U.S. patent application number 13/687789 was filed with the patent office on 2013-06-20 for turbomachine and turbomachine stage.
This patent application is currently assigned to MTU Aero Engines GmbH. The applicant listed for this patent is MTU Aero Engines GmbH. Invention is credited to Michael ENGELKE, Jochen GIER, Kai KOERBER, Inga MAHLE.
Application Number | 20130156562 13/687789 |
Document ID | / |
Family ID | 45442918 |
Filed Date | 2013-06-20 |
United States Patent
Application |
20130156562 |
Kind Code |
A1 |
MAHLE; Inga ; et
al. |
June 20, 2013 |
TURBOMACHINE AND TURBOMACHINE STAGE
Abstract
A turbomachine stage including guide vanes, radially inner
and/or radially outer airfoil platforms, which together form a
guide vane cascade, and further including rotor blades, radially
inner and/or radially outer airfoil platforms, which together form
a rotor blade cascade adjacent to the guide vane cascade. Radially
outer airfoil platforms have cascade regions extending between
circumferentially adjacent airfoils, and gap regions which radially
and/or axially bound an axial gap extending axially between the
guide vane cascade and the rotor blade cascade. A contour of at
least one of these gap regions varies, in particular periodically,
in the radial and/or axial direction around the circumference.
Inventors: |
MAHLE; Inga; (Muenchen,
DE) ; GIER; Jochen; (Karlsfeld, DE) ; KOERBER;
Kai; (Karlsfeld, DE) ; ENGELKE; Michael;
(Bremen, DE) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
MTU Aero Engines GmbH; |
Muenchen |
|
DE |
|
|
Assignee: |
MTU Aero Engines GmbH
Muenchen
DE
|
Family ID: |
45442918 |
Appl. No.: |
13/687789 |
Filed: |
November 28, 2012 |
Current U.S.
Class: |
415/191 |
Current CPC
Class: |
F01D 9/00 20130101; Y02T
50/673 20130101; F05D 2250/611 20130101; Y02T 50/60 20130101; F05D
2250/184 20130101; F01D 5/143 20130101 |
Class at
Publication: |
415/191 |
International
Class: |
F01D 9/00 20060101
F01D009/00 |
Foreign Application Data
Date |
Code |
Application Number |
Dec 20, 2011 |
EP |
EP11194433.6 |
Claims
1. A turbomachine stage comprising: guide vanes; a radially outer
guide vane airfoil platform defining with the guide vanes a guide
vane cascade, the radially outer guide vane airfoil platform having
a guide vane cascade region and a guide vane gap region extending
axially beyond the guide vane cascade region; rotor blades; a
radially outer rotor vane airfoil platform defining with the rotor
blades a rotor blade cascade, the radially outer rotor vane airfoil
platform having a rotor vane cascade region and a rotor vane gap
region extending axially beyond the rotor vane cascade region; the
rotor blade cascade being adjacent to the guide vane cascade, the
guide vane gap region and the rotor vane gap region at least one of
radially and axially defining an axial gap extending axially
between the guide vane cascade and the rotor blade cascade; a
contour of at least one of the guide vane gap region and the rotor
vane gap region varying in at least one of the radial and the axial
direction around a circumference, and an at least one of axially
and radially opposite contour of the other of the guide vane gap
region and the rotor vane gap region varying around the
circumference.
2. The turbomachine stage as recited in claim 1 wherein the contour
and the opposite contour are radially opposite.
3. The turbomachine stage as recited in claim 2 wherein the contour
and the opposite contour vary identically.
4. The turbomachine stage as recited in claim 1 wherein the contour
varies both radially and axially.
5. The turbomachine as recited in claim 1 wherein the contour
varially radially but is constant in the axial direction.
6. The turbomachine stage as recited in claim 1 wherein at least
one of the guide vane or rotor vane gap region merges smoothly into
the respective guide vane or rotor vane cascade region.
7. The turbomachine stage as recited in claim 1 wherein the contour
varies radially and an extreme extent of the radially varying
contour is circumferentially located in a pressure-side half of a
segment between two respective adjacent airfoil leading edges or in
a suction-side half of a segment between two respective adjacent
airfoil trailing edges.
8. The turbomachine stage as recited in claim 1 wherein the contour
varies radially and a maximum variation in the radial direction of
the respectve guide vane or rotor vane gap region is no more than
50% of the pitch of the respective guide vane or rotor vane
cascade.
9. The turbomachine as recited in claim 8 wherein the maximum
variation is no more than 40%.
10. The turbomachine stage as recited in claim 1 wherein the
contour varies axially and wherein an extreme extent of the axially
varying contour is circumferentially located in the region of a
respective airfoil edge.
11. The turbomachine as recited in claim 10 wherein the extreme
extent is circumferentially spaced from the respective airfoil edge
by no more than 25% of the cascade pitch.
12. The turbomachine stage as recited in claim 1 wherein the
contour varies axially and a maximum variation in the axial
direction is no more than 50% the pitch of the respective guide
vane or rotor vane cascade.
13. The turbomachine stage as recited in claim 1 wherein the
contour varies periodically.
14. The turbomachine stage as recited in claim 13 wherein the
opposite contour varies identically with the contour.
15. The turbomachine stage as recited in claim 13 wherein the
opposite contour varies identically with the contour.
16. A turbomachine comprising at least one turbomachine stage as
recited in claim 1.
17. A gas turbine comprising at least one turbomachine stage as
recited in claim 1.
18. An aircraft engine gas turbine comprising at least one
turbomachine stage as recited in claim 1.
19. A compressor stage comprising the turbomachine stage as recited
in claim 1.
20. A turbine stage comprising the turbomachine stage as recited in
claim 1.
Description
[0001] This claims the benefit of European Patent Application EP
11194433.6, filed Dec. 20, 2011 and hereby incorporated by
reference herein.
[0002] The present invention relates to a turbomachine, in
particular a gas turbine, preferably an aircraft engine gas
turbine, having at least one turbomachine stage, in particular a
compressor stage or a turbine stage, including a guide vane cascade
and a rotor blade cascade, and to such a turbomachine stage.
BACKGROUND
[0003] A turbomachine stage has a cascade of rotating rotor blades
and a cascade of guide vanes disposed adjacent to the rotor blade
cascade on the upstream or downstream side. The rotor blades
terminate in a radially outer airfoil platform at the root end.
Similarly, guide vanes may be provided at the tip end with a
radially outer airfoil platform, for example, in the form of a
shroud.
[0004] An axial gap is formed between the guide vane cascade and
the rotor blade cascade. When the rotor blade cascade rotates,
pressure gradients are formed therein, the pressure gradients
varying around the circumference and causing secondary flows. For
example, a rotating cascade of turbine rotor blades may force
working fluid into the axial gap on its pressure side and,
conversely, draw working fluid from the gap on its suction side. As
a result, a compensating flow is generated, which degrades the
efficiency of the turbomachine.
[0005] A gas turbine having shroudless rotor blades is disclosed in
EP 2 372 102 A2, which proposes that the radially inner platforms
of guide vanes and rotor blades have a non-axisymmetric contour, in
particular a radially and/or axially undulated contour.
[0006] European Patent Publication EP 2 136 033 A1 discloses a
turbomachine stage having guide vanes which, together with radially
inner and/or radially outer airfoil platforms, form a guide vane
cascade. The turbomachine further has rotor blades which, together
with radially inner and/or radially outer airfoil platforms, form a
rotor blade cascade. Provided between the guide vane cascade and
the rotor blade cascade is an axial gap which is bounded by gap
regions of the airfoil platforms of the rotor blade cascade and the
guide vane cascade. In this connection, only the contour of the gap
region of the airfoil platform of the rotor blade cascade varies in
the radial and/or axial direction around the circumference.
[0007] European Patent Publication EP 1 067 273 A1 describes a
turbomachine stage having a rotor blade. The radially outer airfoil
platform associated with the rotor blade has a contour which varies
in the axial direction, while the radially outer airfoil platform
associated with a guide vane has a contour which does not vary in
the axial direction.
SUMMARY OF THE INVENTION
[0008] It is an object of the present invention to improve the
efficiency of a turbomachine, in particular an aircraft engine gas
turbine.
[0009] A turbomachine stage according to the present invention
includes a plurality of rotor blades which are preferably
equidistantly distributed around the circumference and, at their
root or rotor end, are connected to, in particularly integrally
formed with, radially inner airfoil platforms. At their tip or
casing ends, the rotor blades may be connected to, in particularly
integrally formed with, radially outer airfoil platforms. Rotor
blades may be removably or non-removably attached to, in particular
integrally formed with, a rotor (member) of the turbomachine,
either individually or in groups.
[0010] On the upstream and/or downstream side(s) of the cascade
formed by these rotor blades, a plurality of guide vanes are
preferably equidistantly distributed around the circumference and
removably or non-removably attached to, in particular integrally
formed with, a casing (member) of the turbomachine. To this end,
the guide vanes are connected to, in particular integrally formed
with, radially outer airfoil platforms. At their tip or rotor ends,
the guide vanes may be connected to, particularly integrally formed
with, radially inner airfoil platforms.
[0011] At their tip or rotor ends, the guide vanes may be connected
to, in particular integrally formed with, radially inner airfoil
platforms.
[0012] Platform regions extending axially between the airfoil
leading and trailing edges and circumferentially between adjacent
airfoils, together with the airfoils themselves and, possibly,
casing or rotor surface regions, define flow channels for the
working fluid, and thus the rotor blade cascade or guide vane
cascade, respectively. Therefore, these platform regions are
hereinafter referred to as cascade regions.
[0013] However, the airfoil platforms may extend axially beyond
these cascade regions on the upstream and/or downstream side(s);
i.e., beyond the airfoil leading and/or trailing edges. These
regions of the airfoil platforms bound an axial gap extending
axially between the guide vane cascade and the rotor blade cascade
and, therefore, are hereinafter collectively referred to as gap
regions of the airfoil platforms.
[0014] An airfoil platform may have radially outer gap regions
including a plurality of sections. For example, the radially outer
airfoil platforms of a rotor blade cascade or a guide vane cascade
may have one or more radial shoulders whose circumferential
surfaces bound the axial gap radially and whose end faces bound the
axial gap axially. In the case of such radially outer gap regions
having a plurality of sections, the following explanations may
refer to one or more, in particular to all of the sections of a gap
region. Thus, for example, when the description speaks of a
variation of a radially outer gap region in the radial and/or axial
direction, the contour(s) of one or more circumferential surfaces
may vary in the radial direction and/or the contour(s) of one or
more end faces may vary in the axial direction.
[0015] If, in a preferred embodiment, rotor blade tips or shrouds
are, in particular sealingly, disposed in a recess of the casing,
the casing member in which the recess is formed may form a radially
outer gap region of the guide vane platforms according to the
present invention. Similarly, a radially outer rotor blade platform
which may be disposed in particular in a recess of the casing may
form a gap region according to the present invention.
[0016] In general, a component which is radially outwardly
connected to, or integrally formed with, at least one guide vane or
rotor blade and whose contour, possibly together with additional
contours, radially and/or axially bounds the axial gap between the
rotor blade cascade and the guide vane cascade, may constitute a
radially outer gap region of an airfoil platform according to the
present invention.
[0017] In accordance with the present invention, a contour of one
or more of these gap regions varies in the radial and/or axial
direction around the circumference. A variation in the radial
direction is understood, in particular, to be an outside radius R
of the contour which, in polar coordinates, varies with the
circumferential angle .phi. around the axis of rotation of the
turbomachine stage, and analogously, a variation in the axial
direction is understood, in particular, to be an axial coordinate X
of the contour which varies with the circumferential angle.
Preferably, the contour varies periodically, in particular
sinusoidally:
R(.phi.)=R.sub.0+.DELTA.R.times.sin(.OMEGA..sub.R.times..phi.+.PHI..sub.-
R)and/or
X(.phi.)=X.sub.0+.DELTA.X.times.sin(.OMEGA..sub.x.times..phi.+.PHI..sub.-
x),
where
.phi..epsilon.[0.degree.,360.degree.],R.sub.0,.DELTA.R,X.sub.0,.DELTA.X,-
.OMEGA..sub.R,.OMEGA..sub.x,.PHI..sub.R,.PHI..sub.x=const.
or asymmetrically.
[0018] As explained above, this variation (hereinafter also
referred to as undulation) may be formed solely in the radial
direction, solely in the axial direction, or in both the axial and
radial directions. For example, the contour of a cylindrical gap
region having a smooth end face and an undulated circumferential
surface varies solely in the radial direction, that of a
cylindrical gap region having an undulated end face and a smooth
circumferential surface varies solely in the axial direction, while
that of a cylindrical gap region having an undulated end face and
an undulated circumferential surface and that of conical gap region
having an undulated circumferential surface vary in both the axial
and radial directions.
[0019] The undulation may be formed solely on one or more gap
regions of radially outer guide vane platforms, solely on one or
more gap regions of radially outer rotor blade platforms, or also
on one or more gap regions of radially outer platforms of both
guide vanes and rotor blades. In a preferred refinement, an
undulation may additionally be provided on one or more gap regions
of radially inner guide vane platforms and/or on one or more gap
regions of radially inner rotor blade platforms.
[0020] In a preferred embodiment, a contour of a gap region of an
airfoil platform of one of the guide vane and rotor blade cascades
and an axially and/or radially opposite contour of a gap region of
an airfoil platform of the other of the guide vane and rotor blade
cascades may vary around the circumference, preferably identically,
in particular in parallel, or with a phase offset of preferably at
least 45.degree., in particular at least 90.degree., preferably at
least 135.degree. and/or preferably of no more than 270.degree., in
particular no more than 210.degree., and preferably no more than
180.degree..
[0021] If a gap region has two opposite contours, such as an inner
and an outer circumferential surface of an annular flange such as,
in particular, a shroud extension, or on a casing, then these two
opposite contours may vary around the circumference, preferably
differently or identically, in particular in parallel, or with a
phase offset of preferably at least 45.degree., in particular at
least 90.degree., preferably at least 135.degree. and/or preferably
of no more than 270.degree., in particular no more than
210.degree., and preferably no more than 180.degree.. If the two
contours vary in parallel, the wall thickness of this gap region of
the airfoil platform remains constant. It may equally be provided
that only one of such opposite contours, in the case of an
annular-flange-like shroud extension preferably the radially inner
contour, varies while the other remains constant around the
circumference.
[0022] In general, an entire contour of a gap region, for example,
the entire inner circumferential surface of an annular flange, may
vary around the circumference. It is equally possible that only a
section of the contour has an undulation. For example, the inner
circumferential surface of an annular flange may vary in the radial
direction only in one or more axial sections, or an end face may
vary in the axial direction only in one or more radial
sections.
[0023] A radial variation of a contour of a gap region of an
airfoil platform of a cascade may be constant in the axial
direction, so that troughs and crests are oriented parallel to the
axis of rotation of the turbomachine stage. Equally, a radial
variation of a contour of a gap region of an airfoil platform of a
cascade may also vary in the axial direction, so that troughs and
crests extend at an angle to the axis of rotation. In particular, a
phase offset may be provided which varies with the axial position
x, preferably linearly:
R(.phi.,x)=R.sub.0+.DELTA.R.times.sin(.OMEGA..sub.R.times..phi.+.PHI..su-
b.R.times.x)
[0024] Similarly, an axial variation of a contour of a gap region
of an airfoil platform of a cascade may be constant in the radial
direction, so that troughs and crests are oriented perpendicularly
to the axis of rotation of the turbomachine stage. Equally, an
axial variation of a contour of a gap region of an airfoil platform
of a cascade may also vary in the radial direction, so that troughs
and crests are inclined at an angle to the axis of rotation. Here,
too, a phase offset may be provided which varies with the radial
position r, preferably linearly:
X(.phi.,x)=X.sub.0+.DELTA.X.times.sin(.OMEGA..sub.x.times..phi.+.PHI..su-
b.x.times.x)
[0025] In a preferred embodiment, in addition to at least one gap
region, the cascade region of the airfoil platform varies as well,
at least partially, around the circumference in one of the ways
described above. In an advantageous refinement, a gap region whose
contour varies around the circumference merges smoothly into this
cascade region, especially in such a way that a trough of the gap
region contour merges into a trough of the cascade region, and a
crest of the gap region contour merges into a crest of the cascade
region. As is customary in the art, the term "smooth transition" is
used, in particular, to refer to a transition which has no sharp
edges or bends, but which preferably has a continuous
curvature.
[0026] In a preferred embodiment, an extreme extent; i.e., a
maximum or minimum extent of a radially varying contour of a gap
region of an airfoil platform of a cascade is circumferentially
located in the pressure-side half of the segment between two
adjacent airfoil leading edges or in the suction-side half of the
segment between two adjacent airfoil trailing edges of the cascade
in order to compensate the pressure increases and decreases induced
there.
[0027] In general, two airfoil leading or trailing edges of a guide
vane or rotor blade cascade define a segment therebetween which
extends in the circumferential direction and is divided into two
halves by the channel center. The segment half adjoining the
pressure side of the airfoil is referred to as "pressure-side
half", the other one as "suction-side half" accordingly. These
halves define a circumferential angular range in which, according
to the present invention, an extreme extent of a varying contour is
disposed. Since the varying contour does not lie at the axial
height of this segment itself, this segment may be imagined as
being displaced parallel to an extension of the mean camber line of
the airfoil for the positioning of the extreme extent.
[0028] Preferably, a maximum variation in the radial direction of
gap region of an airfoil platform of a cascade is no more than 50%,
particularly no more than 40% of the pitch of the cascade.
[0029] In a preferred embodiment, in addition or as an alternative
to the aforementioned positioning of extreme radial extents around
the circumference, an extreme; i.e., a maximum or minimum extent of
an axially varying contour of a gap region may be circumferentially
located in the region of an airfoil edge, in particular
circumferentially spaced from the airfoil edge by no more than 25%
of the cascade pitch. Here, too, reference is made to the segment
which extends between two adjacent airfoil edges, whose arc length
defines the cascade pitch, and which may be imagined as being
displaced parallel to the extension of the mean camber line of the
airfoil. A maximum extent of a guide vane platform in the axial
direction extends preferably axially toward the rotor blade cascade
and, analogously, a maximum extent of a rotor blade platform in the
axial direction extends preferably axially toward the guide vane
cascade.
[0030] Preferably, a maximum variation in the axial direction of
gap region of an airfoil platform of a cascade is no more than 50%,
particularly no more than 40% of the pitch of the cascade.
BRIEF DESCRIPTION OF THE DRAWINGS
[0031] Further features and advantages will become apparent from
the dependent claims and the exemplary embodiments. To this end,
the drawings show, partly in schematic form, in:
[0032] FIG. 1: a developed view of a portion of a gas turbine stage
according to the present invention including a guide vane cascade
and a rotor blade cascade having radially outer airfoil platforms
whose gap region contours vary in the axial direction around the
circumference;
[0033] FIGS. 2A, 2B: meridional sections at different
circumferential positions through a gas turbine stage according to
the present invention including a guide vane cascade and a rotor
blade cascade having radially outer airfoil platforms whose gap
region contour varies in the radial direction around the
circumference; and
[0034] FIG. 3: a meridional section, similar to those of FIG. 2,
through a gas turbine stage according to the present invention, in
which a gap region of a radially outer guide vane platform has a
recess formed therein for receiving radially outer rotor blade
platforms.
DETAILED DESCRIPTION
[0035] FIG. 1 shows a developed view of a portion of a gas turbine
stage according to the present invention, as seen from an axis of
rotation; i.e., viewed from radially inside to radially outside,
showing a stationary cascade of guide vanes 1 and, opposite
thereto, a rotating cascade of rotor blades 2. The rotation is
indicated by a filled vertical arrow, the flow of working fluid is
indicated by an empty arrow in the region of the guide vane
cascade. This configuration is merely exemplary for purposes of
illustration. The present invention may be used equally in turbine
and compressor stages, where the guide vane cascade is disposed
upstream and/or downstream of the rotor blade cascade.
[0036] Integrally formed with airfoils 1, 2 are radially outer
airfoil platforms, which are shown from above in FIG. 1; i.e., as
viewed from the axis of rotation of the gas turbine stage. Each
airfoil may either have a separate airfoil platform, or several or
all of the airfoils of a cascade may be connected to, in particular
integrally formed with, the same airfoil platform which, in
accordance with the present invention, may then be imagined as
being divided into separate airfoil platforms associated with the
individual airfoils. Therefore, FIG. 1 does not show any airfoil
platform boundaries in the circumferential direction (vertically in
FIG. 1). The radially outer platforms of guide vanes 1 may be, for
example, a part, in particular an integral part, of a casing of a
gas turbine (stage), or be attached to such a casing. The radially
outer platforms of rotor blades 2 may be, for example, shrouds, in
particular interconnected shrouds.
[0037] A cascade region 10.1 of the guide vane platforms and a
cascade region 20.1 of the rotor blade platforms extend axially
between the respective leading edge (left in FIG. 1) and the
respective trailing edge (right in FIG. 1), said cascade regions
being hatched from top left to bottom right in FIG. 1.
[0038] The cascade regions merge axially into respective gap
regions 10.2T and 20.2L beyond the respective airfoil leading or
trailing edges, said gap regions being hatched from bottom left to
top right in FIG. 1. Gap regions 10.2T and 20.2L each have
substantially the shape of a radial shoulder whose circumferential
surface facing toward the spoke-like pattern of the respective
cascade and whose end face facing toward the respective other
airfoil cascade radially and axially bound a radially inner axial
gap A between the rotor blade cascade and the guide vane
cascade.
[0039] As can be seen in the developed view of FIG. 1, the contour
of this gap region 10.2T, respectively 20.2L, and more particularly
its end face facing the respective other airfoil cascade, varies in
the axial direction around the circumference; i.e., in the vertical
direction in FIG. 1. That is, the generating lines of the end face,
which extend from the axis of rotation of the turbomachine to the
peripheral edge of the radial shoulder, have different axial
positions, so that the end face has a maximum axial extent
A.sub.max10, respectively A.sub.max20, toward the respective other
airfoil cascade at selected circumferential positions, as measured
from a generating line which is axially farthest away from the
respective other airfoil cascade. The generating lines may be
perpendicular to the axis of rotation of the turbomachine, or
inclined thereto at the same angle or at an angle that varies in
the circumferential direction. In the exemplary embodiment, the
generating lines are perpendicular to the axis of rotation. Their
axial position varies sinusoidally around the circumference, so
that maximum axial extents A.sub.max20 of gap region 20.2L of the
rotor blade platforms are disposed near respective leading edges of
rotor blades 2, and maximum axial extents A.sub.max10 of gap region
10.2T of the guide vanes are disposed near respective trailing
edges of guide vanes 1, as viewed in the circumferential direction.
Maximum extents A.sub.max10 and A.sub.max20 are each 50% of the
respective cascade pitch.
[0040] With regard to the position in the circumferential
direction, instead of making reference to a position of airfoil
edges, the channel center, and the like, which position is
displaced parallel to the axis of rotation, reference may also be
made to a position which is displaced parallel to the extension of
the mean camber line of the respective airfoil. For this purpose,
the extension of mean camber line 2.1 of rotor blades 2 is
indicated by a dot-dash line. It can be seen that the maximum axial
extents A.sub.max20 of gap region 20.2L of the rotor blade
platforms are still circumferentially located near the so-displaced
positions of the leading edges of rotor blades 2.
[0041] FIGS. 2A, 2B show meridional sections at different
circumferential positions through a gas turbine stage according to
the present invention including a guide vane cascade and a rotor
blade cascade having radially outer airfoil platforms whose gap
region contour varies in the radial direction around the
circumference. In particular, this gas turbine stage may be the one
described hereinabove with reference to FIG. 1, so that a radial
undulation is combined with an axial undulation. Therefore, in the
following, reference is made to the above description and only the
aspects of the radial undulation will be described. It is equally
possible to provide only an axial undulation, as described
hereinabove with reference to FIG. 1, or only a radial undulation,
such as will be described hereinafter.
[0042] FIG. 2A shows a meridional section at a circumferential
position where the circumferential surface of gap region 20.2L has
a minimum radial extent in a radially outward direction; i.e., in a
direction away from the rotor and, analogously, FIG. 2B shows a
meridional section at a circumferential position where the
circumferential surface of gap region 20.2L has a maximum radial
extent. It can be seen that the circumferential surface of gap
region 20.2L varies sinusoidally in the circumferential direction
with an amplitude AR=(R.sub.max20+R.sub.min20)/2. The maximum
positive amplitude; i.e., the maximum radial extent R.sub.max20 in
the radially outward direction, is circumferentially located in the
pressure-side half of the segment between two successive rotor
blade leading edges. Again, the positions in the circumferential
direction may be imagined as being displaced parallel to the
extension of the mean camber line to the respective axial
positions.
[0043] It can be seen that the radial undulation varies not only
around the circumference (compare FIG. 2A to FIG. 2B), but also in
the axial direction (see the horizontal direction in FIG. 2A, 2B),
so that the crests and troughs are inclined at an angle to the axis
of rotation. In particular, FIG. 2A shows the trough R.sub.min20
sloping upwardly in the direction of the flow, while FIG. 2B shows
the crests R.sub.max20 sloping downwardly in the direction of the
flow.
[0044] Moreover, it can be seen that this radial undulation of gap
region 20.2 merges smoothly into a corresponding undulation of
cascade region 20.1 between rotor blades 2.
[0045] Although not shown, in addition or as an alternative to the
above-described axial undulation (see FIG. 1), the gap region 10.2T
of the guide vane platforms, which faces the rotor blade cascade,
may also have a radial undulation, such as described hereinabove
with reference to gap region 20.2 of the rotor blade platforms.
[0046] It can be seen in FIG. 2 that this trailing-edge gap region
is shaped like an annular flange, and therefore has two radially
opposite surfaces (upper and lower in FIG. 2). A radial undulation
may in particular be provided on the radially inner surface (lower
in FIG. 2) or on both surfaces. In the latter case, preferably,
they vary identically, so that the wall thickness of the annular
flange remains constant.
[0047] FIG. 3 shows, in a view similar to that of FIG. 2, a portion
of a gas turbine stage according to a modified embodiment of the
present invention. Corresponding elements are identified by the
same reference numerals, so that reference is made to the above
explanations in their entirety, and only the differences in the
modified embodiment will be discussed below.
[0048] Firstly, FIG. 3 shows the downstream trailing edge of a
rotor blade 2 and the upstream leading edge of a following guide
vane 1. Accordingly, the radially outer gap regions are designated
20.2T (for trailing edge) and 10.2L (for leading edge, and further
with an M or S designation as discussed below) to illustrate by way
of example that the explanations are equally applicable to leading
and trailing edges of rotor blade platforms and guide vane
platforms, respectively.
[0049] Secondly, the radially outer airfoil platforms are inclined
at an angle to the turbine axis to illustrate a divergent flow
channel. The explanations apply equally to convergent flow channels
(not shown), in particular in compressor stages.
[0050] Moreover, the radially outer rotor blade platform, which is
in the form of a shroud, has an annular flange formed in its
trailing-edge gap region 20.2T. This annular flange has radially
opposite circumferential surfaces, such as described hereinbefore
with reference to trailing-edge gap region 10.2T of guide vane 1 of
FIG. 1. Trailing-edge gap region 20.2T may also have an undulation,
in particular the same undulation, on its radially inner (lower in
FIG. 3) and/or outer circumferential surface.
[0051] The leading-edge gap region includes a radially inner
annular flange in a radially outer groove-like recess of the gas
turbine casing. Accordingly, the leading-edge gap region has three
circumferential surfaces 10.2LM, namely the radially inner and
radially outer circumferential surfaces of the annular flange and
the circumferential surface of the recess itself, as well as two
end faces 10.2LS, namely that of the annular flange and that of the
recess itself.
[0052] Each of these sections 10.2LM, 10.2LS may have an undulation
in the radial direction (10.2LM) and in the axial direction
(10.2LS), respectively. Moreover, it is possible that several, in
particular all, sections of the gap region have an undulation. In
this regard, the modification shown in FIG. 3 is intended to
illustrate in one view different variants where a contour of a gap
region may vary in the radial and/or axial direction around the
circumference.
LIST OF REFERENCE NUMERALS
[0053] 1 guide vane [0054] 2 rotor blade [0055] 2.1 extension of
the mean camber line [0056] 10.1/20.1 cascade region of the
radially outer airfoil platform of the rotor blade or guide vane
cascade [0057] 10.2T trailing-edge gap region of the radially outer
airfoil platform of the guide vane cascade [0058] 10.2LM
circumferential surface section of the leading-edge gap region of
the radially outer airfoil platform of the guide vane cascade
[0059] 10.2LS end face section of the leading-edge gap region of
the radially outer airfoil platform of the guide vane cascade
[0060] 20.2T trailing-edge gap region of the radially outer airfoil
platform of the rotor vane cascade [0061] 20.2L leading-edge gap
region of the radially outer airfoil platform of the rotor vane
cascade [0062] A axial gap
* * * * *