WO2019185713A1

WO2019185713A1 - Turbine rotor blade for a gas turbine

Info

Publication number: WO2019185713A1
Application number: PCT/EP2019/057718
Authority: WO
Inventors: Johannes Döhnert; Torsten JOKISCH; Francis Ladru; Thorsten Schulz
Original assignee: Siemens Aktiengesellschaft
Priority date: 2018-03-29
Filing date: 2019-03-27
Publication date: 2019-10-03
Also published as: EP3546703A1

Abstract

The invention relates to a turbine rotor blade (10) for a gas turbine, comprising a metal blade airfoil (16), which has a leading edge (18), against which a hot gas (S) can flow and from which a side wall (22) on the suction side and a side wall (24) on the pressure side extend to a trailing edge (20) of the blade airfoil (16), wherein the blade airfoil (16) extends, in a radial direction (R) extending perpendicularly thereto, from an end (26) on the root side to an end (28) on the tip side, which end forms an airfoil tip (30); and wherein a first cutting layer (36) is applied to the surface (35) of the airfoil tip (30) facing radially outward. The aim of the invention is to provide a turbine rotor blade (10) that has an improved service life, in particular when used in a gas turbine having an axially movable rotor. This aim is achieved, according to the invention, in that a further cutting layer (36) is applied on the tip side at least to one of the two side walls, in particular to the side wall (22) of the blade airfoil (16) on the suction side and/or to the leading edge (18) of the blade airfoil.

Description

description

Turbine blade for a gas turbine

The invention relates to a turbine blade for a gas turbine according to the preamble of claim 1.

A turbine blade according to the preamble of claim 1 is known, for example, from the document EP 0 166 676 A2 or from the document EP 2 316 988 A1. The previously known turbine blade has a metallic blade leaf on whose outwardly facing tip a layer was applied with hard particles. This is when using the turbine blade in a gas turbine for targeted removal of material, which was applied to the tip of the blade opposite housing wall. Usually, these flow path limiting Ge housing walls in stationary gas turbines are formed by so-called ring segments on which the abradable material is applied not only as a temperature protection, but also functional for radio aspalteinstellung and optimization. Background is a during the operation of the gas turbine a stellende thermal distribution, which lead to different thermally induced strains in the housing. Due to the different heat distribution and the different ge stalteten housing components creates a not exactly kreisrun de flow contour of the housing walls in which rotates the stocked with the turbine blades rotor. The hydro-dynamic bearing also leads to the fact that the rotor is not positioned exactly centered. As a result, an individual radial gap is established between each blade tip and the housing wall opposite it. The se can be distributed along the circumference of different sizes. To increase the turbine efficiency these differences are to be set to uniform and smallest possible radial column, which can be achieved with the help of the abradable material. This is called a ceramic, abradable Layer - also called inlet layer - applied to the ring segments and allows a rubbing and digging of the blades in it.

Thus, for example, the aforementioned EP 0 166 676 A2 proposes to use on the outwardly facing surface of the Schaufelspit ze a particularly hard material, thus the softer, applied to the ring segments material during operation of the gas turbine for setting minimum column and also rub there to improve the warm start behavior.

In addition, it is known in the art, the blade blade of turbine blades by means of a layer system of mostly two layers, a first applied Haftver middle layer and a ceramic heat insulating layer applied above, to protect against the harmful influence of Heißga ses.

However, it has been shown that for burying in a rubbed off layer used blades Kings tend to lose at its outer end of the thermal insulation layer locally. This reduces the protection of the blade against hot gas influences, so that at the points missing the thermal barrier coatings, the metallic material of the turbine blade can age prematurely. This can shorten the life of the turbine blade, which in the worst case can result in premature replacement of the damaged component.

In addition, it is known from EP 1 820 938 A1 to equip running blade tips with cleaning elements in order to be able to remove impurities occurring during operation on the ring segments.

The object of the invention is therefore to provide a tur binenlaufschaufel for a gas turbine, which, designed for burying in a housing-side provided abradable layer having a comparatively long life.

The task directed to the turbine blade is achieved with the features of claim 1. Thus, to solve the problem, a turbine blade for a gas turbine, comprising a metallic airfoil comprising one of a Heißgasanströmbare leading edge from which a suction-side side wall and a pressure-side side wall to a trailing edge of the airfoil, wherein the show felblatt in a to Transversely extending radial direction from a foot-side end to form a blade tip forming the tip-side end and wherein at the radially outwardly facing surface of the blade tip a first

Cutting layer is applied, proposed, in which on at least one of the two side walls, in particular on the suction-side side wall of the airfoil and / or on the leading edge of the airfoil on the tip side at least one further cutting layer is applied.

The invention is based on the finding that the first cutting layer has a detectable in the radial direction of thickness, which is relatively low. This can lead to the fact that during one of the first rubbing operations the respective turbine blade during its rotation digs into the material of the housing wall to be abraded so that a contact between the applied on the side walls or on the front edge Wär medämmschicht and the abzureibenden material of the housing wall occurs. In this case, meet two relatively soft materials to each other, so that not necessarily abzurei Bende material of the housing wall is rubbed off. Rather, the also on the tip-side ends of the airfoil on placed thermal barrier coating due to the unwanted contact locally flake. Furthermore, the first on the ra dial outwardly facing surface of the blade tip applied cutting layer when digging into the opposite applied on the housing wall abradable layer significantly heat. The heating leads to thermal strains in the cutting layer and also in the adjacent areas. In particular, the heat-insulating layer attached on the side walls and on the front edge of the blade can be heated to an impermissibly high level, so that the risk of spalling of the heat-insulating layer on the tip side is increased.

In order to avoid these effects, the invention proposes that not only on the radially outwardly facing surface of the blade tip, a first cutting layer is applied, son NEN that at least on one of the two side walls, in particular special on the suction side, because leading side wall of the airfoil and / or on the leading edge of the blade at the tip end at least one further cutting layer is applied. In other words, the first cutting layer and the at least one further cutting layer go into each other and surround the edge between the side wall and the outwardly facing surface. The terms "first" and "further" cutting layer serve only to identify two or more contiguous areas of the cutting layer, which applied in a manufacturing step who can. In this regard, the tip side means that the further cutting layer extends radially inward from the radially outermost edge of the turbine blade in a small, but for the purpose of sufficient circumference.

The cutting layer, usually enriched with hard particles, in particular cubic or hexagonal boron nitride, thus have a sufficient strength and hardness, which makes it possible that only the material applied to the housing wall material is removed there. The applied on the Be tenwänden the airfoil thermal barrier coating is thus protected from destructive contact with the coated housing wall.

More preferably, a mixture of cubic Borni trid (cBN) and MAX phase powder is introduced as a cutting layer, so that the abrasion resistance of Turbine blade significantly improved. The cubic boron nitride has a very high hardness and is characterized by an outstanding wear resistance. The cubic boron nitride embedding material should also have good wear resistance combined with good oxidation resistance and high thermal resistance. This combination of properties is optimally covered by the MAX phases. MAX phases are materials whose properties lie between those of ceramics and metals. Where M is an element of the transition metals (eg Ti, Cr, Nb, Ta, etc.), A is an element of the A group (eg Al, Si, etc.) and X is carbon or nitrogen ( C or N). Typical representatives of the MAX phases are Ti2AlC, Cr2AlC and Nb2AlC. The mixture of cubic boron nitride and MAX phase could, for example, be powder-applied by laser welding. Alternatively, you could sinter sintered preforms, which are then connected to the turbine blade via a soldering process.

By maintaining blade protection due to the continued existence of the ceramic thermal barrier coating on the sidewalls of the airfoil, the life of the turbine blade can be improved as compared to prior art turbine blades.

Advantageous embodiments are given in the dependent claims, wherein the features are combinable with each other any cash.

According to a first advantageous development of the invention, the first cutting layer and the at least one further cutting layer merge into one another, so that an edge arranged between the side wall or the front edge and the radially outwardly facing surface is covered by it. More preferably, the blade tip comprises at least one, preferably at least two squealer edges, each having a first, the suction side facing side surface on which side surface, the further cutting layer is applied. With out the invention is applicable not only only in flat and so with unprofiled blade tips, but also in those blade tips in which the turbine blade be known used have squeal edges. Since the suction-side side wall of the pressure-side side wall leads in the case of rotating turbine blades, it is particularly advantageous if surfaces facing the suction side have the further cutting layer on the tip side.

In order to provide a constant thickness of coatings on the side wall of the airfoil, it is advantageous if the respective side wall of the airfoil or the respective side surface of the squeal edge has a recessed step towards the blade tip. The depth of the step corresponds approximately to the thickness of the applied further cutting layer.

It has been found to be advantageous if at least one further layer, in particular a bonding agent layer and / or a ceramic heat-insulating layer, are provided at the tip end between the further cutting layer and the metallic airfoil. In the event that due to the lack of fatigue strength of the hard particles these prematurely age al or due to higher operating temperatures dissolve, it is advantageous if between the other cutting layer still at least one primer layer, which at the same time in the form of an MCrAlY at the same time represents a corrosion protection layer, this area the turbine blade additionally protects against the harmful effects of the hot gas.

Be particularly advantageous, the use of a vorge turbine blade in a gas turbine Issuer has, whose rotor is slidably ausgestal tet in the axial direction. As is known, gas turbines with axial sliding rotor to felspitzen the radial gaps between the rotor blade and a conically expanding turbine hot gas during stationary operation of the gas turbine to reduce. The blade tips of the turbine blade blades and the conically expanding turbine hot gas channel are correspondingly inclined, so that at a Axialver shift of the rotor in the direction of the incoming Heißgasströ tion of the radial gap between the conical housing wall and just inclined blade tip is reduced. In particular for gas turbines configured in this way, it is advantageous if their turbine blades provided for burying in a coating arranged on the housing side have on the leading edge in the area of the blade tip and / or thus also on the side walls of the blade tip the further cutting layer. Thus, a sharp-sided flaking of the ceramic protective layer on the front edge can be avoided if during the axial displacement of the rotor in the upstream direction, contact between the leading edge of the blade and the abradable coating Be the housing wall occurs.

Further advantages and features of the invention will become apparent from the following description with reference to FIGS. Showing:

1 is a perspective view of a turbine blade according to a first Ausführungsbei game,

Figure 2 shows the longitudinal section through the blade tip of the Turbi nenschaufel according to Figure 2 and

Figure 3 in longitudinal section an alternative blade tip ge according to a second embodiment.

FIG. 1 shows a perspective view of a turbine blade 10. Along an axis which coincides with the radial direction R of a gas turbine in which the turbine blade 10 is installed, it comprises a blade root 12, a platform 14 and an airfoil 16 seen from the inside outwards Such turbine blades 10 are cooled internally. So who made this usually in a casting process, so that they are made of a total of a metallic base material ge. For the invention, however, it is sufficient if the airfoil 16 is metallic. This has a vorströmbare of a hot gas S leading edge 18 as a trailing edge 20. In between, a suction-side side wall 22 and a pressure-side side wall 24 extend, which extend in the transverse direction and thus in the radial direction R from a foot-side end 26 to a tip-side end 28. The latter also forms the blade tip 30, which does not necessarily have to be flat. According to the first embodiment shown, the blade tip 30 is provided with a border, which is also characterized as a squealer edge 32 be provided. On their radially outwardly facing the surface 35 a not shown in Figure 1 cutting layer is applied, however, which can be seen in Figures 2 and 3 in detail.

Figures 2 and 3 each show in a longitudinal section accordingly the section lines II-II and III-III unterschiedli che embodiments of the invention, in which the cutting layer 36 is not only on the radially outwardly facing FLAG surface 35 of the blade tip 30 is disposed but also on the suction-side side wall 24 of the metallic blade 16 and there on its leading edge 18. This is true for both the suction side and the pressure side arranged squeaking edge 32. The irrelevanten for understanding the inven tion elements of the turbine blade are - as always in schematic representations - not illustrated.

The blade tip 30 is exemplified on both the suction side and the pressure side, each with a squeal edge 32 is equipped, which are not only covered at their radially outwardly facing surfaces 35 of the first cutting layer 36. According to the invention, the cutting layer 36 also extends over a corner 38 onto those side surfaces 34 of the invention. grazing edges 32, which are indicative of the suction side 24 or part of the leading edge 18. This part of the cutting layer 36 is referred to as a further cutting layer. The entire cutting layer thus covers the outermost edge 37 of the turbine blade on both sides.

According to Figure 3, the section line III-III is angeord net such that it extends through the front edge 18 of the blade and the pressure-side side wall 24. In this respect, the cutting layer 36 is preferred according to FIG 3 Schaufelspitzenseitig on the leading edge 18 on the hotgas and extends even slightly up to the pressure side side wall 24. The front edge side region of the hot gas S vorströmbaren surface of the blade 16, on which the further cutting layer 36 is provided, is framed in Figure 1 as area VK. This Ausgestal device is particularly useful when the rotor of the turbine equipped with such a turbine blade 10 Gastur is also displaceable in the axial direction A.

According to the embodiments of Figures 2 and 3, the airfoil 16 is coated on both the suction side and drucksei TIG or at the front edge over a large area first of a primer layer 40, which in turn is covered by a ceramic thermal barrier coating 42. According to the first embodiment, two layers 40, 42 abut directly on the further cutting layer. At the same time, the second exemplary embodiment according to FIG. 3 has a step 44 on the tip side, which recesses relative to the front edge 18 or, if appropriate, relative to the suction side wall 22.

The adhesive layer 40 can - as shown in Figure 3 exemp larisch - also be arranged between the cutting layer 36 egg nerseits and the squeal edge 32 and the blade 16 on the other. In this case, only the ceramic thermal barrier coating 42 extends to the other cutting layer zoom. Thus, the invention relates to a turbine blade 10 for a gas turbine with a metallic airfoil 16 comprising an approachable by a hot gas S leading edge 18, from which a suction-side side wall 22 and a pressure-side side wall 24 felblatts 16 extend to a trailing edge 20 of the show, the Airfoil 16 extends in a transverse thereto extending radial direction R from a foot-side end 26 to a blade tip 30 forming spit zenseitigen end 28 and wherein on the radially outwardly facing surface 35 of the blade tip 30 a first

Cutting layer 36 is applied. In order to provide a provided for abrading a softer layer of a housing wall turbine blade, which has an improved life, especially when used in a gas turbine with an axially displaceable rotor, it is proposed that at least on one of the two side walls, in particular on the suction side wall of the airfoil and / or on the leading edge of the blade on the tip side, a further cutting layer is applied.

Claims

claims

A turbine blade (10) for a gas turbine,

comprising a metallic airfoil (16) comprising a front edge (18), which can be flowed on by a hot gas, from which a suction-side side wall (22) and a pressure-side side wall (24) extend to a trailing edge (20) of the airfoil (16),

wherein the airfoil (16) extends in a transversely extending radial direction from a foot-side end (26) to a tip-side (28) forming a blade tip (30), and

wherein on the radially outwardly facing surface (35) of the blade tip (30) a first cutting layer (36) is applied,

characterized,

that at least one of the two side walls (22, 24), in particular on the suction-side side wall (22) of the blade (16), and / or on the front edge (18) of the blade (16) tip-side at least ei ne further cutting layer (36) is applied.

2. turbine blade (10) according to claim 1,

in which the first cutting layer (36) and the at least one further cutting layer (36) merge into one another, so that an edge arranged between the radially outwardly facing surface (35) and the relevant side wall (22, 24) or the front edge (18) (37) of which is covered.

3. turbine blade according to claim 1 or 2,

in which on the side walls of the airfoil (16) at least one ceramic thermal barrier coating (42) is introduced.

4. turbine blade (10) according to claim 1, 2 or 3, wherein the blade tip (30) at least one, preferably, at least two squealer edges (32) with a first side of the suction side wall (22) facing side surface on which side surface of the further cutting layer (36) is applied.

5. turbine blade (10) according to one of the preceding claims,

in which the respective side wall of the airfoil (16) or the respective side surface of the Anstreifkan te to the blade tip (30) towards a recessed step (44).

6. turbine blade (10) according to one of the preceding claims,

being on the tip side between

the further cutting layer (36) and the metallic blade

at least one more layer,

in particular a bonding agent layer (40) and / or a thermal barrier coating (42) is provided.

7. turbine blade (10) according to one of the preceding claims,

wherein the first cutting layer and the further cutting layer hard particles, in particular of cubic or hexagonal boron nitride, and contain MAX phases.

8. Use of a turbine blade (10) according to one of claims 1 to 7 in a gas turbine with an axially displaceable rotor.