U.S. patent application number 13/024545 was filed with the patent office on 2011-07-14 for blade arrangement of a gas turbine.
This patent application is currently assigned to ALSTOM TECHNOLOGY LTD. Invention is credited to Thomas DUDA, Thomas HEINZ-SCHWARZMAIER, Alexander SCHNELL.
Application Number | 20110171039 13/024545 |
Document ID | / |
Family ID | 40090118 |
Filed Date | 2011-07-14 |
United States Patent
Application |
20110171039 |
Kind Code |
A1 |
HEINZ-SCHWARZMAIER; Thomas ;
et al. |
July 14, 2011 |
BLADE ARRANGEMENT OF A GAS TURBINE
Abstract
A blade arrangement of a gas turbine, with at least one blade
which in the radial direction projects into a hot gas passage
arranged concentrically to an axis, and terminates in a blade tip
which with a clearance lies opposite a heat shield which delimits
the hot gas passage. The blade and the heat shield are movable in
relation to each other in the circumferential direction, and the
blade tip and the heat shield are covered with coatings, which
enable a directed cutting of the blade tip into the heat shield. By
such a blade arrangement, a reduction of the clearance as a result
of cutting in is simply achieved by the heat shield having a porous
thermal barrier coating as an outer, abradable coating, and by the
blade tip being provided with a homogenous, metallic cover
coating.
Inventors: |
HEINZ-SCHWARZMAIER; Thomas;
(Wettingen, CH) ; DUDA; Thomas; (Wettingen,
CH) ; SCHNELL; Alexander; (Baden, CH) |
Assignee: |
ALSTOM TECHNOLOGY LTD
Baden
CH
|
Family ID: |
40090118 |
Appl. No.: |
13/024545 |
Filed: |
February 10, 2011 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
PCT/EP2009/060387 |
Aug 11, 2009 |
|
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|
13024545 |
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Current U.S.
Class: |
416/241B |
Current CPC
Class: |
F01D 11/122 20130101;
F01D 11/14 20130101; F01D 5/288 20130101 |
Class at
Publication: |
416/241.B |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Foreign Application Data
Date |
Code |
Application Number |
Aug 15, 2008 |
CH |
01285/08 |
Claims
1. A blade arrangement (30) of a thermal turbomachine with at least
one blade (11), which projects in the radial direction into a
passage (13) which is arranged concentrically to an axis (16) and
is exposed to throughflow by hot gas, the at least one blade
terminates in a blade tip (27) which with a clearance (25) lies
opposite a heat shield (12) which delimits the passage (13),
wherein the at least one blade (11) and the heat shield (12) are
movable in relation to each other in a circumferential direction,
and the blade tip (27) and the heat shield (12) are covered with
coatings (22, 23, 24) which enable a directed cutting of the blade
tip (27) into the heat shield (12), the heat shield (12) has a
porous thermal barrier coating (23) as an outer, abradable coating,
and the blade tip (27) is provided with a homogenous, metallic
cover coating (24).
2. The blade arrangement as claimed in claim 1, wherein the thermal
turbomachine is a gas turbine, the at least one blade is a rotor
blade (11) which rotates around the axis (16), and the heat shield
(12) is installed on a stator of the gas turbine in a fixed
manner.
3. The blade arrangement as claimed in claim 1, wherein the thermal
barrier coating (23) is a porous ceramic coating, comprising
YSZ.
4. The blade arrangement as claimed in claim 3, wherein a porosity
of the thermal barrier coating (23) is more than 20%.
5. The blade arrangement as claimed in claim 3, wherein an adhesion
coating (22), comprising MCrAlY, is arranged between the heat
shield (12) and the thermal barrier coating (23).
6. The blade arrangement as claimed in claim 1, wherein the
metallic cover coating (24) comprises MCrAlY.
7. The blade arrangement as claimed in claim 2, wherein the rotor
blade (11) is part of a first rotor-blade row in a turbine section
of the gas turbine.
8. The blade arrangement as claimed in claim 2, wherein the thermal
barrier coating (23) is a porous ceramic coating, comprising
YSZ.
9. The blade arrangement as claimed in claim 8, wherein a porosity
of the thermal barrier coating (23) is more than 20%.
10. The blade arrangement as claimed in claim 8, wherein an
adhesion coating (22), comprising MCrAlY, is arranged between the
heat shield (12) and the thermal barrier coating (23).
11. The blade arrangement as claimed in claim 2, wherein the
metallic cover coating (24) comprises MCrAlY.
Description
CROSS REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of International
Application No. PCT/EP2009/060387 filed Aug. 11, 2009, which claims
priority to Swiss Patent Application No. 01285/08, filed Aug. 15,
2008, the entire contents of all of which are incorporated by
reference as if fully set forth.
FIELD OF INVENTION
[0002] The present invention relates to the field of gas turbine
technology in that it refers to a blade arrangement of a gas
turbine.
BACKGROUND
[0003] For the efficiency of a gas turbine, it is of great
importance, especially in the turbine section in which the hot
gases from the combustion chamber are expanded, to minimize as far
as possible the gaps which occur in the region of the blading
between the bladed, rotating rotor and the encompassing stator.
[0004] In the simplest case, as is reproduced in FIG. 1 with
reference to the blade arrangement 10, no special measures are
adopted for optimizing the gap width. The rotor blades 11 which
project radially into the hot gas passage 13 of the gas turbine and
rotate around the axis 16, terminate in a blade tip 27 which is
provided with a first cover coating 15 and which with a clearance
25 lies opposite a heat shield 12 which forms the outer wall 26 of
the hot gas passage 13 and is provided with a second cover coating
14. The cover coatings 14 and 15, which may comprise MCrAlY, for
example, protect the components 11 and 12 against the damaging
effects of the hot gases in the hot gas passage 13, especially
against undesirable oxidation. The blade arrangement 10 of FIG. 1
is not designed for cutting of the blade tips 27 into the heat
shield 12. The clearance 25 between the blade tips 27 and the heat
shield 12 makes allowance for the different thermal expansions and
is therefore comparatively large in order to avoid rubbing of the
components during operation.
[0005] In order to be able to reduce the clearance 25, blade
arrangements 20 according to FIG. 2 have been proposed (see, for
example, EP-A2-1 312 760 or US-A1-2008166225), in which abrasive
bodies (grains) 21 (for example consisting of cubic boron nitride
cBN), which are embedded in a carrier layer 19 and with which the
blade tip 27 can cut into an abradable thermal barrier coating
(TBC) 18 on the oppositely disposed heat shield 12 during
operation, are arranged on the blade tip 27 of the rotor blades. A
coating of metallic MCrAlY, for example, can be used as a carrier
layer 19 for the abrasive bodies 21 and also as an adhesion coating
17 beneath the thermal barrier coating 18.
[0006] The abrasive coating 19, 21 of such a blade arrangement is
of a comparatively complex construction as a result of the embedded
abrasive bodies and is therefore costly in production. The aim,
however, would have to be to create a comparable cutting-in
behavior without a special abrasive coating having to be provided
on the blade tip.
SUMMARY
[0007] The present disclosure is directed to a blade arrangement of
a thermal turbomachine with at least one blade, which projects in
the radial direction into a passage which is arranged
concentrically to an axis and is exposed to throughflow by hot gas.
The at least one blade terminates in a blade tip which with a
clearance lies opposite a heat shield which delimits the passage.
The blade and the heat shield are movable in relation to each other
in a circumferential direction, and the blade tip and the heat
shield are covered with coatings which enable a directed cutting of
the blade tip into the heat shield, the heat shield has a porous
thermal barrier coating as an outer, abradable coating, and the
blade tip is provided with a homogenous, metallic cover
coating.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] The invention shall subsequently be explained in more detail
based on exemplary embodiments in conjunction with the drawings. In
the drawings
[0009] FIG. 1 shows in a greatly simplified view an earlier blade
arrangement without the possibility of cutting in;
[0010] FIG. 2 shows in a view comparable to FIG. 1 another earlier
blade arrangement with a special abradable coating on the blade
tip, and
[0011] FIG. 3 shows in a view comparable to FIG. 1 a blade
arrangement intended for cutting in, with a simple cover coating on
the blade tip according to an exemplary embodiment of the
invention.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
Introduction to the Embodiments
[0012] The invention should provide a remedy in this case. It is
therefore the object of the invention to disclose a blade
arrangement which avoids the disadvantages of known blade
arrangements and with a simultaneously simple construction enables
a significant reduction of the clearance between the blade tips and
the oppositely disposed stator-side elements.
[0013] It is preferable for the heat shield to have a porous
thermal barrier coating as an outer, abradable coating and for the
blade tip to be provided simply with a homogenous metallic cover
coating. The porous thermal barrier coating enables the blade tip,
which is covered by the cover coating, to cut into the heat shield
even without special abrasive bodies or abrasive coating and so to
optimally minimize the clearance between blade tip and oppositely
disposed heat shield.
[0014] The blade is essentially a rotor blade or a stator blade of
a thermal turbomachine, in particular a gas turbine, wherein in the
case of a stator blade a heat shield, which is fastened on the
rotor, lies opposite the blade tip. According to a preferred
embodiment, the blade is a rotor blade which rotates around the
axis, whereas the heat shield is installed on the stator of the gas
turbine in a fixed manner.
[0015] In another embodiment of the invention, the thermal barrier
coating is a porous ceramic coating, in particular comprising YSZ.
In this case, the porosity of the thermal barrier coating is
preferably more than 20%.
[0016] An adhesion coating, particularly comprises MCrAlY, is
advantageously arranged between the heat shield and the thermal
barrier coating.
[0017] The metallic cover coating preferably comprises MCrAlY.
[0018] In a further embodiment, the rotor blade is part of the
first rotor-blade row in the turbine section of the gas
turbine.
DETAILED DESCRIPTION
[0019] In FIG. 3, a preferred exemplary embodiment for a blade
arrangement 30 according to the invention is reproduced. In the
example, a heat shield 12, with a clearance 25, again lies opposite
a rotor blade 11 which has the blade tip 27 and is rotatable around
the axis 16 of the gas turbine. The clearance 25, and consequently
the efficiency of the turbine, are optimized by the blade tip 27
cutting into the coating 22, 23 of the heat shield 12 (in FIG. 3,
the possible cutting-in region 28 on the heat shield is indicated
by means of a broken line).
[0020] As a coating which is to be abraded during the cutting in,
provision is made on the heat shield 12 for a thermal barrier
coating 23 which is connected to the substrate of the heat shield
12 via an adhesion coating 22 which lies in between. As an adhesion
coating 22, provision may customarily be made for a metallic,
anti-oxidation coating comprising MCrAlY.
[0021] In trials, it has now been proved that cutting of the blade
tip into the thermal barrier coating 23 is possible even without a
special abrasive coating on the blade tip 27 and leads to good
results if the thermal barrier coating 23--without losing its
thermal properties--is to be slightly abraded to an adequate
degree. This can be achieved by a porous thermal barrier coating 23
being used.
[0022] In this case, a porous ceramic coating, which in particular
may comprise YSZ (yttrium oxide stabilized zirconium), is
especially suitable as a thermal barrier coating 23, wherein the
porosity is created for example by means of embedded polymers which
are subsequently heated. It has been proved to be advantageous in
this case if the porosity of the thermal barrier coating is more
than 20%, that is to say lies within the range of 22-24%, for
example.
[0023] In the case of such a porous thermal barrier coating 23, the
abrasion on the blade tip 27 during cutting in, in relation to the
depth of the cutting-in region 28, is comparatively small so that a
special abrasive coating on the blade tip 27 can be dispensed with.
It suffices, therefore, if the blade tip 27 is covered with a
homogenous cover coating 24 (without abrasive bodies) comprising
MCrAlY, which is provided anyway as a protective coating against
oxidation of the blade material.
[0024] In this way, special provisions do not need to be made on
the blade 11 for cutting in, as a result of which, production of
the blade 11 is substantially simplified.
List of Designations
[0025] 10, 20, 30 Blade arrangement (gas turbine) [0026] 11 Rotor
blade [0027] 12 Heat shield [0028] 14 Hot gas passage [0029] 14,
15, 24 Cover coating [0030] 16 Axis [0031] 17, 22 Adhesion coating
[0032] 18, 23 Thermal barrier coating (TBC) [0033] 19 Carrier layer
[0034] 21 Abrasive bodies [0035] 25 Clearance [0036] 26 Wall (hot
gas passage) [0037] 27 Blade tip [0038] 28 Cutting-in region
* * * * *