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.

Description

CROSS REFERENCE TO RELATED APPLICATION

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

The present invention relates to the field of gas turbine technology in that it refers to a blade arrangement of a gas turbine.

BACKGROUND

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.

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.

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.

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

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

The invention shall subsequently be explained in more detail based on exemplary embodiments in conjunction with the drawings. In the drawings

FIG. 1 shows in a greatly simplified view an earlier blade arrangement without the possibility of cutting in;

FIG. 2 shows in a view comparable to FIG. 1 another earlier blade arrangement with a special abradable coating on the blade tip, and

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

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.

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.

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.

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%.

An adhesion coating, particularly comprises MCrAlY, is advantageously arranged between the heat shield and the thermal barrier coating.

The metallic cover coating preferably comprises MCrAlY.

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

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).

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.

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.

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.

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.

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

• 10, 20, 30 Blade arrangement (gas turbine)
• 11 Rotor blade
• 12 Heat shield
• 14 Hot gas passage
• 14, 15, 24 Cover coating
• 16 Axis
• 17, 22 Adhesion coating
• 18, 23 Thermal barrier coating (TBC)
• 19 Carrier layer
• 21 Abrasive bodies
• 25 Clearance
• 26 Wall (hot gas passage)
• 27 Blade tip
• 28 Cutting-in region

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.