Shockwave-induced boundary layer bleed

An apparatus and method is provided for improving efficiency of a transonic gas turbine engine compressor by bleeding off a shockwave-induced boundary layer from the gas flow passage of the compressor using an array of bleed holes having a downstream edge aligned with a foot of an oblique shock wave which originates on the leading edge of an adjacent transonic rotor blade tip.

TECHNICAL FIELD

The present invention relates generally to gas turbine engines, and more particularly to an improved compressor therefor.

BACKGROUND OF THE INVENTION

Bleeding air from gas turbine engine compressors is well known. Air bled from the compressor can sometimes be used to provide a source of pressurized and/or cool air to the engine or the aircraft, however air is principally bled from the compressor in order to improve the operating envelope and overall compressor efficiency, which is often expressed as improved surge margin. Increased incidence angle between the airflow and the blade leading edges at “off design” conditions tends to cause separation of the flow on the suction side of the blades, which results in blade stall and eventually complete surging of the compressor. By bleeding off this stalled airflow adjacent the blade tips, the surge margin of the compressor is thus increased. This accordingly improves the overall efficiency of the compressor.

However, separation of airflow on the compressor blades can also result from factors other than increased blade leading edge incidence. Particularly, the interaction between the boundary layer formed on a stationary outer shroud and a shock wave produced by supersonic compressor blade tips rotating within the shroud, also tends to cause additional flow separation which can result in blade stall and to full compressor surge. Although the inlet flow may be subsonic in a subsonic compressor, the flow relative to the rotor blade tips of a high speed compressor can nevertheless become supersonic, causing separation-inducing shock waves at the blade tips.

Accordingly, there is a need to provide an improved compressor which addresses these and other limitations of the prior art, and it is therefore an object of this invention to do so.

SUMMARY OF THE INVENTION

It is an aim of the present invention to provide a gas turbine engine compressor having improved efficiency.

In accordance with a first aspect of the present invention, there is provided a transonic gas turbine engine compressor comprising: a rotor having a central axis of rotation and a plurality of blades extending into a gas flow passage through said compressor, each of said blades having a blade tip and a leading edge defined between opposed pressure and suction surfaces, said rotor being rotatable about said axis of rotation at a speed such that gas flow adjacent said blade tips becomes supersonic, creating oblique shock waves originating at said leading edge and terminating at a shock foot on said suction surface of an adjacent blade; an outer shroud surrounding said rotor, said outer shroud defining a radially outer boundary of said gas flow passage; a plurality of bleed holes extending through at least a portion of said outer shroud adjacent said blade tips to provide gas flow communication between said gas flow passage and an outer bleed passage, said plurality of bleed holes defining a bleed hole array; and said bleed hole array defining a downstream edge thereof substantially aligned in a gas flow direction with said shock foot and extending upstream thereof, said bleed holes in said array being selected in size, number and location to bleed at least a portion of a shockwave-induced boundary layer from said gas flow passage adjacent said outer shroud.

There is also provided, in accordance with a second aspect of the present invention, a method of bleeding a shockwave-induced boundary layer from a transonic gas turbine engine compressor comprising: providing a rotor having a plurality of blades extending into a gas flow passage of the compressor and an outer shroud surrounding said rotor; rotating said rotor such that gas flow adjacent tips of said blades becomes supersonic, creating an oblique shockwave originating at a leading edge of each said tip and terminating at a shock foot on a suction surface of an adjacent blade; providing a plurality of bleed holes in said outer shroud in a predetermined region corresponding to a boundary layer induced by said oblique shock wave; and bleeding said shockwave-induced boundary layer out of said gas flow passage.

Further details of these and other aspects of the present invention will be apparent from the detailed description and Figures included below.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

The compressor section14is typically a multi-stage compressor, and thus may comprise several axial and/or centrifugal compressors. Although the present invention is preferably adapted for use with an axial turbomachine rotor, and will therefore be described below with regards to its use in an axial compressor, it is to be understood that the use of the present invention in a centrifugal compressor and/or a mixed flow rotor is also envisaged. The present invention is also intended for transonic compressor rotors rather than fully supersonic ones. The term transonic rotor, as used herein, is defined as a rotor having generally subsonic inlet flow, but wherein a relative Mach number of the flow near at least a portion of blade tip region is supersonic.

Referring toFIGS. 2 and 3, an axial compressor20of the compressor section14comprises generally a rotor24and a stator22downstream relative thereto, each having a plurality of blades defined within the gas flow path downstream of the compressor inlet passage42and upstream of the compressor discharge passage21. The gas flowing in direction25is accordingly fed to the compressor20via the compressor inlet passage42and exits therefrom via the compressor discharge passage21. The rotor24rotates about a central axis of rotation23within the stationary and circumferentially extending outer shroud32, which defines a radial outer boundary of the annular gas flow path through the compressor20. The rotation of the rotor24is shown inFIG. 3by displacement of the blades28in direction27relative to the airflow direction25. The rotor24includes a central hub26and a plurality of blades28radially extending therefrom and terminating in blade tips30immediately adjacent the outer shroud32. Each blade28includes a leading edge46defined between a pressure surface48and a suction surface50, as best seen inFIG. 3.

The stationary outer shroud32defines a plurality of bleed holes36in a portion thereof immediately adjacent the blade tips30of the rotor24. The bleed holes36provide fluid flow communication between the annular main compressor gas path, defined radially within the outer shroud32, and a bleed passage38disposed radially outward from the shroud32. The bleed passage38is preferably disposed adjacent to the rotor24, and may define an annular cavity which receives the bleed air. Thus, to bleed holes36passively bleed air from the compressor20, more specifically the bleed holes36bleed shock-wave induced airflow separation therefrom, as will be described in greater detail below. The air bled via the bleed holes36into the bleed passage38is subsequently re-introduced back into the main gas path of the compressor, preferably into the compressor inlet passage42through at least one bleed exhaust opening40located just upstream of the rotor24. As air is passively bled by the bleed holes36, there is accordingly a continuous bleed of air from the main gas path of the compressor. Such a passive bleed design allows for self regulation of the bleed flow based on flow conditions and shock properties of the assembly. As these can be predetermined by one skilled in the art, the necessary amount of bleed flow can be established and regulated to bleed off substantially only shockwave-induced boundary layer from the gas path of the transonic compressor rotor. Generally, only a very small amount of bleed flow is required to achieve this, much less than many compressor bleed systems. At least less than 5% of to total flow through the compressor is bled off by bleed holes36, but more preferably only about 1% of the total flow through the compressor is bled off by the bleed holes36. This is accordingly significantly less than traditional compressor bleed systems which bleed off separated flow in a conventional subsonic compressor, in which typically between 5% and 25% of the total flow is bled off. Air bled by bleed holes36is also dumped back into the main gas flow, preferably upstream from the rotor28, and therefore no further uses for the small amount of air bleed from the compressor are intended. In contrast, prior art compressor bleed designs often route bled air to other parts of to turbomachine for alternate uses, such as cooling airflow and the like. Such designs are significantly more complex, and therefore heavy and expensive, and require a considerably greater volume of bleed airflow to supply such alternate uses adequately.

Referring now toFIG. 3in greater detail, the plurality of bleed holes36are arranged in a predetennined region of the shroud32relative to the rotor blades28, such that the bleed holes36provide passive bleeding of a shockwave-induced boundary layer from the gas path flow adjacent the outer shroud32. As noted above, even when compressor inlet flow is subsonic, the flow near the rotor blade tips30of the high speed compressor20can nevertheless become supersonic, causing oblique shock waves to form at the leading edges46of the blade tips30. Generally, oblique shock waves44are formed at the leading edge46of a first blade tip and terminate with a shock foot45on the suction surface50of an adjacent blade tip. One skilled in flat art will appreciate that the design of the compressor rotor24is such that the shocks formed at the transonic blade tips30are oblique shocks rather than normal shocks, to prevent significant loss of compressor efficiency.

However, such oblique shock waves44interact with the flow boundary layer, which forms on the inner surface34of the stationary outer shroud32, to cause an additional separation of flow induced by the shock-wave. The bleed holes36are therefore positioned such that this shockwave-induced flow separation is passively bled out of the compressor gas flow path, thereby at least reducing and delaying blade stall. This accordingly provides an improved surge margin, and therefore an improved overall efficiency, for such a transonic compressor. Additionally, by substantially eliminating, or at least significantly reducing, the growth of this shock-induced boundary layer, flow diffusion in the compressor gas path passage is enhanced and the flow separation at “off-design” conditions is thus delayed. Although work is done by the compressor on the air which is then bled off, the overall effect on the net efficiency loss of the compressor section will be minimal, due to the counteracting efficiency improvement for downstream compressor blade rows which is a result of bleeding off the flow separation caused by the interaction between the oblique shock waves and the boundary layer at an upstream blade row. In fact, the net compressor stage efficiency can actually improve, while the surge margin gain provided is nevertheless maintained. Conversely, a more loaded blade passage can be permitted with similar surge margins.

As shown inFIG. 3, the bleed holes36are defined in the outer shroud32in an array37of a plurality of holes, the array being preferably provided about the full circumference of the outer shroud32. The bleed hole array37is located, relative to a fluid flow direction25, between the leading edges46and tailing edges47of the rotor blades28. The array37of bleed holes36preferably includes at least several rows of holes deep in an axial direction parallel to the flow direction25, and defines a band about the outer shroud32.

More specifically, the array37of bleed holes36is positioned in the outer shroud32such that a downstream edge39of the array of holes is aligned, in a fluid flow direction25. with each oblique shock wave foot45on the blade suction surfaces50An upstream edge41of the array37of bleed holes36is preferably disposed just downstream from the leading edges46of the blades28. As such, the array37of bleed holes36is disposed within a shock-induced boundary layer region, axially defined between the leading edges46of the blade tips30and the location of the oblique shock foots45on the blade suction surfaces50. As noted above, the number and size of the bleed holes36are preferably defined such that approximately 1% of the total main flow through the compressor is bled off. In a particular embodiment the array37defines at least three rows of bleed holes36between the downstream edge39and the upstream edge41.

Referring now toFIGS. 4 and 5which depict an alternate compressor60which is similar to the compressor20described above, but having an alternate bleed hole configuration. Particularly, a plurality of bleed holes62are provided in the outer shroud32in an array63which, as per the first embodiment described above, defines a downstream edge65which is approximately aligned, in a fluid flow direction25, with the shock wave foots45on the blade suction surfaces50. The upstream edge67of the array63of bleed holes62is also preferably disposed just downstream from the leading edges46of the blades28.

As seen inFIG. 5, the array or band63between adjacent blades28is disposed at an angle relative to a plane perpendicular to the flow direction25. In contrast, the array37of bleed holes depicted inFIG. 3is oriented substantially parallel to such a plane perpendicular to the direction of fluid flow through the compressor. However, both arrays37,63of bleed holes are nevertheless positioned in the outer shroud such that their downstream edges39,65are substantially aligned, in a fluid flow direction25, with the shock wave foots45on the blade suction surfaces50.

Thus the bleed holes36,62are thus defined in a region of outer shroud32corresponding to the area in which an oblique shook forms when flow adjacent the blade tips becomes supersonic. Although the pressure downstream of the bleed holes and the particular pattern of the shock wave firmed will vary the specific hole geometry and position, one skilled in the art will appreciate that the location of the oblique shock foot on the blade suction surfaces may be determined. Therefore the arrays37,63of bleed holes36,62may be thus suitably positioned in a region of the outer shroud32which is defined at least upstream of the foot45of the oblique shook wave formed at each supersonic blade tips30of the transonic rotor.

The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. For example, as noted above, although the present invention is preferably adapted for use with an axial compressor, it may also be employed in a centrifugal compressor or a mixed flow rotor. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the equivalents accorded to the appended claims.