CA1241275A - Contoured endwall turbine stator - Google Patents

Abstract

CONTOURED ENDWALL TURBINE STATOR

ABSTRACT OF THE DISCLOSURE
A subsonic, radial inflow, compressible flow turbine for turbomachinery having an inlet nozzle stator ring with complementally curved endwall and stator vanes.

Description

US

CONTOURED END WALL TURBINE STATOP~

BACKGROUND OF THE INVENTION
This invention relates to gas turbo machinery and relates more particularly to improved radial turbines and methods associated therewith.
Minimization of aerodynamic losses for compressible flow to brines and the like are critical for producing a turbo machine having acceptance efficiency performance. For radial inflow turbines it is conventional to provide a radial fluid flow nozzle section immediately upstream of the inlet to the turbine to optimize both the velocity of the incoming fluid flow in relation to the optimum design point operation of the turbine, as well as to tend to alter the circumferential velocity of the fluid flow in relation to the turbine speed for maximizing energy I transfer from the fluid flow to drive the turbine. In certain circumstances this is accomplished by use of a v~neless nozzle space which effectively transforms fluid flow potential energy in the form of a pressure head, to kinetic energy by acceleration of the fluid flow. In other circumstances stators vanes are included in the nozzle section to divide the latter into a plurality of nozzles each of which is operable to radially, and usually circumferential, accelerate the incoming fluid flow so that its net speed and direction are optimized to impart maximum energy transfer to the turbine wheel.
z5 For high speed, subsonic radial turbines as required in present day turbo machinery such as gas turbine engines, it has - been conventional that requirements for minimizing aerodynamic losses as discussed above dictate straight, if not parallel, axially spaced endless in the nozzle section to minimize frictional

2~7S

and boundary layer diffusion losses within the nozzle, whether of the vane or vinyls type. Such has also been conventional practice to alleviate secondary losses as a result of pressure and velocity differentials across the nozzles in directions perpendicular to the primary flow direction there through.
In some turbo machinery designs, however, it is believed that flow blockage near the nozzle entrance, in the case of vane nozzles as may occur generally at the stators vane leading edges, promotes relative y large velocity differentials between -the pressure and suction sides of adjacent vanes. Velocity peaks introduced by such tendency toward flow blockage necessarily promotes subsequent downstream diffusion transforming fluid velocity back into relatively decelerated, higher pressure conditions along the vane, or at best reduces the desired acceleration before the fluid flow reaches the vane throat.
SUMMARY OF TOE It NOTION
The present invention contemplates complemental contoured, mirror image endless for a nozzle section that directs fluid flow to the inlet of a radial turbine wheel of a turbo machine, and method for minimizing aerodynamic losses into a radial turbine wheel by delaying radial acceleration of fluid flow in the entrance region of the radial inflow nozzle section, while subsequently smoothly radially accelerating the fluid flow to the exit throat of the nozzle. Such arrangement and method provides an optimum I balance between diffusion losses and wall friction in the nozzle section while controlling the velocity level at the entrance thereof, such as at the leading edges of the stators vanes, to I. prevent downstream diffusion and promote ideal, smooth acceleration to the nozzle throat. At the same time, the present invention controls over-velocity at the vane leading edges to prevent 124 1;~'7S

excessive pressure and velocity differentials across the nozzles in circumferential and axial or span-wise airctions. The present invention has been found to reduce the maximum pressure differential between the pressure sides and suction sides of the adjacent vanes to minimize secondary flows.
These and other objects and advantages of the present invention are set forth in or will be apparent from the following detailed description of a preferred embodiment of the invention when read in conjunction with the accompanying drawings.

BRIEF DESCRIPTION OF TOE DRAWINGS
Fly. l is a partial cross-sectional elevation Al view of a gas turbine engine embodying the present invention with portions show schematically;
Fig. 2 is an enlarged cross-sectional view of a portion of the engine of Fig. l;
Fig. 3 is a partial perspective view of the turbine stators with one end wall thereof removed for clarity of illustration;
Fig. 4 is a graph depicting total aerodynamic losses across the axial width or span of the nozzle of the present invention, with conventional non-contoured end wall nozzle efficiency illustrated in dashed lines for comparison purposes; and Fig. 5 is a graph depicting pressure distribution of the nozzle of the present invention, with comparable data of a non-contoured end wall nozzle again shown in dashed lines.

DETAILED DESCRIPTION OF TOE PREFERRED EMBODIMENT
- Referring now more particularly to the drawings, high speed fluid turbo machinery is illustrated in partially schematic form in Fig. l as a gas turbine engine lo having a fluid compressor 1~24~L2~5 12 for compressions and delivering pressurized flow to a combustor 13 of which a portion of the combustor wall 14 is shown. In a conventional fashion fuel from a source 16 is also delivered to the combustor; a combustion process is maintained therein and the heated exhaust gases therefrom are directed to a turbine 13 for driving the latter. Normally, rotational energy of turbine 18 is used to drive the compressor 12 through a shaft 20 and to perform other useful work. Gas exhausting from turbine 18 is directed through a downstream passage 22 to perform additional useful work and/or for exhaust.
Turbine 18 is a radial inflow, centripetal turbine wheel having a radially outer, circumferential extending inlet for receiving gas flow from combustor 13, and a plurality of circus mferentially spaced blades 24 disposed in momentum exchange relationship with the gas flow for driving the entire turbine wheel. A housing generally designated by the numeral 26, to which both turbine 18 and compressor 12 are mounted on bearings such as 28 for rotation about the axis of shaft 20r defines a flow passage in which turbine blades 24 are closely fitted.
Intermediate combustor 13 and turbine 18 is a circus mferentially extending, radial nozzle section 30 bounded axially by a pair of axially spaced generally parallel, facially extending endless 32 and 34. The nozzle section extends radially inwardly to a location closely adjacent the radially cuter inlet to turbine 18 in surrounding relationship to the turbine inlet. In the embodiment illustrated, the nozzle 30 is a vane nozzle having a plurality of equally spaced stators vanes 36. Vanes 36 -ore - curved circumferential, as best illustrated in Fig. 3, for accelerating the fluid flow circumferential, shown by an arrow in Fig. I to provide optimal matching of the fluid flow entrance '5 direction in relation to the speed of the turbine at its design point speed to produce maximum energy transfer from the fluid to the turbine, and thus maximum aerodynamic efficiency. As a result, as well known, each vane presents a pressure side 44 and suction side 46 at relatively high and low pressures respectively, and define nozzle spaces or nozzles there between.
Endless 32 and 34 are complemental contoured with compound curvature to define an outer section 38 of substantially constant width, an inner section 40 of substantially constant- -width less than that of section 38, and a smoothly curved variable width transition section 42 disposed between sections 38 and 40. Compound curvature within transition section 42 presents a concave section and downstream convex section in relation to fluid flow through the nozzles defined between each stators vane 36. The contraction in flow area due to the compound curvature is substantial r with the axial span or width of downstream section - 40 being about 75 percent of the axial width of upstream section 38~ This percentage can be in the range of about 85 percent to about 50 percent dependent upon the aerodynamic characteristics of the turbo machine. As is evident, the stators vanes are likewise contoured at the endless by extending from the outer section 38 through the transition section 42 and into the inner section 40. With inclusion of stators vanes 38, the nozzle section thereby presents a stators ring or nozzle ring through which the subsonic motive fluid flow is delivered to turbine 18. The compound curvature may also be substantial in the radial direction, extending radially to as little as 20 percent, or even somewhat less, of - the projected meridional length of the stators vanes the radial length thereof in the direction as viewed in Figs. 1 and 2), and preferably extends about 30 percent of the meridional length of the stators vanes.
- I

ll'~41'~S

In operation, was flow from combustor 13 enters nozzle section 30 and is smoothly turned and accelerated therein prior to delivery to turbine 18. Subsonic, compressible gas flow is first delivered radially inwardly to the nozzles defined between adjacent stators vanes I Due to the circumferential angulation of vanes 36, angular acceleration is imparted to the fluid flow by its interaction with vanes 36, primarily on pressure sides I The flow is thereby turned to optimally match the effective entrance angle of the turbine flow passages between blades 24-so as to impart maximum momentum exchange to blades 24.
By virtue of the greater axial width in outer section 38, the nozzle section minimizes flow blockage at its entrance' and in particular is sized to avoid vane leading edge flow blockage.
As a result, increase in flow velocity in outer section 38 is deterred or delayed. This is clearly depicted in Fig. 5 which is a plot of flow velocities measured at the pressure and suction sides along the meridional length of adjacent stators vanes 36 defining the nozzles there between, the pressure measured for the structure of the present invention being shown in solid lines while the same pressures for straight endless, typical of prior art, under the same conditions shown in dashed lines.
In the outer section 38 adjacent the stators vane leading edge 48, both the pressure side and suction side velocities are reduced as expected. Yet the difference between the two velocities is significantly less than the prior art conditions. This reduced difference in pressure and suction side velocities minimizes secondary flow migration across the nozzles from the pressure = side of one vane toward the suction side of the adjacent vane.
Through transition section 42, as shown in Fig.- 5, both velocities increase at a somewhat uniform rate in comparison to 2~5 one another to maintain a reduced difference in pressure side and suction side velocities. Yet, because end wall curvature may itself introduce cross-passage velocity variation, the smooth curvature of the transition section plus limitation as to the axial contraction produced thereby in relation to the projected meridional length is controlled to assure that end wall velocities do not become dominant in generating frictional and diffusion losses within the nozzles.
Inspection of Fig. 5 clearly shows that continuous velocity increases on both the pressure and suction sides are developed.
Thus, downstream diffusion and declaration within nozzle section 30 is avoided, and favorable acceleration of the fluid flow throughout the lengths of the subsonic nozzles is achieved through and beyond the throat 50 of the nozzles all the way to the trailing edges 52 of the vanes where, of course, the pressure and suction side effects disappear and their two velocities become equal.
Throughout the downstream section 40, due to lesser radial distances, the nozzles between the vanes 36 smoothly reduce in cross-sectional area to relatively smoothly radially accelerate this fluid flow while the vanes are still in momentum exchange relationship for turning the flow to the desired air-cumferential angulation for the flow exiting the nozzle section.
As a result, secondary fluid flow across the nozzles is minimized throughout the length of the nozzle section 30.
Fugue is a plot of aerodynamic losses attributable to stators vanes 36 across the axial width or span of the vanes as measured at a location immediately downstream of the trailing edges 52. The same measurements for prior art, n~n-contoured endless are shown in dashed lines. In the central, mid-stream section it is clear that the contoured endless slightly reduce ~2~1Z75 efficiency and increase losses. however, the integrated average loss across the entire span shows a significant efficiency increase and reduction in aerodynamic loss for the entire nozzle section.
The combined reduction in secondary flow migration and reduction in boundary flow losses more than offset the controlled and limited introduction ox cross-passage velocity variation caused by the end wall contouring. The effect therefore is a net increase in nozzle section aerodynamic efficiency by producing a relit;
uniform total aerodynamic loss across the axial width of each-nozzle.
The foregoing detailed description of a preferred embodiment of the present invention should be considered exemplary and not as limiting to the scope and spirit of the invention as set forth in the appended claims.
I waving described the invention with sufficient clarity that those skilled in the art may make and practice it, we claim:

Claims (22)

THE EMBODIMENTS OF THE INVENTION IN WHICH AN EXCLUSIVE
PROPERTY OR PRIVILEGE IS CLAIMED ARE DEFINED AS FOLLOWS:

1. A radial inflow turbine rotatable about an axis and having a radial inlet for fluid flow; and a pair of axially spaced, generally parallel, radially extending endwalls disposed outwardly of and in surrounding relation-ship to said inlet to define a passage between said endwalls for directing fluid flow into said inlet, said endwalls being complementary contoured whereby said passage has radially inner and outer sections, wherein the axial width of said inner section is substantially less than the axial width of said outer section, and a variable width transition section between said inner and outer sections.

2. A radial inflow turbine as set forth in Claim 1, wherein said axial width of said inner section is less than 85 percent of the axial width of said outer section.

3. A radial inflow turbine as set forth in Claim 2, wherein said axial width of said inner section is between approximately 50 and 85 percent of the axial width of said outer section.

4. A radial inflow turbine as set forth in Claim 3, wherein said axial width of said inner section is approximately 75 percent of the axial width of said outer section.

5. A radial inflow turbine as set forth in Claim 1, further including a plurality of circumferentially spaced stator vanes extending axially across said passage.

6. A radial inflow turbine as set forth in Claim 5, wherein said vanes extend radially cross said transition section and into both said inner and outer sections.

7. In turbomachinery: a housing; a radial inflow turbine mounted for rotation in said housing about an axis in response to motive fluid flow into a radially outer, circumferential inlet to said turbine; and a pair of axially spaced, endwalls extending radially outwardly from said inlet substantially perpendicular to said axis to define an annular fluid flow passage for carrying motive fluid flow to said inlet, said endwalls being complementally, smoothly contoured to define radially inner and outer sec-tions of substantially constant, axial widths, the axial width of said outer section being greater than the axial width of said inner section to minimize secondary fluid flow adjacent both said endwalls and provide relatively uniform flow conditions across substantially the entire axial width of said inner section at a location adjacent said inlet.

8. In a subsonic radial inflow turbine having a circumferentially extending radial flow inlet: a stator ring surrounding said inlet, said ring comprising a pair of axially spaced, radial endwalls defining therebetween an inlet flow passage to said inlet, said endwalls being complementally configured to present radially outer and inner sections of said passage of respectively larger and smaller, substantially constant axial widths, and an inter-mediately disposed compound curvature transition section of said passage having an axial width smoothly varying from said larger to said smaller axial width, said ring further comprising a plurality of circumferentially spaced stator vanes extending axially across said passage between said endwalls, said vanes extending radially across said transition section into said inner and outer sections of said passage.

9. An annular nozzle ring for delivering sub-sonic, compressible, motive fluid flow to a radial inflow turbine, comprising a pair of annular, axially spaced, radially extending endwalls complementally contoured to axially converge in a radially inward direction, and a plurality of circumferentially spaced stator vanes axially spanning the space between said endwalls and complementally contoured therewith to define inlet nozzles therebetween, said complemental contouring of said endwalls and stator vanes being for substantially reducing secondary flow circumferentially across said nozzles between adjacent stator vanes in a manner producing relatively uniform total pressure loss axially across substantially the en-tire axial span of each of said nozzles at a location immediately downstream of the trailing edges of the associated stator vanes.

10. An annular flow inlet nozzle ring for delivering subsonic, compressible, motive fluid flow to a radial inflow turbine, comprising a pair of annular, axially spaced, radially extending endwalls complementally contoured to converge toward one another in a radially in-ward direction; and a plurality of circumferentially spaced stator vanes axially spanning said endwalls and complementally contoured therewith to define inlet nozzles therebetween, said complemental contouring of said endwalls and said stator vanes including compound curvature and being such as to minimize secondary fluid flow between adjacent stator vanes and produce a relatively uniform total pressure loss across substantially the entire axial span of each of said nozzles at locations immediately down-stream of the radially inner ends of said stator vanes.

11. A nozzle ring as set forth in Claim 10, wherein said endwalls and stator vanes are contoured whereby the axial span of each of said stator vanes at the radially inner trailing edge thereof is about three-fourths of the axial span at the radially outer leading edge there-of.

12. A nozzle ring as set forth in 11, wherein said stator vanes are curved circumferentially for smoothly turning inlet flow to said turbine and presenting pressure and suction surfaces defining boundaries of said nozzles, said contouring of said vanes being operable to minimize migration of secondary flow.

13. A gas turbine engine including a housing;
a fluid compressor mounted for rotation in said housing;
a combustor for receiving and heating fluid flow from said compressor; a radial inflow turbine mounted for rotation in said housing about an axis in response to motive fluid flow received from said combustor into a radially outer, circumferential inlet; a pair of endwalls extending radially outwardly from said inlet substantially perpendi-cular to said axis to define an annular fluid flow passage for carrying motive fluid flow to said inlet, said endwalls being complementally, smoothly contoured to define radially inner and outer sections of substantially constant, axial widths, the axial width of said outer section being greater than the axial width of said inner section to minimize secondary fluid flow adjacent both said endwalls and pro-vide relatively uniform flow conditions across substantially the entire axial width of said inner section at a location adjacent said inlet.

14. A radial inflow turbine rotatable about an axis and having a radial inlet for fluid flow; and a pair of axially spaced, generally parallel, radially extending endwalls disposed outwardly of and in surrounding relation-ship to said inlet to define a passage between said endwalls for directing fluid flow into said inlet, said endwalls being complementally contoured whereby said passage has radially inner and outer sections of substantially constant, different axial widths and a variable width transition sec-tion between said inner and outer sections; said comple-mental contouring of said endwalls and stator vanes being for substantially reducing secondary flow circumferentially across inlet nozzles between adjacent stator vanes in a manner producing relatively uniform total pressure loss axially across substantially the entire axial span of each of said nozzles at a location immediately downstream of the trailing edges of the associated stator vanes.

15. A radial inflow turbine as set forth in any of Claims 1, 8 and 14, wherein said endwall complemental contouring includes an outer concave section and an inner convex section in relation to the fluid flowing toward said turbine.

16. Turbomachinery as set forth in Claim 7, wherein said endwall complemental contouring includes an outer concave section and an inner convex section in rela-tion to the fluid flowing toward said turbine.

17. An annular nozzle ring as set forth in either one of Claims 9 and 10, wherein said endwall comp-lemental contouring includes an outer concave section and an inner convex section in relation to the fluid flowing toward said turbine.

18. A gas turbine engine as set forth in Claim 13, wherein said endwall complemental contouring includes an outer concave section and an inner convex section in relation to the fluid flowing toward said turbine.

19. A method for smoothly turning and accelerat-ing fluid flow immediately prior to its delivery to a radial inflow, rotating turbine, comprising the steps of:
delivering subsonic, compressible fluid flow radially inwardly to a plurality of radial inlet nozzles to said turbine separated by a plurality of circumferenti-ally curved stator vanes;
imparting angular acceleration to said fluid flow by momentum transfer to said fluid flow by its interaction with the pressure sides and suction sides of said circum-ferentially curved stator vanes;
minimizing radial acceleration of said fluid flow in outer portions of said nozzles immediately downstream of the leading edges of said stator vanes at both said pres-sure and suction sides thereof; and relatively smoothly radially accelerating said fluid flow downstream of said outer portions yet while said fluid flow is still in momentum exchange relationship with said stator vanes.

20. A method as set forth in Claim 19, wherein said minimizing radial acceleration step and said radially accelerating step are together operable to produce rela-tively uniform total aerodynamic loss across the width of each nozzle at a location immediately downstream of the trailing edges of said vanes.

21. A method as set forth in Claim 20, wherein said minimizing radial acceleration step and said radially accelerating step are together operable to minimize second-ary fluid flow across each nozzle from the pressure side of one stator vane toward the suction side of the adjacent stator vanes.

22. A method as set forth in Claims 19,20 or 21, wherein axially spaced endwalls defining boundaries of said nozzles are contoured with compound curvature.