U.S. patent application number 13/054162 was filed with the patent office on 2011-08-04 for axial turbine for a gas turbine with limited play between blades and housing.
Invention is credited to Stefan Braun, Christian Cornelius, Annika Emde, Andreas Heilos, Olaf Hein, Thomas Hofbauer, Christian Lerner, Silvio-Ulrich Martin, Thorsten Mattheis, Ralf Musgen, Eckart Schumann, Rostislav Teteruk, Adam Zimmermann.
Application Number | 20110188999 13/054162 |
Document ID | / |
Family ID | 40010789 |
Filed Date | 2011-08-04 |
United States Patent
Application |
20110188999 |
Kind Code |
A1 |
Braun; Stefan ; et
al. |
August 4, 2011 |
Axial turbine for a gas turbine with limited play between blades
and housing
Abstract
An axial turbine for a gas turbine including a rotor blade
cascade is provided. The rotor blade cascade is formed from rotor
blades each including, a front edge, a blade tip, and an annular
space wall that surrounds the rotor blade cascade and includes an
annular space inner side, enabling the annular space wall to be
arranged directly adjacent to the blade tip forming a radial gap
between the covering of the blade tip and the annual space inner
side. When the turbine is in operation, the area of the blade tip
with the highest pressure load is disposed in the region of the
front edges, and the rotor blades in the region of their front
edges include a radial projection and the annular space wall
includes on the annular space inner side, a peripheral radial
recess that interacts with the radial projections such that a
minimum is established in the direction of the main through-flow of
the turbine.
Inventors: |
Braun; Stefan;
(Neukirchen-Vluyn, DE) ; Cornelius; Christian;
(Sprockhovel, DE) ; Emde; Annika; (Mulheim,
DE) ; Heilos; Andreas; (Mulheim an der Ruhr, DE)
; Hein; Olaf; (Mulheim an der Ruhr, DE) ;
Hofbauer; Thomas; (Wustenroth, DE) ; Lerner;
Christian; (Dorsten, DE) ; Martin; Silvio-Ulrich;
(Oberhausen, DE) ; Mattheis; Thorsten; (Mulheim,
DE) ; Musgen; Ralf; (Essen, DE) ; Schumann;
Eckart; (Mulheim an der Ruhr, DE) ; Teteruk;
Rostislav; (Mulheim an der Ruhr, DE) ; Zimmermann;
Adam; (Mulheim a.d. Ruhr, DE) |
Family ID: |
40010789 |
Appl. No.: |
13/054162 |
Filed: |
July 8, 2009 |
PCT Filed: |
July 8, 2009 |
PCT NO: |
PCT/EP2009/058682 |
371 Date: |
April 14, 2011 |
Current U.S.
Class: |
415/170.1 |
Current CPC
Class: |
F01D 5/143 20130101;
F01D 5/20 20130101; F01D 5/141 20130101 |
Class at
Publication: |
415/170.1 |
International
Class: |
F01D 11/00 20060101
F01D011/00 |
Foreign Application Data
Date |
Code |
Application Number |
Jul 17, 2008 |
EP |
08012960.4 |
Claims
1.-8. (canceled)
9. An axial turbine for a gas turbine including a rotor blade
cascade, comprising: a plurality of rotor blades each including a
leading edge, a trailing edge, and a radially outwardly disposed,
free-standing blade tip; and a divergent annulus wall, encasing the
rotor blade cascade, with an annulus inner side by which the
annulus wall is arranged directly adjacently to the blade tips,
forming a radial gap between the contours of the blade tips and the
annulus inner side, wherein the plurality of rotor blades at the
plurality of blade tips each include a region with a highest
pressure load of the plurality of blade tips between the leading
edge and the trailing edge, and wherein the plurality of rotor
blades in the region of the highest pressure load include in each
case a radial projection, wherein the annulus wall on the annulus
inner side includes an encompassing radial recess which, lying
opposite the radial projections, is formed in such a way that a
progression of the radial gap, as seen in the principal throughflow
direction of the axial turbine, extends essentially with constant
width in a wave-like, edge-free and step-free manner, wherein the
progression of the radial recess, as seen in the principal
throughflow direction of the axial turbine, on the annulus inner
side includes a first curvature section, a second curvature section
adjoining the first, and a third curvature section adjoining the
second, and wherein the first curvature section is delimited from
the second curvature section by a first inflection point and the
second curvature section is delimited from the third curvature
section by a second inflection point so that the curvatures of the
first curvature section and the third curvature section include the
same sign which is different from the sign of the curvature of the
second curvature section.
10. The axial turbine as claimed in claim 11, wherein the region of
the highest pressure load of the rotor blade is arranged in a
region of the leading edge.
11. The axial turbine as claimed in claim 12, wherein the region of
the highest pressure load of the rotor blade, amounts at most to
20% of a first height of the rotor blade and a remaining region of
a second height of the rotor blade includes a further highest
pressure load which is arranged in a region of the trailing
edge.
12. The axial turbine as claimed in claim 12, wherein the radial
recess is arranged in the front third of the radial gap, as seen in
the principal throughflow direction of the axial turbine.
13. The axial turbine as claimed in claim 12, wherein a curvature
of the first curvature section is greater than that of the third
curvature section.
14. The axial turbine as claimed in claim 12, wherein the
progression of the radial projections, as seen in the principal
throughflow direction of the axial turbine, on the plurality of
sides of the radial projections facing the radial gap, is adapted
to the progression of the radial recess.
15. The axial turbine as claimed in claim 11, wherein the first
inflection point is located in the region of the leading edge.
16. The axial turbine as claimed in claim, wherein a plurality of
sections of the radial gap, as seen in the principal throughflow
direction of the axial turbine, which are upstream and downstream
of and adjacent to the radial recess, are conical.
Description
CROSS REFERENCE TO RELATED APPLICATIONS
[0001] This application is the US National Stage of International
Application No. PCT/EP2009/058682, filed Jul. 8, 2009 and claims
the benefit thereof. The International Application claims the
benefits of European Patent Office application No. 08012960.4 EP
filed Jul. 17, 2008. All of the applications are incorporated by
reference herein in their entirety.
FIELD OF INVENTION
[0002] The invention refers to an axial turbine for a gas turbine,
the axial turbine having low tip clearance losses.
BACKGROUND OF INVENTION
[0003] A gas turbine has a turbine, for example in an axial type of
construction. The turbine has a casing and a rotor which is
enclosed by the casing. The rotor has a shaft, from which shaft
output can be tapped off. Provision is made for a hub, encompassing
the shaft, the hub contour of which, together with the inner
contour of the casing, forms a flow passage through the turbine.
The flow passage has a cross section which diverges in the flow
direction on account of a mostly conical inner contour of the
casing.
[0004] The rotor has a multiplicity of rotor stages which are
formed in each case by a rotor blade cascade. The rotor blade
cascades have a multiplicity of rotor blades which by one of their
ends are fastened in each case on the rotor on the hub side and by
their other end point radially outwards. A blade tip, which faces
the inner side of the casing and is arranged directly adjacent
thereto, is formed at the other end of the rotor blade. The
distance between the blade tips and the inner side of the casing is
formed as a radial gap which is dimensioned in such a way that on
the one hand the blade tips do not rub against the casing during
operation of the gas turbine and on the other hand the leakage flow
through the radial gap which ensues during operation of the gas
turbine is as low as possible. So that the gas turbine has high
efficiency, it is desirable that the leakage flow through the
radial gap is as low as possible so that the power gain in the
turbine is as high as possible.
[0005] The casing of the turbine is solidly constructed in order to
be able to withstand the pressure stresses and temperature stresses
during operation of the gas turbine. Furthermore, the casing is
rigidly constructed so that the load yield to the casing during
operation of the gas turbine results in only minimal deformation of
the casing. In contrast to this, the rotor blades are thinner and
less solidly constructed in comparison to the casing.
[0006] During operation of the axial turbine, the inner side of the
casing and the rotor blades are in contact with hot gas, the rotor
blades being completely exposed to circumflow by the hot gas. Due
to the fact that the rotor blades are of a more filigree design
than the casing and are in more extensive contact with the hot gas
than the casing, the rotor blades heat up more quickly than the
casing. This has the result that for startup and shutdown of the
gas turbine the rotor blades and the casing have different rates of
thermal expansion so that during startup and shutdown of the gas
turbine the height of the radial gap changes, the radial gap
becoming smaller during startup and larger during shutdown. So that
during startup the blade tips of the rotor blades do not butt
against the casing and damage this, the radial gap is provided with
a minimum height which is dimensioned in such a way that during
startup of the gas turbine the blade tips seldom if ever come into
contact with the casing. This has the result that provision is made
for a correspondingly dimensioned radial gap at the blade tips
which leads to a reduction in the power density and efficiency of
the gas turbine.
[0007] Modern rotor blades have a very high aerodynamic efficiency
which is achieved as a result of a high pressure load of the rotor
blades. Brought about by the high pressure load, the leakage flow
through the radial gap is high so that the overall efficiency of
the rotor blade is seriously impaired as a result of the character
and the intensity of the leakage flow through the radial gap. A
reduction to the losses which are brought about by the leakage flow
has the effect of a great improvement in the overall efficiency of
the rotor blade.
[0008] Attempts are customarily made to reduce the aerodynamic
losses in the gap region of the rotor blade by means of measures
for reducing the leakage flow. In this case, provision is made for
measures for reducing the radial gap or for a special profiling of
the blade tips, such as crowns or directed cooling air blowouts.
Alternatively to this, a rotor blade with a curved blade tip which,
for forming a minimum radial gap, can rub a groove into an
oppositely disposed passage wall, is known from DE 10 2004 059 904
A1. The curvature is achieved by means of an abrasive coating, with
varying coating thickness, which is applied to the blade tip. On
the edges, i.e. on the leading edge and trailing edge, the coating
is formed in a manner in which it runs out so that the rubbed-in
groove merges into the adjacent passage walls without steps. The
more costly production process of the rotor blade is considered to
be disadvantageous and requires the application of the rubbing-in
process, which results in an increased minimum strength of rotor
blades.
[0009] Conventional turbine rotor blades are configured according
to the "rear-loaded design", the maximum pressure stress of the
rotor blade being located in the region of its trailing edge. Known
as being obsolete are also rotor blades configured according to the
"front-loaded design", in which the highest pressure load is
located in the region of the leading edge. In this respect, a
turbine rotor blade with a blade airfoil is known for example from
EP 1 057 969 A2, which blade airfoil on the hub side has a
"front-loaded design" or "intermediate-loaded design", and on the
tip side has a "rear-loaded design", as a result of which the
distribution of the rate of change of the tip speed is
facilitated.
SUMMARY OF INVENTION
[0010] It is the object of the invention to create an axial turbine
for a gas turbine which has high aerodynamic efficiency.
[0011] The axial turbine according to the invention for a gas
turbine has a rotor blade cascade, which is formed from rotor
blades having in each case a leading edge, a trailing edge and a
radially outwardly disposed, free-standing blade tip, an annulus
wall which encases the rotor blade cascade, with an annulus inner
side by which the annulus wall is arranged directly adjacently to
the blade tips, forming the radial gap between the contours of the
blade tips and the annulus inner side, wherein the rotor blades at
their blade tips have a region with the highest pressure load of
the blade tips between the leading edge and the trailing edge, and
wherein in the region of the highest pressure load the rotor blades
have in each case a radial projection and the annulus wall on the
annulus inner side has an encompassing radial recess which lies
opposite the radial projections. The pressure load, in the sense of
this document, corresponds in this case to the pressure difference
between suction side and pressure side of the rotor blade, which is
of variable value along the profile section.
[0012] As a result, by making use of the blade tip, which is
optimized directly with regard to minimum losses, and the annulus
contour, the unfavorable, loss-affected gap flow is reduced. In
this case, the annulus in the region of the blade tip is
constructed as a contour which deviates from the conventional
annulus. When establishing the shape of the annulus contour,
moreover, consideration is given to the fact that the minimum gap
width during operation of the axial turbine is arranged in the
region of the maximum pressure difference between the pressure side
and the suction side of the rotor blade. These measures have almost
no influence upon the aerodynamic principle of operation of the
rotor blade and bring about a significant reduction in the gap flow
from the pressure side to the suction side over the blade tip
compared with a conventionally designed axial turbine. Furthermore,
it is possible to additionally apply all previously known measures
for reducing the negative effects of leakage flow in the axial
turbine according to the invention.
[0013] Furthermore, the radial recess and the radial projections
are formed in such a way that the progression of the radial gap, as
seen in the principal flow direction of the axial turbine, extends
essentially with constant width in a wave-like and step-free
manner.
[0014] The progression of the radial recess, as seen in the
principal flow direction of the axial turbine, on the annulus inner
side has a first curvature section, a second curvature section
adjoining the first, and a third curvature section adjoining the
second, wherein the first curvature section is delimited from the
second curvature section by a first inflection point and the second
curvature section is delimited from the third curvature section by
a second inflection point so that the curvatures of the first
curvature section and of the third curvature section have the same
sign which is different from the sign of the curvature of the
second curvature section. In this case, the size of the radial gap
between blade tip and annulus wall--as seen along the axial
direction--can also be constant.
[0015] Consequently, the annular gap, as seen in the principal flow
direction, has a uniform, non-abruptly changing progression so that
the flow in the region of the blade tip is low in loss.
[0016] The volume of leakage flow is advantageously reduced in a
directly aimed manner and its unfavorable effects upon the overall
efficiency of the rotor blade cascade are reduced. As a result, an
improved aerodynamic quality of the rotor blade cascade ensues
without having to provide additional constructional measures.
[0017] The profile section at the blade tip can be advantageously
constructed, contrary to the conventional construction as a
"front-loaded design". That is to say, the largest pressure load is
shifted from the rear part (close to the trailing edge) of the
blade into the region of the profile inlet edge (close to the
leading edge). This region, as seen over the height of the rotor
blade, can amount to about 20%. The remaining region of the rotor
blade can then be conventionally constructed in the "rear-loaded
design". The transition from "front-loaded design" to "rear-loaded
design" within some 20% of the height of the rotor blade is carried
out preferably steplessly.
[0018] It is preferred that with regard to the extent of the radial
gap, as seen in the principal flow direction of the axial turbine,
the radial recess is arranged in the front third.
[0019] Consequently, the radial recess is located in the region of
the highest pressure load of the blade tip so that the gap flow is
reduced.
[0020] It is preferred that the progression of the radial
projections, as seen in the principal flow direction of the axial
turbine, on their sides facing the radial gap is adapted to the
progression of the radial recess.
[0021] Moreover, it is preferred that the curvature of the first
curvature section is greater than that of the third curvature
section. Furthermore, it is preferred that the first inflection
point is located in the region of the leading edge.
[0022] It is preferred that the sections of the annular passage
which upstream and downstream are adjacent to the radial recess, as
seen in the principal flow direction of the axial turbine, are
conical.
BRIEF DESCRIPTION OF THE DRAWINGS
[0023] The invention is explained in the following text, based on a
preferred embodiment of an axial turbine according to the invention
with reference to the attached schematic drawings. In the
drawing:
[0024] FIG. 1 shows a profile section of a rotor blade according to
the invention in the region of the blade tip,
[0025] FIG. 2 shows a side view of an axial turbine according to
the invention, and
[0026] FIG. 3 shows the side view from FIG. 2 compared with a
conventional axial turbine.
DETAILED DESCRIPTION OF INVENTION
[0027] As is apparent from FIGS. 1 to 3, an axial turbine 1 has a
rotor blade 2 which has a leading edge 3 and a trailing edge 4.
[0028] The rotor blade 2 has a pressure side 5 and a suction side 6
which extend in each case from the leading edge 3 to the trailing
edge 4. The pressure side 5, compared with the suction side 6, is
more sharply concavely curved. The rotor blade 2, at its radially
outer side end, has a blade tip 13 which is freestanding. In the
region of the blade tip 13, the rotor blade 2 is constructed in the
"front-loaded design" 7. In comparison to this, the "rear-loaded
design" 8 is shown, in which the pressure side 5 is less sharply
curved in the region of the leading edge 3 than in the case of the
"front-loaded design" 7.
[0029] Due to the fact that the rotor blade 2 is constructed in the
"front-loaded design" 7 in the region of the blade tip 13, the
region 9 with the highest pressure load of the rotor blade 2 is
located in the region of the blade tip 13 in the proximity of the
leading edge 3.
[0030] Furthermore, the axial turbine 1 on the hub side has a hub
contour 10 on which the rotor blade 2 is fastened. The axial
turbine 1 has an annulus wall 11, terminating radially on the
outside, which has an annulus inner side 12 facing the blade tip
13. The rotor blade 2 is encased by the annulus wall 11 and with
the annulus inner side 13, together with the hub contour 10, forms
a divergent annulus of the axial turbine 1. The annulus wall 11 in
this case is principally--i.e. apart from a radial recess 15--of
conical design with a greater inclination than the hub contour
10.
[0031] Between the blade tip 13 and the annulus inner side 12,
provision is made for a space so that a radial gap 14 is formed
between the blade tip 13 and the annulus inner side 12.
[0032] In FIG. 3, the rotor blade 2 with a conventional blade tip
23 and the annulus wall 11 with a conventional annulus inner side
24 is also shown, wherein the conventional blade tip 23 and the
conventional annulus inner side 24 have a straight progression.
[0033] In contrast to this, the annulus wall 11 according to the
invention on the annulus inner side 12 has the radial recess 15
which is arranged in the region of the leading edge 3 of the rotor
blade 3. In correlation to the radial recess 15, and engaging in
this, a radial projection 16 is provided at the blade tip 13. The
radial projection 16 extends essentially parallel to the radial
recess 15 so that the radial gap 14 has a uniform progression, as
seen in the principal flow direction of the axial turbine 1.
[0034] As seen in the principal flow direction of the axial turbine
1, the radial recess has a first curvature section 17, a second
curvature section 19 adjoining the first, and a third curvature
section 21 adjoining the second. The first curvature section 17 is
delimited from the second curvature section 19 by a first
inflection point 18, and the second curvature section 19 is
delimited from the third curvature section 21 by a second
inflection point 20. Consequently, the curvature middle point of
the first curvature section 17 and of the third curvature section
21, as seen radially, lies outside the axial turbine 1 and the
curvature middle point of the second curvature section 19 lies
inside the axial turbine 1.
[0035] The curvature of the first curvature section 17 is greater
than the curvature of the third curvature section 21 so that the
radial gap 14 in the region of the leading edge 3 has a steeper
progression, as seen radially outwardly, than in the region of the
third curvature section 21.
[0036] As seen in the principal flow direction of the axial turbine
1, the radial recess 15 and the radial projection 16 are arranged
in the front third of the blade tip 13. Due to the fact that in the
region of the blade tip 13 the rotor blade 2 is formed in the
"front-loaded design", the region 9 with the highest pressure load
is located specifically in this region.
[0037] The radial recess 15 and the radial projection 16 are
arranged in relation to each other in such a way that a gap minimum
22 is formed in the region 9 of the highest pressure load. As a
result, a leakage flow through the radial gap 14, which develops
during operation of the axial turbine 1, is low, specifically in
the region 9 with the highest pressure load. Consequently, the
rotor blade 2 has high aerodynamic efficiency, especially in the
region of the blade tip 13.
* * * * *