U.S. patent application number 12/792352 was filed with the patent office on 2010-12-02 for turbine stage.
This patent application is currently assigned to ALSTOM Technology Ltd. Invention is credited to Said HAVAKECHIAN, Benjamin MEGERLE, Thomas MOKULYS.
Application Number | 20100303627 12/792352 |
Document ID | / |
Family ID | 41138647 |
Filed Date | 2010-12-02 |
United States Patent
Application |
20100303627 |
Kind Code |
A1 |
MEGERLE; Benjamin ; et
al. |
December 2, 2010 |
TURBINE STAGE
Abstract
A turbine stage includes a circumferentially distributed row of
adjacent airfoils between which there is a flow passage. The
passage has a surface, which in its unmodified form, defines a
datum plane. A channel, located in the passage, extends in the
direction of an airfoil pressure face from a point towards a
leading edge line to a point towards a trailing edge line. The
channel includes two channel walls angled relative to the datum
plane. Relative to the datum plane, the channel has a low point,
two high points, and a channel height, which is measured between
the low point and the highest one of the high points. The channel
provides a means to reduce secondary flow losses.
Inventors: |
MEGERLE; Benjamin;
(Wettingen, CH) ; MOKULYS; Thomas; (Wuerenlingen,
CH) ; HAVAKECHIAN; Said; (Baden, CH) |
Correspondence
Address: |
BUCHANAN, INGERSOLL & ROONEY PC
POST OFFICE BOX 1404
ALEXANDRIA
VA
22313-1404
US
|
Assignee: |
ALSTOM Technology Ltd
Baden
CH
|
Family ID: |
41138647 |
Appl. No.: |
12/792352 |
Filed: |
June 2, 2010 |
Current U.S.
Class: |
416/179 ;
415/221 |
Current CPC
Class: |
F01D 5/145 20130101;
F01D 5/143 20130101; F05D 2250/611 20130101; F05D 2250/71
20130101 |
Class at
Publication: |
416/179 ;
415/221 |
International
Class: |
F01D 5/22 20060101
F01D005/22; F04D 29/54 20060101 F04D029/54 |
Foreign Application Data
Date |
Code |
Application Number |
Jun 2, 2009 |
EP |
09161706.8 |
Claims
1. A turbine stage comprising: a circumferentially distributed row
of adjacent airfoils, each of the airfoils respectively having a
pressure face, a suction face, and at least one end wall from which
the airfoils radially extend, respectively; and a flow passage
defined by a region between a pressure face of a first one of the
airfoils, a suction face of a second one of the airfoils adjacent
to the first airfoil, a leading edge line extending between lead
edges of adjacent airfoils, and a trailing edge line extending
between trailing edges of adjacent airfoils, wherein the flow
passage has a defined datum plane extending between a base of the
pressure face and a base of the suction face of adjacent airfoils,
and wherein the turbine stage further comprises a channel, in the
flow passage, adjacent to the pressure face and extending in the
direction of the pressure face from a point towards the leading
edge line to a point towards the trailing edge line, the channel
including two angled and joined channels walls that, relative to
the datum plane, form a low point at the joined channel walls, two
high points, and a channel height, the channel height defining the
radial distance between the low point and highest one of the high
points.
2. The turbine stage of claim 1, wherein each flow passage includes
at least two adjacent channels.
3. The turbine stage of claim 1, wherein the high points of the
channel do not extend above the datum plane.
4. The turbine stage of claim 1, wherein the low point of the
channel in each flow passage is substantially in the midpoint, in
the pitchwise direction, between the high points of each
channel.
5. The turbine stage of claim 1, wherein in each flow passage, an
angle of the channel wall closer to the pressure face is less,
relative to the datum plane, than an angle of the channel wall
closer to the suction face.
6. The turbine stage of claim 1, wherein in each flow passage, the
channel height in the primary flow direction increases to a
maximum, at a relative channel length measured in the direction
from the leading edge line to the trailing edge line, of between
approximately 0.35-0.55, at which point the channel height
decreases.
7. The turbine stage of claim 2, wherein in the pitchwise direction
from the pressure face to the suction face, the channel height of
each successive adjacent channel one of remains the same and
decreases.
8. The turbine stage of claim 7, wherein, in each flow passage, the
channel height of the channel adjacent the pressure face is at
least twice that of the channel furthest, in the pitchwise
direction, from the pressure face.
9. The turbine stage of claim 2, wherein, in each flow passage, in
the pitchwise direction from the pressure face to the suction face,
each successive adjacent channel extends from a point that is one
of the same and further from the leading edge line to form an
extended region, adjacent the suction face, free of channels.
10. The turbine stage of claim 9, wherein the extended region is a
region that, in operation, encompasses a region essentially free of
secondary flow vortices caused by cross flow originating from the
pressure face.
11. The turbine stage of claim 2, wherein the high points of the
channel do not extend above the datum plane.
12. The turbine stage of claim 11, wherein in the pitchwise
direction from the pressure face to the suction face, the channel
height of each successive adjacent channel one of remains the same
and decreases.
13. The turbine stage of claim 12, wherein, in each flow passage,
the channel height of the channel adjacent the pressure face is at
least twice that of the channel furthest, in the pitchwise
direction, from the pressure face.
14. The turbine stage of claim 11, wherein, in each flow passage,
in the pitchwise direction from the pressure face to the suction
face, each successive adjacent channel extends from a point that is
one of the same and further from the leading edge line to form an
extended region, adjacent the suction face, free of channels.
15. The turbine stage of claim 14, wherein the extended region is a
region that, in operation, encompasses a region essentially free of
secondary flow vortices caused by cross flow originating from the
pressure face.
16. The turbine stage of claim 2, wherein the low point of the
channel in each flow passage is substantially in the midpoint, in
the pitchwise direction, between the high points of each
channel.
17. The turbine stage of claim 2, wherein in each flow passage, an
angle of the channel wall closer to the pressure face is less,
relative to the datum plane, than an angle of the channel wall
closer to the suction face.
18. The turbine stage of claim 2, wherein in each flow passage, the
channel height in the primary flow direction increases to a
maximum, at a relative channel length measured in the direction
from the leading edge line to the trailing edge line, of between
approximately 0.35-0.55, at which point the channel height
decreases.
19. The turbine stage of claim 3, wherein the low point of the
channel in each flow passage is substantially in the midpoint, in
the pitchwise direction, between the high points of each
channel.
20. The turbine stage of claim 3, wherein in each flow passage, an
angle of the channel wall closer to the pressure face is less,
relative to the datum plane, than an angle of the channel wall
closer to the suction face.
21. The turbine stage of claim 3, wherein in each flow passage, the
channel height in the primary flow direction increases to a
maximum, at a relative channel length measured in the direction
from the leading edge line to the trailing edge line, of between
approximately 0.35-0.55, at which point the channel height
decreases.
22. The turbine stage of claim 4, wherein in each flow passage, an
angle of the channel wall closer to the pressure face is less,
relative to the datum plane, than an angle of the channel wall
closer to the suction face.
23. The turbine stage of claim 4, wherein in each flow passage, the
channel height in the primary flow direction increases to a
maximum, at a relative channel length measured in the direction
from the leading edge line to the trailing edge line, of between
approximately 0.35-0.55, at which point the channel height
decreases.
24. The turbine stage of claim 1, wherein each airfoil respectively
has two end walls between which the airfoils respectively extend.
Description
RELATED APPLICATION
[0001] This application claims priority under 35 U.S.C. .sctn.119
to European Patent Application No. 09161706.8 filed in Europe on
Jun. 2, 2009, the entire content of which is hereby incorporated by
reference in its entirety.
FIELD
[0002] The present disclosure relates generally to axial gas and
steam turbines in which there are one or more rows of generally
radially extending airfoils of non rotating vanes and rotating
blades. More particularly, the present disclosure relates to
configurations of end walls joined to radial ends of airfoils with
improved aerodynamic behaviour.
[0003] Within this specification, the term "pitchwise" is used to
mean a circumferential direction between two adjacent airfoils,
vanes or blades. Further, the term "end wall" is broadly defined to
encompass any surface at a radial end of an airfoil, and any
surface from which the airfoil radially extends. End walls thus
include, but are not limited to, airfoil platforms and shrouds.
BACKGROUND INFORMATION
[0004] The ideal flow through a turbine is termed the "primary
flow". The difference between the primary flow and the actual flow
is termed the "secondary flow". The secondary flow represents, to a
large extent, a loss that has an impact on axial turbine
efficiency.
[0005] The development of secondary flow in a turbine cascade
starts with the end wall boundary layer interacting with a leading
edge of the airfoil. As the flow impinges on the leading edge of
the airfoil, radial variation in the stagnation pressure creates
flow along the stagnation line of the airfoil towards the end wall.
When this flow reaches the end wall, the flow travels locally
upstream along the end wall. A separation occurs where the incoming
boundary layer meets this flow, and a so-called horseshoe vortex is
formed around the leading edge of the airfoil. The strength of this
vortex is dependent on the thickness of the leading edge and the
variation of the radial static pressure gradient along the leading
edge, which is, among other things, linked to the end wall boundary
layer thickness and quality.
[0006] The pressure side leg of the vortex is influenced by the
airfoil-to-airfoil pressure gradient as it enters the flow passage
and travels towards the suction side. The resulting cross-passage
flow along the end wall imposes vortex motion in the cascade. These
vortices may commonly be referred to as passage vortices that
include horseshoe vortices in their core. These vortices can be
present in any flow channel with a curved shape and a boundary
layer. The strength of this secondary flow in a cascade is
dependent on a number of other factors, including the amount of
turning and the shape of the incoming boundary layer.
[0007] The end wall vortex generation negatively affects turbine
efficiency, contributing up to approximately 35% of the total
losses for a typical high-pressure turbine. The key cause for the
additional loss generation is the passage vortex that grows
downstream of the cascade. The kinetic energy stored in this vortex
is lost for further use as the kinetic energy is mostly mixed out
downstream. The passage vortex can be easily detected as a high
loss core existing away from the suction surface close to the
center of the passage vortex.
[0008] Besides loss generation, secondary flow perturbs the exit
flow distribution downstream of the cascade. Since the low momentum
boundary layer fluid is deflected substantially more than the main
flow close to the end wall, the low momentum boundary layer fluid
sees the same blade to blade pressure gradient but has less impulse
and thus causes overturning of the exit flow close to the end wall.
Further away from the end wall, the rotation of the passage vortex
comes into play and thus less turning occurs, which, due to the
passage vortex driving the fluid in an opposite direction, results
in so-called under turning.
[0009] The inhomogeneous flow field after the cascade is
responsible for additional losses in the following cascade. This is
partly due to the overturned flow close to the end wall leading to
more secondary flow in the next blade row.
[0010] As the passage vortex lifts off the end wall and grows in
size, the flow channel is increasingly influenced by secondary
flow. It is known to be beneficial if the passage vortex is closer
to the end wall as this increases the region of undisturbed primary
flow. One method of inferring this is through the measurement of
peak radial helicity.
[0011] Airfoil design has evolved to reduce secondary flow by
optimizing the three-dimensional shape of airfoils and, more
recently, by contouring of the end walls. This technology, which is
referred to as Tangential End Wall Contouring (TEWC), involves
adjusting end wall surfaces to reduce secondary flow, resulting,
for example, in a modified airfoil face pressure profile.
[0012] US Patent Application Publication No. US 2007/0059177 A1
describes such an end wall non-axisymmetric profile. This solution
includes forming circumferentially extending sinusoids at a number
of axial positions, wherein corresponding points on successive
sinusoids are joined by spline curves so that the curvature of the
end wall is smooth.
[0013] Another alternative solution involves providing the end wall
with a fence that lifts the vortex up off the end wall and into the
main flow, which has the effect of washing out the vortices. The
fence, of which an example is described in ASME Turbo Expo 2000
"Secondary flow measurements in a turbine passage with end wall
flow modification", 2000-GT-0212, has a leading edge at the center
of a line connecting the leading edges of the pressure and suction
side airfoils. While such walls can reduce aerodynamic losses,
practical problems can arise due to the need to cool the fence.
[0014] An alternative method of reducing the affects of secondary
flow involves using non-axisymmetric profiling to reduce cross flow
instead of adjusting the pressure profile. EP 1 995 410 A1, for
example, provides a solution in which an end wall of a turbine
stage cascade includes a first projection having a ridge extending
downward from the trailing edge of a turbine blade toward the
downstream side, gently at the beginning and steeply at the end,
and along the suction side of an adjacent turbine blade. Such an
arrangement is, however, limited by the fact that downstream axial
space is required, and therefore, such a solution may not always be
applicable.
SUMMARY
[0015] An exemplary embodiment provides a turbine stage, which
includes a circumferentially distributed row of adjacent airfoils.
Each of the airfoils respectively has a pressure face, a suction
face, and at least one end wall from which the airfoils radially
extend, respectively. The exemplary turbine stage also includes a
flow passage, which is defined by a region between a pressure face
of a first one of the airfoils, a suction face of a second one of
the airfoils adjacent to the first airfoil, a leading edge line
extending between lead edges of adjacent airfoils, and a trailing
edge line extending between trailing edges of adjacent airfoils.
The flow passage has a defined datum plane extending between a base
of the pressure face and a base of the suction face of adjacent
airfoils. The exemplary turbine stage also includes a channel, in
the flow passage, adjacent to the pressure face and extending in
the direction of the pressure face from a point towards the leading
edge line to a point towards the trailing edge line, the channel
including two angled and joined channels walls that, relative to
the datum plane, form a low point at the joined channel walls, two
high points, and a channel height, the channel height defining the
radial distance between the low point and highest one of the high
points.
BRIEF DESCRIPTION OF THE DRAWINGS
[0016] Additional refinements, advantages and features of the
present disclosure are described in more detail below with
reference to exemplary embodiments illustrated in the drawings, in
which:
[0017] FIG. 1 is a top view of two adjacent airfoils of a turbine
stage according to an exemplary embodiment of the present
disclosure;
[0018] FIG. 2 is a perspective view of the adjacent airfoils of
FIG. 1;
[0019] FIG. 3 is a perspective view of the two adjacent airfoils of
a turbine stage with exemplary channels of the disclosure in end
walls according to an exemplary embodiment of the disclosure;
[0020] FIG. 4 is a pitchwise sectional view of a turbine stage
showing adjacent airfoils and an end wall of an exemplary
embodiment of the present disclosure;
[0021] FIGS. 5 and 6 are expanded views of channel wall sections V
and VI of FIG. 4, respectively;
[0022] FIG. 7 is a pitchwise profile of a channel of FIG. 3 or
4;
[0023] FIG. 8 is a height profile of a channel of FIG. 3 or 4;
[0024] FIG. 9 is a perspective view of an exemplary embodiment with
an extended region;
[0025] FIG. 10 is a top view of FIG. 9 showing secondary flow
lines; and
[0026] FIGS. 11-14 are exemplary performance graphs of exemplary
embodiments.
[0027] Other aspects and advantages will become apparent from the
following description, when taken in connection with the
accompanying drawings, which, by way of illustration and example,
illustrate exemplary embodiments of the present disclosure.
DETAILED DESCRIPTION
[0028] Exemplary embodiments of the present disclosure are directed
towards correcting the problem, in a turbine, of over- and
under-turning capability and/or reduced helicity losses as a result
of secondary flows caused by cross flow that flows in the pitchwise
direction from a pressure face of an airfoil towards the suction
face of an adjacent airfoil across an end wall surface.
[0029] Exemplary embodiments of the present disclosure address this
problem by providing a turbine stage as described herein.
[0030] According to an exemplary embodiment, the turbine stage
includes one or more adjacent channels formed in end walls in flow
passages between adjacent airfoils. Each channel extends in the
primary flow direction and can be located adjacent to airfoil
pressure faces, for example. The channels each have two angled
walls, which, in conjunction with the configuration and location of
the channels, are configured to reduce the potential for secondary
flow formation in the channels.
[0031] An exemplary embodiment provides a turbine stage that
includes a circumferentially distributed row of adjacent airfoils.
Each of the airfoils includes a pressure face, a suction face, and
at least one end wall, from which the airfoils radially extend,
respectively. One or more of the airfoils can also include two end
walls between which the airfoils extend. The turbine stage also
includes a flow passage defined by a region between a pressure face
of a first airfoil, a suction face of a second airfoil adjacent to
the first airfoil, a leading edge line, which is defined as a line
extending between the lead edges of adjacent airfoils, and a
trailing edge line, which is defined as a line extending between
the trailing edges of adjacent airfoils. The flow passage has a
surface, which, in its unmodified form, defines a datum plane. The
turbine stage, in each flow passage has at least one adjacent
channel(s), adjacent a pressure face, that modify the surface and
extend in the direction of primary flow lines from a point towards
the leading edge line to a point towards the trailing edge line.
Each channel includes two channels walls angled relative to the
datum plane that provide the channels with a low point, two high
points, and a channel height which is the radial distance between
the low point and a highest one of the high points. The location of
the channels adjacent the pressure face reduces the extent of
influence of cross flow pitchwise across the flow passage and
thereby reduce the influence of secondary flow. The closer the
channel is to the pressure face, the more pronounced this effect
is. In this way, any negative effect on aerodynamic performance is
more than offset by the benefit of reduced secondary cross
flow.
[0032] In accordance with an exemplary embodiment, the high points
of each channel do not extend above the datum plane, thereby
reducing the impact of the channels on primary flow and reducing
scraping losses. In accordance with an exemplary embodiment, the
low point of each channel is substantially in the midpoint,
pitchwise, between high points of each channel. In accordance with
another exemplary embodiment, the angle of the walls of each
channel closer to the pressure face is less, relative to the datum
plane, than channel walls closer to the suction face.
[0033] According to an exemplary embodiment, the channel height, in
the primary flow direction, increases to a maximum at a relative
channel length, in the direction of primary flow lines, of between
approximately 0.35-0.55, for example, at which point the channel
height decreases. In the last fifth of the length of the channel,
this rate of decrease may be less. In this way, the channel depth
provides a balance between scraping losses, which change with the
velocity profile in the flow passage, and cross flow presence.
[0034] In accordance with an exemplary embodiment, the channel
height of each successive adjacent channel, adjacent in the pitch
wise direction extending from the pressure face to the suction
face, remains the same or decreases. For example, the channel
height of the channel adjacent the pressure face is at least twice
that of the channel furthest, in the pitch wise direction, from the
pressure face. In this way, channels are configured to reduce cross
flow where its affect is strongest, e.g., adjacent the pressure
face.
[0035] In accordance with an exemplary embodiment, in each flow
passage, the point of extension of each adjacent channel is the
same or further from the leading edge line the closer the channel
is to the suction face thus defining an extended region, adjacent
the suction face, which is free of channels. The extended region
can be generally defined as a region, that in operation, is
configured in a region that is essentially free of secondary flow
vortices caused by cross flow originating from the pressure
face.
[0036] Exemplary embodiments of the present disclosure are now
described with reference to the drawings, wherein like reference
numerals are used to refer to like elements throughout. In the
following description, for purposes of explanation, numerous
specific details are set forth in order to provide a thorough
understanding of the disclosure. It may be evident, however, that
the disclosure may be practiced without these specific details.
[0037] FIGS. 1 and 2 respectively show top and perspective views of
two adjacent airfoils 10 of a turbine stage, in which airfoils 10
are adjacently and circumferentially distributed in rows. Each
airfoil 10 is integrally joined at one or both radial ends to
corresponding end walls 12, which are partially shown as grid
lines. The area between the pressure face 14 and suction face 16 of
adjacent airfoils 10 defines a flow passage 18 that is further
bound by a region extending between a leading edge line 20, which
is defined as a line extending between the lead edges 21 of
adjacent airfoils 10, and a trailing edge line 22, which is defined
as a line extending between the trailing edge 23 of adjacent
airfoils 10. The flow passage 18 has a surface common with a
surface of the end walls 12. In its unmodified form, the surface
defines a datum plane DR (see FIG. 4), wherein "unmodified form"
means the contour that the passage surface would take if the
surfaces were not changed, for example, by TEWCs, thus forming a
plane extending between bases of the pressure face 14 and suction
face 16 of adjacent airfoils 10. The grid lines in FIGS. 1 and 2
representing an unmodified surface include primary flow lines PFL
that represent ideal lines of flow unaffected by secondary flow,
and pitch wise sections A-D.
[0038] FIG. 3 shows an exemplary embodiment applied to the turbine
stage shown in FIGS. 1 and 2. The exemplary embodiment illustrated
in FIG. 3 includes channels 30 formed in the passage surfaces of
end walls 12, at one or both radial ends of adjacent airfoils 10.
The channels 30 extend in the direction of primary flow lines PFL
(e.g., essentially in the direction of the pressure face 14) of an
airfoil 10, resulting in the channels 30 being substantially
parallel to each other. The extension of the channels 30 is from a
point towards the leading edge line 20 to a point towards the
trailing edge line 22. Each channel 30 includes two channel walls
32 that angle relative to the datum plane DR, which is shown in
more detail in FIGS. 5 and 6, and join to define a low point LP of
the channel 30 relative to the datum plane DR.
[0039] FIG. 4 shows an exemplary cross section through a pitch wise
section A-D extending between the pressure face 14 of one airfoil
10 to the suction face 16 of another adjacent airfoil 10. As shown
in the exemplary embodiment illustrated in FIG. 4, channels 30 are
formed in end walls 12 between the adjacent airfoils 10. Each
channel 30 has two channel walls 32, one closer to the pressure
face 14 and the other closer to the suction face 16. The channel
height CH is the radial height, that is, the height measured
perpendicular to the datum plane DR, between the channel's low
points LP and the highest one of the high points HP. As used
herein, "low" and "high" are defined relative to the datum plane
DR, wherein "low" refers to a negative extension from the datum
plane DR into the end wall 12, while "high" refers to a positive
extension in the direction away from the end wall 12. The
indication is independent of absolute location. That is, even
though the high point HP extends in a direction away from the end
wall 12, the high point HP, as shown in FIG. 7, for example, may or
may not extend above the height of the datum plane DR
[0040] The "channel wall angle" .theta., as shown in the exemplary
embodiments illustrated in FIGS. 5 and 6, is the angle of a nominal
channel wall 33 relative to the datum plane DR. The nominal channel
wall 33, without curvature, approximates the actual channel wall
32. For example, with reference to the exemplary embodiment
illustrated in FIG. 5, an expanded view of V of FIG. 4 is shown, in
which the channel wall angle .theta. of a bowed channel wall 32 is
illustrated. The wall angle .theta. is taken to be the angle of the
nominal channel wall 33, which is the average angle of the nominal
channel wall 32. In another example shown in FIG. 6, which is an
expanded view of VI of FIG. 4, a channel wall 32 is shown with a
rounded-off end section that is otherwise straight. In this case,
the nominal channel wall 33 corresponds to the channel wall 32
straight portion, disregarding the rounded end section.
[0041] One purpose of the channels 30 is to reduce cross flow and
thereby reduce secondary flow and resulting losses. The desired
channel height CH and desired number of channels 30 is dependent on
the degree of cross flow, estimatable using known techniques,
described, for example, in Harvey, N. W. et al, 2000
"Nonaxisymmetric Turbine End Wall Design: Part I", ASME J.
Turbomach., 122, pp. 278-285, and Hartland, J. C. et al, 2000
"Nonaxisymmetric Turbine End Wall Design: Part II", ASME 122 J.
Turbomach, 122, pp. 286-293. With increasing channel height CH and
number of channels 30 (and hence an increased number of channel
walls 32), passage surface area increases, which, in the absences
of secondary flow, results in increased scraping losses. Where the
effect of scraping losses may be higher than the beneficial effect
of channels 30, it may be advantageous to minimise both channel
height CH and/or the number of channels 30. An embodiment in its
simplest form suitable for turbine stages with minimal cross flow
therefore comprises one channel 30 located adjacent to the pressure
surface 14, which is the region with the most significant cross
flow.
[0042] The channel depth CH, shown in detail in FIG. 4, is a
function of the number of channels 30 and the degree of cross flow.
If the channel depth CH is too great, further secondary flow can be
created resulting in additional losses. If the channel depth CH is
too low, the ability of the channel 30 to limit cross flow will be
limited. A further consideration is the channel wall angle .theta..
If the channel wall angle .theta. is too steep, additional
secondary flow may be created. Channel design is therefore a
compromise between at least these factors and so is dependent on
airfoil design and operation conditions. In consideration of these
factors, an optimum design can be derived by simulation using known
methods.
[0043] In accordance with an exemplary embodiment of the present
disclosure, the low point LP of each channel 30 is at the pitchwise
midpoint between the high points HP of the channel, as shown in
FIG. 4, for example.
[0044] In accordance with an exemplary embodiment of the present
disclosure, the channel height CH of each successive adjacent
channel 30 in the pitchwise direction from the pressure face 14 to
the suction face 16 remains the same or decreases. In accordance
with another exemplary embodiment, the channel height CH of the
channel 30 adjacent the pressure face 14 is at least twice that of
the channel 30 furthest from the pressure face 14, as can be seen
in FIG. 4, for example. As cross flow is typically greatest towards
the pressure face 14, the benefit of channels 30 may decrease
towards the suction face 16.
[0045] In accordance with another exemplary embodiment of the
present disclosure, the low point LP is closer to the suction face
16, represented as the pitchwise position "1" in FIG. 7 than the
pressure face 14, which is represented as "0". This arrangement can
result in the channel wall angle .theta. of the channel wall 32
closer the suction face 16 being greater than the channel wall
angle .theta. of the channel wall 32 closer the pressure face 14.
In this way, a smooth transition into the channel 30 is provided
for cross flow originating from the pressure face 14, which
minimizes the formation of additional losses, while cross flow
suppression can be promoted by the steeper channel wall angle
.theta. of the channel wall 32 located closer to the suction face
16. The channel wall angle .theta. of the channel wall 32 located
closer to the suction face 16 can be less than less than 90
degrees, since an angle approaching 90 degrees or greater may
create additional vortices resulting in additional losses.
[0046] In accordance with an exemplary embodiment of the present
disclosure, the channel walls 32 are configured such that the
channels 30 do not extend above the datum plane DR, as shown in
FIG. 7, for example, wherein "0" is the channel height CH at the
datum plane DR. By this means, it was found that scraping losses of
the primary flow can be further reduced while still maintaining
good cross flow suppression performance.
[0047] Through a turbine cascade, the flow is accelerated
significantly. Scraping losses, which have a squared relationship
with velocity, are of greatest significance in the region of
highest velocity. The highest velocity may correspond to a region
where the separation distance measured in the pitchwise direction,
between adjacent airfoils 10, is smallest. In such a region,
overall efficiency may be optimized if the channel height CH is
limited so as to be lower than would optimally be designed in view
only of predicted cross flow. Therefore, in accordance with an
exemplary embodiment of the present disclosure, in the direction of
primary flow lines PFL extending from towards the leading edge line
20 to the trailing edge line 20, the channel height initially
increases to a maximum at a relative channel length of between
approximately 0.35-0.55, for example, after which it decreases. In
accordance with another exemplary embodiment of the present
disclosure, the decrease is not as pronounced in the last fifth of
the relative channel length. The relative channel length is the
length point along a channel 30 measured relative to the total
length of the channel 30. FIG. 8 shows an example of a
configuration of an exemplary embodiment in which it was found that
for one set of operating conditions not only can scraping losses be
reduced but also over- and under-turning performance can be
improved without detrimentally affecting helicity.
[0048] FIG. 9 shows an exemplary embodiment in which the channels
30 towards the suction face 16 start further from the leading edge
line 20 than the channels 30 closer to the pressure face 14. That
is, their point of extension from the leading edge line 20 is
further. This results in the formation of an extended region ER
adjacent the suction face 16, towards the leading edge line 20,
which is free of channels 30. The extended region ER may be bounded
by a midpoint on the leading edge line 20, a point along the
suction face 16, and/or a point on the suction face 16 at which the
suction face 16 and leading edge line 20 join, as shown in both
FIGS. 9 and 10. Such an arrangement is beneficial when the flow
across the extended region ER is essentially free of secondary
flow, as shown in FIG. 10, and as such the loss in this region
primarily includes scraping losses. In accordance with another
exemplary embodiment of the present disclosure, the extended region
ER is the region adjacent the suction face 16 towards the leading
edge line 20 that is essentially free of secondary flow, as shown
by the flow lines FL in FIG. 10. As the size and shape of the
extended region ER is dependent not only on turbine stage
configuration but also the operating conditions of the turbine
stage, the optimum location of the extended region ER is unique for
each turbine configuration. Accordingly, the extended region ER is
defined by a region derived and determined by known flow simulation
methods.
[0049] FIGS. 11-14 show examples of the performance that can be
achieved with a combination of various exemplary embodiments
described herein. Improvements include over- and under-turning,
shown in FIGS. 11 and 12 for both a stator and a rotor, and
helicity, shown in FIGS. 13 and 14 also for a stator and rotor.
[0050] Although the disclosure has been herein shown and described
in what is conceived to be practical exemplary embodiments, it will
be appreciated by those skilled in the art that the present
invention can be embodied in other specific forms without departing
from the spirit or essential characteristics thereof. The presently
disclosed embodiments are therefore considered in all respects to
be illustrative and not restricted. The scope of the invention is
indicated by the appended claims rather that the foregoing
description and all changes that come within the meaning and range
and equivalences thereof are intended to be embraced therein.
REFERENCE NUMBERS
[0051] 10 Airfoil [0052] 12 End wall [0053] 14 Pressure face [0054]
16 Suction face [0055] 18 Flow passage [0056] 20 Leading edge line
[0057] 21 Leading edge [0058] 22 Trailing edge line [0059] 23
Trailing edge [0060] 30 Channel [0061] 32 Channel wall [0062] 33
Nominal channel wall [0063] A-D Pitchwise sections [0064] CH
Channel height [0065] DR Datum plane [0066] ER Extended region
[0067] FL Flow lines [0068] PFL Primary flow lines [0069] LP Low
point (of a channel) [0070] HP High point (of a channel) [0071] RD
Radial direction [0072] .theta. Channel wall angle
* * * * *