is provided at each one of the downstream strakes and the upstream stakes. In one embodiment, the number of strakes N and the number X of gas compression ramps R are both equal to three. The pressure inherent in the compressed gases exiting from each compressive shock structure between the upstream and downstream strakes is efficiently captured at one or more diffuser structures located between diverging portions of upstream and downstream strakes. Moreover, the compressed gas is effectively prevented from “short circuiting” or returning to the inlet side of subsequent gas compression ramps by the strakes S. More fundamentally, the strakes S act as a large screw compressor fan or pump to move compressed gases along with each turn of the rotor. 
     To accommodate the specific strength requirements of high speed rotating service, various embodiments for an acceptable high strength rotor are feasible. In one embodiment, the rotor section may comprise a carbon fiber disc. In another, it may comprise a high strength steel hub. In each case, the strakes and accompanying gas compression ramps and diffuser(s) may be integrally provided, or rim segments including strake segments (with or without gas compression ramps) may be releasably and replaceably affixed to the rotor. 
     Attached at the radial edge of the outer surface of the rotor are one or more of the at least one strakes, which strakes each extend further radially outward to a strake tip very closely adjacent the interior peripheral wall of a stationary housing. At least one of the gas compression ramps are situated at one of the downstream or at one of the upstream strakes so as to engage and to compress that portion of the entering gas stream which is impinged by the gas compression ramp upon its rotation, to cause a supersonic shock wave that is captured between adjacent strakes. The compressed gases escape rearwardly from the gas diffuser portion, decelerate, and expand outwardly into a gas expansion diffuser space or volute, prior to entering a compressed gas outlet nozzle. 
     Finally, many variations in the gas flow configuration and providing gas passageways, may be made by those skilled in the art without departing from the teachings hereof. Finally, in addition to the foregoing, this novel gas compressor is simple, durable, and relatively inexpensive to manufacture and to maintain. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWING 
       In order to enable the reader to attain a more complete appreciation of the invention, and of the novel features and the advantages thereof, attention is directed to the following detailed description when considered in connection with the accompanying drawings, wherein: 
         FIG. 1  shows a sectioned view of a set of prior art rotor blades used in one supersonic turbo-compressor design. 
         FIG. 2  illustrates the total pressure recovery versus flight Mach number for a variety of supersonic flight inlet designs, including the United States Military Specification MIL-E-5007 design. 
         FIG. 3  illustrates the total pressure recovery versus design Mach number for three basic supersonic compression inlets, namely (a) a normal shock inlet, (b) an external compression inlet, and (c) a mixed compression inlet. 
         FIG. 4  shows the flight inlet compression performance for a Military Specification MIL-E-5007 flight inlet, showing (a) static pressure ratio versus flight Mach number, and (b) adiabatic pressure efficiency versus flight Mach number. 
         FIG. 5  shows the comparative compression performance of (a) supersonic flight inlets, (b) axial flow compressors, and (c) centrifugal compressors, using a plot of adiabatic compression efficiency versus static pressure ratio. 
         FIG. 6  shows one type of supersonic compressor which utilizes primarily a radial extending shock structure, created by flowing the gas to be compressed along a pre-inlet flow surface, and then against a radially outward compression ramp design to create a series of oblique shocks and a normal shock prior to an expansion portion downstream which provides a subsonic diffuser. 
         FIG. 6A  illustrates the shock structure which is developed when utilizing the supersonic compressor of the design first set forth in  FIG. 6 , as created by flowing the gas to be compressed along a pre-inlet flow surface, and then against a radially outward compression ramp to create a series of oblique shocks and a normal shock prior to an expansion portion downstream which provides a subsonic diffuser. 
         FIG. 7  illustrates a novel mixed compression supersonic compressor wheel which utilizes an axial shock structure, wherein a supersonic compression ramp is provided to create a shock system axially between adjacent strakes. 
         FIG. 8  illustrates a detailed view of the inlet and diffuser portions of the rotary supersonic compressor wheel just set forth in  FIG. 7 , viewed radially inward and looking at the circumference of the compressor wheel, and now showing each rim segment number as manufactured utilizing a plurality of rim segments to define the various aerodynamic components of the strakes, compression ramp, shock capture inlet lip on the upstream strake, and the diffuser which splits gas flow into two flow channels before a large diffusion chamber is reached between the upstream strake and the downstream strake. 
         FIG. 9  illustrates the shock system generated by the compressor design just illustrated in  FIGS. 7 and 8  above, when the inlet is operating at the design Mach number, so that a plurality of oblique shocks, and reflected oblique shocks, are captured between the compression ramp on the downstream strake and the upstream strake. 
         FIGS. 9   a  and  9   b  illustrate the shock system generated by the compressor design just illustrated in  FIG. 9 , but when operating at less than the design inlet flow Mach number, thus showing that the leading shock(s) are not captured by the shock capture lip of the upstream strake. 
         FIG. 10  shows the shock system generated by an internal compression inlet, where the compressive surface is incorporated into the upstream strake. 
         FIG. 11  shows the shock system generated by an internal compression inlet, where the compressive surface is incorporated into the downstream strake wall. 
         FIG. 12  shows the shock system generated by an internal compression inlet, where the compression ramp surfaces are incorporated into both the upstream and downstream strake walls. 
         FIG. 13  provides a vertical cross-sectional view of one embodiment of a supersonic gas compressor, showing the inlet, the upstream strake, the downstream strake, and stationary peripheral housing. 
         FIG. 14  shows a hypothetical vertical elevation view of the circumferential view just provided in  FIG. 8  above, but now showing from the side, as if unrolled, the housing with interior peripheral wall, the outer extremity of the rotor, the downstream strake extending outward to a tip end adjacent the interior peripheral wall, the upstream strake extending outward to a tip end adjacent the interior peripheral wall, shock capture inlet lip on the upstream strake, and in phantom lines, the location of the diffuser in the gasdynamic path. 
         FIG. 15  shows one embodiment for incorporating boundary layer bleed holes into the base of the axial compression ramp, along the ramp itself, and at the throat between the ramp and the upstream strake, and at the adjacent rotor outer surface portion, and additionally shows outlets for discharge of accumulated bleed air from within a rim segment to the adjacent wheel space. 
         FIG. 16  provides a partially cut away perspective view of once embodiment of a compressor utilizing opposing rotors mounted on a common shaft, with each rotor having axial compression ramps as described herein. 
     
    
    
     The foregoing figures, being merely exemplary, contain various elements that may be present or omitted from actual implementations depending upon the circumstances. An attempt has been made to draw the figures in a way that illustrates at least those elements that are significant for an understanding of the various embodiments and aspects of the invention. However, various other elements and parameters are also shown and briefly described to enable the reader to understand how various optional features may be utilized in order to provide an efficient, reliable supersonic gas compressor. 
     DETAILED DESCRIPTION 
     A detailed view of an exemplary embodiment of a supersonic compressor rotor wheel  20  designed for utilization of axial supersonic shock patterns is provided in  FIG. 7 . Rotor disc portion  18  of wheel  20  supports a plurality of rim segments M 1  through M 52  mounted thereon, as further indicated in  FIGS. 8 and 15 . In  FIG. 8 , a series from 1 to 52 of rim segments (M 1  through M 52 ) are described in a circumferential manner as if looking radially face down toward the center  21  of rotor  20 . Inlet fluid (such as air) as indicated by reference arrow  22  is supplied to the pre-inlet flow surface  24  at the outer periphery of the rotor wheel  20 . The inlet fluid encounters a compression ramp  26  provided as a part of downstream strake  28 . 
     A profiled, preferably smoothly curved cowl portion  30  of upstream strake  32 , and having a strake shock capture inlet lip S IN , is provided to capture a series of axially extending oblique shocks (see discussion below in conjunction with  FIGS. 9 ). The compression ramp  26  provided as a part of downstream strake  20  serves to laterally compress inlet air and direct it primarily (substantially uni-directionally) in the direction of reference arrow  34 . Under design supersonic speed inlet conditions, lip S IN  of upstream, inlet strake  32  captures the oblique shockwave and directs entering air between inner wall  40  of upstream strake  30  strake and inner wall  42  of compression ramp  26 . Captured, compressed fluid is eventually diffused via use of diffuser centerbody  44 . In one embodiment, diffuser  44  comprises a substantially triangular structure having a leading edge  46 . A first diffuser sidewall  48  and a second diffuser sidewall  50  act, in conjunction with inner wall portion  52  of upstream strake  32  and inner wall portion  54  of downstream strake  26 , respectively, to provide first  56  and second  58  diffusion channels for the compressed fluid. A rear wall  60  is provided for diffuser  44 . Behind the rear wall  60 , the speed of captured fluid decreases and pressure increases. Compressed fluid is dumped at the exhaust outlet S EX  of the downstream strake  28 . 
     Note that the inlet end S IN  of the upstream, or inlet strake  32  is preferably slightly inward of the outermost point  64  of the strake  30  toward the lateral edge  70  of rotor  20 . This provides a unique contoured inlet cowl shape  30  to capture and compress inlet air, more specifically in the form of a mixed compression supersonic inlet. Such a shape provides for easier self starting and capture of the supersonic shock structure. 
     The compressor design taught herein uniquely applies various techniques of flight inlet design, in order to achieve performance optimization, with the advantages of high single stage pressure ratios, simplicity, and low cost of supersonic compressors to provide a high efficiency, low cost compression system especially adapted for ground based (stationary or mobile) compressor applications. Such a combination requires many novel, unique mechanical and aerodynamic features in order to achieve the aerodynamic requirements for a particular system design, without violating the mechanical design limits necessary to provide a safe, durable, robust compression system that can be manufactured utilizing proven and cost effective manufacturing techniques. 
     One of the primary techniques utilized in the design of the compressors taught herein is to employ certain optimization techniques heretofore employed in supersonic flight inlets within the architecture of an enclosed, rotating disc system. Thus, one common element is to utilize a non-rotating compressor case or housing. As represented in  FIG. 16 , a substantially cylindrical stationary housing  80  having an interior peripheral wall  82  is utilized as one of the boundaries for the gas dynamic flow path. Basically, in the most simple terms, three of the surfaces of the supersonic inlet are formed by the moving surfaces integrated onto the rim of a high speed rotor  20 , and one of the surfaces is the interior peripheral wall  82  of the non-rotating housing  80 . 
     In another design for a compressor, the generated supersonic shocks are generally radial in nature, as is illustrated in  FIG. 6 . In that design, the shocks generated by the compression ramp  90  on the rotor rim  92  coalesce and/or reflect off of the stationary interior wall  94 , as illustrated in  FIG. 6A . 
     However, it has now been found that it is possible to advantageously configure the compressive surfaces in a supersonic compressor so that an oblique shock system is provided that creates a compressive field in the axis of rotation, rather than against the outer stationary wall. The apparatus described above with respect to  FIGS. 7 ,  8 , and  13 - 15  show a suitable mixed compression inlet for use with an axial compression system. Such a mixed compression inlet is shown in additional detail in  FIG. 9 . A mixed compression inlet is one in which part of the shock system is external to the fully enclosed portion of the aerodynamic duct defining the inlet flow path. As was earlier illustrated in  FIG. 3 , mixed compression inlets can be designed to operate with greater efficiencies at higher Mach numbers than normal shock inlets or external compression inlets. Also, operation at higher Mach numbers results in greater compression ratios than internal compression inlets (where all the contraction occurs within the fully enclosed part of the aerodynamic duct defining the inlet flow path), while preserving the ability to swallow the shock system, or “start” without the need for complex variable geometry features. 
     In  FIG. 9 , the downstream strake  28  is provided with compression ramp  26 . A plurality of oblique shock structures  100 ,  102 ,  104 ,  106 , are generated at the design Mach number, which in this case is M=2.5. These shocks  100 ,  102 ,  104 , and  106  are captured by shock lip  30  at the inlet to the upstream strake  32 . A plurality of reflected oblique shocks  110 ,  112 ,  114 ,  116 , and  118  are illustrated downstream. Finally, a normal shock  120  is shown, after which the flow stream is operating at a Mach number of about M=0.75. 
     As shown in  FIGS. 9   a  and  9   b , the oblique shocks generated, i.e.,  122 ,  124 , and  126 , are not captured, or are not completely captured, when compressor design first illustrated in  FIGS. 7 ,  8 , and  14  is operated at less than design Mach number. 
     Turning now to  FIG. 10 , the use of an internal compression inlet is illustrated. Here, the downstream strake  128  does not include a compression ramp. Rather, the upstream strake  132  incorporates a compression ramp  134  having an inlet lip  136 , which generates an oblique shock  138  that is captured by sidewall  140  of downstream strake  128 , and reflected back against compression ramp  134 . After a normal shock  142 , the Mach number is reduced to about M=0.5. 
     In  FIG. 11 , an internal compression inlet is provided. Here, the downstream strake  128  includes a compression ramp  131 . Upstream strake  132  has an inlet lip  136  which captures the oblique shock  144  that is generated by compression ramp  131 . After normal shock  146 , the Mach number is reduced to about M=0.5. 
     In  FIG. 12 , an internal compression inlet is provided where compression ramps  131  and  134  are both provided, incorporated into the downstream  128  and upstream  132  strake walls, respectively. Compression ramps  131  and  134  generate opposing oblique shocks  150  and  152 , which in turn are reflected in shocks  154  and  156 . After normal shock  148 , the Mach number is reduced to about M=0.5. 
     Finally, in  FIG. 13 , a vertical cross section of a portion of one embodiment for a supersonic compressor  200  is provided. The gas compressor  200  includes a circumferential housing  202  having a stationary peripheral wall  204  with an inner surface portion  206  defined by a surface of rotation. An inlet  210  is provided for supply of gas to be compressed. A rotor  20  is provided having a central axis  212  adapted for rotary motion within housing  202  by application of mechanical energy to driving shaft  213 . The rotor  20  extends radially outward from the central axis  212  to an outer surface portion  214 . 
     One or more strakes, and, as illustrated an upstream stake  32  and a downstream stake  28  extend outward from the outer surface portion  214  of the rotor  20  to a tip end  28   T  and  32   T , respectively. Each of the tip ends  28   T  and  32   T  are adjacent the inner surface portion  206  of the stationary peripheral wall  204 . As better seen in  FIG. 8 , at least one of the one or more strakes  28  and  32  further include (i) an upstream end having an inlet S IN , (ii) a supersonic compression ramp  26 , wherein the ramp  26  is oriented to develop an axially oriented supersonic shock (see  FIG. 9 ) during compression of an inlet gas G I . A shock capture lip  30  is provided, axially displaced from the supersonic compression ramp  26  and positioned at a location on the outer surface  24  of the rotor  20  so that the shock compression ramp  26  and the shock capture lip  30  effectively contain a supersonic shock wave  100  (see  FIG. 9 ) therebetween at a selected design Mach number. An outlet diffuser  44  is optionally provided, situated downstream of the supersonic compression ramp  26 . The one or more strakes  28 ,  32 , etc. operate as a helical screw to separate the inlet gas G I  from compressed gas G p  downstream of each one of the supersonic gas compression ramps  26 . Each one of the one or more strakes  28 ,  32 , etc., in one embodiment are configured as a helical structure extending substantially radially from the outer surface portion  214  of the rotor  20  to their respective tip end  28   T  or  32   T . 
     As illustrated, the number of the one or more helical strakes is N, and the number of said one or more supersonic gas compression ramps is X, and N and X are equal—i.e. one gas compression ramp is provided on a downstream portion of each strake. Each one of the one or more gas compression ramps  26  includes an axially directed portion that provides an upstream narrowing gas compression ramp face  220 . 
     As further illustrated in  FIG. 15 , in one configuration each of the one or more gas compression ramps  26  further include one or more boundary layer bleed or holes  230 . In such a configuration, at least one of the one or more boundary bleed holes  230  is located at said base  232  of a gas compression ramp  26 . Also, at least one of the one or more boundary layer bleed holes  230  can be located along the working face  220  portion of the compression ramp  26 . And, at least one or more of the boundary layer bleed holes  230  can be located in the throat  236  area of the compression ramp  26  adjacent the closest approach to the upstream strake  32 . In still another variation, it is advantageous to include at least one of a plurality of bleed holes in the outer surface portion of  24  of the rotor, at a location adjacent each one of the locations of bleed holes in the compression ramp, namely the base  232 , the face  220 , or the throat  236 . Additionally shown in  FIG. 15  are the use of hollow rotor segments M 8 , M 16 , M 17 , and M 18 , which allow passage of bleed gas out into the adjacent wheel space via outlet passages B 9 , -B 12 , B 16 , B 17 , and B 18 , respectively in the direction of reference arrows G B  so that accumulated bleed gas from within a rim segment passes to the adjacent wheel space. 
     Especially where an inlet body diffuser  44  is not utilized, the gas compression ramps  26  may further include (a) a throat  240 , and (b) an inwardly sloping gas deceleration ramp  244 , as indicated in  FIG. 10 , for example. 
     Also, each of the gas compression ramps  26  may further form, adjacent thereto and in corporation with one of said at least one strakes  28  or  32 , a bleed air receiving chamber  250 . Each of the bleed air receiving chambers  250  effectively contains therein, for ejection therefrom, bleed air routed thereto from the bleed ports  230 , such as located on face  220 . 
     Returning now to  FIG. 13 , the apparatus also includes a gas outlet  252  for receiving and passing therethrough high pressure outlet gas G p  resulting from compression of inlet gas G I . 
     The apparatus just described includes supersonic shock compression of inlet gas G I , utilizing the apparent velocity of gas entering the one or more gas compression ramps in excess of Mach 1. In another embodiment, the apparent velocity of gas entering the one or more gas compression ramps is in excess of Mach 2. In another embodiment, the design apparent velocity of gas entering the one or more gas compression ramps is between about Mach 1.5 and Mach 3.5. 
     A gas compressor configured as described herein may be provided specifically engineered to compress any selected gas, including a gas selected from the group consisting of (a) air, (b) refrigerant, (c) steam, and (d) hydrocarbons. Importantly, the compressor may compress such gases at a selected isentropic efficiency in excess of ninety (90) percent. In some cases, the compressor will compress a selected gas at an isentropic efficiency in excess of ninety five (95) percent. 
     Again, as noted in  FIG. 13 , part of the reason that such high efficiency can be attained is that the rotor includes a central disc portion that is confined within a close fitting housing having a minimal distance D between the rotor  20  the housing  260 , so as to minimize aerodynamic drag on the rotor  20 . 
     In an advantageous method of compressing gas, one or more gas compression ramps are provided on a rotor which is rotatably secured with respect to stationary housing having an inner surface. Each of the gas compression ramps is provided with an inlet gas stream, which stream is compressed by one or more gas compression ramps and contained by a stationary housing, to generate a high pressure gas G P  therefrom; The high pressure gas is effectively separated from low pressure inlet gas G I  by using one or more strakes along the periphery of a rotor. The strakes are helically offset by an angle delta (Δ), as indicated in  FIG. 8 . Each one of the one or more strakes are provided adjacent to one of one or more gas compression ramps. At least a portion of each of the one or more strakes extend outward from at least a portion of an outer surface portion of the rotor to a point adjacent an inner surface of a stationary housing. Mechanical power is applied to an input shaft that operatively drives the rotor and thus drives the one or more gas compression ramps. In practice of the method, the apparent inlet velocity of the one or more gas compression ramps is at least Mach 1.0. In one aspect of the method, the apparent inlet velocity of the one or more gas compression ramps is at least Mach 2.5. In another embodiment of the method, the inlet velocity of the one or more gas compression ramps is between Mach 2.5 and Mach 4. In yet another embodiment, the apparent inlet velocity of the gas compression ramps is approximately Mach 3.5. In practice of the method, a gas being compressed can be selected from the group consisting of (a) air, (b) steam, (c) refrigerant, and (d) hydrocarbons. In one embodiment the gas is essentially natural gas. In another embodiment, the method can be practiced to compress air. In yet another embodiment, the method can be practiced to compress a refrigerant. In a still further embodiment, the method can be practiced to compress steam. For aerodynamic and acoustic purposes, the compression ramps can be arranged and spaced equally apart circumferentially about a rotor so as to engage a supplied gas stream substantially free of turbulence from the previous passage through a given circumferential location of any one of the one or more gas compression ramps. In design of a suitable supersonic gas compressor as taught herein, the cross sectional areas of each of the throat resulting at one of the one or more gas compression ramps is sized and shaped to provide a desired compression ratio. 
     Turning now to  FIG. 16 , a partially cut away perspective view of one embodiment of a compressor  21  utilizing opposing rotors mounted on a common shaft is provided. Here, each rotor has axial compression ramps  26  as described herein, but mounted in opposing fashion along a common shaft for thrust balancing. Major components shown in this  FIG. 16  include a stationary housing or case  322  having first  324  and second  326  inlets for supply of low pressure gas to be compressed, and a high pressure compressed gas outlet nozzle  328 . In this dual unit design, a first rotor  330  and a second rotor  332  are provided, each having a central axis defined along centerline  334 , here shown defined by common shaft  336 , and adapted for rotary motion therewith, in case  322 . Each one of the first  330  and second  332  rotors extends radially outward from its central axis to an outer surface portion  338 , and further to an outer extremity  340  on the strakes S. On each one of first  330  and second  332  rotors, one or more axially directed supersonic shock compression ramps  26  are provided. Each one of the axially directed supersonic shock compression ramps  26  forms a feature extending outward from the outer surface portion  338  of its respective first  330  or second  332  rotor. Within housing  322 , a first circumferential stationary interior peripheral wall  342  is provided radially outward from first rotor  330 . Likewise a second circumferential stationary interior peripheral wall  344  is provided radially outward from second rotor  332 . Each one of the stationary peripheral walls  342  and  344  are positioned radially outward from the central axis defined by centerline  334 , and are positioned very slightly radially outward from the outer extremity  340  of first  330  and second  332  rotors (i.e. tips of strakes) respectively. Each one of the first and second stationary peripheral walls  342  and  344  have interior surface portion  352  and  354 , respectively. Each one of the one or more supersonic shock compression ramps  346  cooperates with the interior surface portion  352  and  354  of one of the stationary peripheral walls  342  or  344  to contain gas which has been compressed by the axially directed compression ramp  346 . 
     One or more helical strakes  28  and  32  are provided adjacent each one of the one or more supersonic compression ramps  26 . An outwardly extending wall portion  28   W  or  32   W  of each of the one or more strakes  28  or  32  extends outward from at least a portion of the outer surface portion  338  of its respective rotor  330  or  332  along a height HH to a point adjacent the respective interior surface portion  352  or  354  of the peripheral wall  342  or  344 . The upstream strakes  32  and the downstream strakes  28  effectively separate the low pressure inlet gas G I  from high pressure compressed gas G P  downstream of each one of the supersonic gas compression ramps  26 . Strakes  28  and  32  are, in the embodiment illustrated by the circumferential flow paths depicted in  FIGS. 7 and 8 , provided in a helical structure extending substantially radially outward from the outer surface portion  24  of its respective rotor  330  or  332 . In one embodiment, such as is shown in  FIG. 9 , the number of the one or more helical strakes is N, and the number of the one or more supersonic gas compression ramps is X, and the number N of strakes S is equal to the number X of compression ramps R. In another embodiment, as is shown in  FIG. 12 , the number of helical strakes is N, and the number of the one or more supersonic gas compression ramps is equal to  2 N. When strakes are designated by the reference numeral S, the strakes S 1  through S N  partition entering gas so that the gas flows to the respective gas compression ramp then incident to the inlet area for that rotor. As can be appreciated from  FIG. 8 , the preferably helical strakes, such as strakes S 1 , S 2 , and S 3  as shown in  FIG. 7 , are thin walled, with about 0.15″ width (axially) at the root, and about 0.10″ width at the tip. With the design illustrated herein, it is believed that leakage of compressed gases will be minimal. Thus, the strakes S 1  through S N  allow feed of gas to each gas compression ramp without appreciable bypass of the compressed high pressure gas to the entering low pressure gas. That is, the compressed gas is effectively prevented by the arrangement of strakes S from “short circuiting” and thus avoids appreciable efficiency losses. This strake feature can be better appreciated by evaluating the details shown in  FIG. 16 , where strakes  28  and  32  revolves in close proximity to the interior wall surface  352 . The strakes  28  and  32  have a localized height HS 1  and a localized height HS 2 , respectively, which extends to a tip end TS 1  and TS 2  respectively, that is designed for rotation very near to the interior peripheral wall surface of housing  22 , to allow for fitting in close proximity to the tip end TS 1  or TS 2  with the adjacent wall. 
     As depicted in  FIG. 16  downstream of each of first  330  and second  332  rotors is a first  390  and second  392  high pressure outlet, respectively, each configured to receive and pass therethrough high pressure outlet gas resulting from compression of gas by the one or more gas compression ramps  26  on the respective rotor  330  or  332 . One or more combined high pressure gas outlet nozzles  328  can be utilized, as shown in  FIG. 16 , to receive the combined output from the first and second high pressure outlets  390  and  392  from rotors  330  and  332 . 
     For improved efficiency and operational flexibility, the compressor  20  may be designed to further include a first inlet casing  400  and a second inlet casing  402  having therein, respectively, first  404  and second  406  pre-swirl impellers. These pre-swirl impellers  404  and  406  are located intermediate the low pressure gas inlets  324  and  326 , and their respective first  330  or second  332  rotors. Each of the pre-swirl impellers  404  and  406  are configured for compressing the low pressure inlet gas G I  to provide an intermediate pressure gas stream IP at a pressure intermediate the pressure of the low pressure inlet gas G I  and the high pressure outlet gas G P , as noted in  FIG. 16 . In one application for the apparatus depicted, air at ambient atmospheric conditions of 14.7 psig is compressed to about 20 psig by the pre-swirl impellers  404  and  406 . However, such pre-swirl impellers can be configured to provide a compression ratio of up to about 2:1. More broadly, the pre-swirl impellers can be configured to provide a compression ratio from about 1.3:1 to about 2:1. 
     Also, for improving efficiency, the gas compressor  21  can be provided in a configuration wherein, downstream of the pre-swirl impellers  404  and  406 , but upstream of the one or more gas compression ramps  26  on the respective rotors  330  and  332 , a plurality of inlet guide vanes, are provided, a first set  410  or  410 ′ before first rotor  330  and a second set  412  or  412 ′ before second rotor  332 . The inlet guide vanes  410 ′ and  412 ′ impart a spin on gas passing therethrough so as to increase the apparent inflow velocity of gas entering the one or more gas compression ramps  26 . Additionally, such inlet guide vanes  410 ′ and  412 ′ assist in directing incoming gas in a trajectory which more closely matches gas flow path through the ramps  26 , to allow gas entering the one or more gas compression ramps  26  to be at a suitable angle, given the design rotating speed, to minimize inlet losses. 
     In one embodiment, as illustrated, the pre-swirl impellers  404  and  406  can be provided in the form of a centrifugal compressor wheel. As illustrated in  FIG. 16 , pre-swirl impellers  404  and  406  can be mounted on a common shaft  336  with the rotor  330  and  332 . It is possible to customize the design of the pre-swirl impeller and the inlet guide vane set to result in a supersonic gas compression ramp inlet inflow condition with the same pre-swirl velocity or Mach number but a super-atmospheric pressure. Since the supersonic compression ramp inlet basically multiples the pressure based on the inflow pressure and Mach number, a small amount of supercharging at the pre-swirl impellers can result in a significant increase in cycle compression ratio. 
     With (or without) the aid of pre-swirl impellers  404  and  406 , it is important that the apparent velocity of gas entering the one or more gas compression ramps  26  is in excess of Mach 1, so that the efficiency of supersonic shock compression can be exploited. However, to increase efficiency, it would be desirable that the apparent velocity of gas entering the one or more gas compression ramps  26  be in excess of Mach 2. More broadly, the apparent velocity of gas entering the one or more gas compression ramps  26  can currently practically be between about Mach 1.5 and Mach 3.5, although wider ranges are certainly possible within the teachings hereof. 
     It is to be appreciated that the various aspects and embodiments of the supersonic compressor designs described herein are an important improvement in the state of the art of gas compressors. Although only a few exemplary embodiments have been described in detail, various details are sufficiently set forth in the drawings and in the specification provided herein to enable one of ordinary skill in the art to make and use the invention(s), which need not be further described by additional writing in this detailed description. Importantly, the aspects and embodiments described and claimed herein may be modified from those shown without materially departing from the novel teachings and advantages provided by this invention, and may be embodied in other specific forms without departing from the spirit or essential characteristics thereof. Therefore, the embodiments presented herein are to be considered in all respects as illustrative and not restrictive. This disclosure is intended to cover the structures described herein and not only structural equivalents thereof, but also equivalent structures. Numerous modifications and variations are possible in light of the above teachings. It is therefore to be understood that within the scope of the appended claims, the invention(s) may be practiced otherwise than as specifically described herein. Thus, the scope of the invention(s), as set forth in the appended claims, and as indicated by the drawing and by the foregoing description, is intended to include variations from the embodiments provided which are nevertheless described by the broad interpretation and range properly afforded to the plain meaning of the claims set forth below.