EP2087149A2

EP2087149A2 - Vane for a compressor or a turbine of an aircraft engine, aircraft engine comprising such a vane, and method for coating a vane of an aircraft engine

Info

Publication number: EP2087149A2
Application number: EP07846269A
Authority: EP
Inventors: Thomas Uihlein; Wolfgang Eichmann; Falko Heutling; Markus Uecker
Original assignee: MTU Aero Engines GmbH
Current assignee: MTU Aero Engines AG
Priority date: 2006-11-03
Filing date: 2007-10-27
Publication date: 2009-08-12
Also published as: DE102006051813A1; US20100143110A1; WO2008055471A3; CA2668298A1; WO2008055471A2

Abstract

The invention relates to an aircraft engine comprising a compressor and at least one turbine. The compressor and the turbine are each provided with vanes (1), each of which has a blade (10) that forms a suction side and a pressure side. At least one first (10) of said blades (10) is coated with a protective layer (16) in order to reduce erosion or wear. The protective layer (16) is applied to at least one side of the blade (10) in such a way that at least two areas (18, 20) are formed that adjoin each other at a boundary line (22). A first area (18) is provided with the protective layer (16) in such a way that the protective layer (16) has a substantially constant first thickness in said first area (18). A second area (20) is free from the protective layer (16) or is provided therewith in such a way that said protective layer (16) has a substantially constant second thickness in said second area (20), the second thickness differing from the first thickness. Said boundary line (22) has at least two points, of which the connecting line differs from the course of the boundary line between said two points.

Description

Blade for a compressor or a turbine of an aircraft engine,

An aircraft engine with such a blade and method for coating a blade of an aircraft engine

The invention relates to an aircraft engine with a compressor and at least one turbine, a blade for a compressor or a turbine of an aircraft engine and a method for coating a blade of an aircraft engine.

From DE 10 2004 001 392 Al a blade of a gas turbine with an airfoil and a blade root is already known, in which the entire blade is provided with a wear protection coating. This wear protection coating is designed there as a multilayer coating system with four different layers.

Layers of any kind often exert a negative influence on the fatigue strength and / or the service life of components. This is especially true for hard coatings against wear or corrosion, with the risk that cracks in the ceramic layers run quickly into the base material and lead to premature failure of the component.

DE 10 2004 001 392 A1 further explains that it is also possible to provide the blade only in sections, namely in the region of the blade or in parts thereof or in the region of the blade root with a wear protection coating.

Such designs are known to the Applicant in that often only the upper third or the radially outer third of a blade blade is coated, as shown by the Applicant Fig. Ia shows, in which a blade 101 with an airfoil 110, a blade root 112th , a platform 114 and an erosion control coating 116 is shown schematically. However, this leads to the risk that a particle flow in the transition area washes the base material, undermines the layer and leads to notching, with premature failure occurring under vibration load. To avoid this, the coating can be completely dispensed with and increased erosion or increased wear can be accepted. Against this background, the invention has for its object to provide compressor or turbine blades of aircraft engines with high corrosion and erosion resistance and good fatigue strength.

It is proposed an aircraft engine according to claim 1. A blade according to the invention is the subject of claim 2. A method according to the invention is the subject of claim 9. Preferred developments are subject of the dependent claims.

According to the invention, in particular, an aircraft engine with at least one compressor and at least one turbine is proposed. The compressor as well as the turbine each have blades, namely compressor blades or turbine blades. The blades each form an airfoil, which - as is usually the case with compressor blades or turbine blades - has a suction side and a pressure side. At least a first of these airfoils is coated for the reduction of erosion or wear with a protective layer which is applied on at least one side of this first airfoil, ie on the pressure side and / or on the suction side, such that at least two, at a boundary line to each other forming abutting regions, of which a first region is provided with the protective layer such that the protective layer has a substantially constant first thickness in this first region, and of which a second region, in particular on the same side as the first region is free of the protective layer or is provided with the protective layer such that this protective layer in this second region has a substantially constant, deviating from the first thickness second thickness.

It is now provided that the boundary line separating the first area from the second area is designed such that at least two points of this boundary line exist, whose connecting line deviates from the boundary line course between these two points or non-coincident with the boundary line course between them is two points. This can for example be such that the corresponding route is such that it does not intersect the boundary line. But it can also be provided that the route cuts the borderline. The first airfoil may be an airfoil of a turbine blade or an airfoil of a compressor blade. It can also be provided that at least one turbine blade and at least one compressor blade is a first blade or designed in accordance with the invention. It can be provided that the first blade is designed according to the invention on its suction side and / or on its pressure side. In an advantageous embodiment, the blades of several, preferably all, turbine and / or compressor blades are configured as first blades or in erfmdungs- according manner.

The respective first and / or second region may be an area that extends to the outer edge of the airfoil, or an area that is essentially closed in itself.

According to claim 2, a blade according to the invention is claimed.

The blade may in particular be a compressor blade or a turbine blade of an engine.

The blade may be configured to form a platform from which the airfoil protrudes. The blade has in particular a blade root. In principle, the invention may, for example, also refer to blisks or the like.

In a particularly advantageous embodiment, it is provided that the blades or the blades formed in accordance with the invention are designed in one piece, and in particular have a blade root and an airfoil. The blade is advantageously made of the same material, apart from the coating.

The blade may in particular be integrally formed, that is to say in particular such that an airfoil and a blade root (and possibly platform) are formed or produced from one part.

In an advantageous embodiment, the blade at least on one side, namely suction side and / or pressure side, exactly two areas of the type mentioned, and thus exactly a borderline. It can also be provided that a plurality of regions of the aforementioned type and accordingly a plurality of boundary lines are provided on the suction side and / or pressure side.

In an advantageous embodiment, the boundary line has curved sections. It can be provided that the boundary line is parabolic-shaped. The blade may, for example, have a blade root, wherein the boundary line is designed parabolic in such a way that it is open in the direction of this blade root.

In a particularly preferable embodiment, it is provided that the position of the boundary line in dependence (in operation in an aircraft engine, in particular presumably, given) maximum Schwingungsspannungeii (in the airfoil) and / or depending on (in operation in an aircraft engine, in particular, given ) Erosionsbelastungen of the airfoil is selected, which are given at the leading and trailing edges of the blade or from the wide area extending on the suction side or pressure side or expected. This can in particular be such that the areas addressed are selected such that the points at which the voltage maxima are given are in an area or respectively in a region in which no protective layer is provided, or a protective layer with a smaller thickness than elsewhere on the same side of the airfoil. In particular, permanent voltages, dynamic voltages and residual stresses can be taken into account.

The anticipated stresses and / or erosion stresses mentioned can be determined, for example, by means of simulation and / or empirical values or in another suitable manner.

It can be provided that on the suction side, two regions are provided, which differ in the thickness of their protective layer, or in that in one of these areas, a protective layer is given, and in the other of these two areas, no protective layer is given for which these two areas separate boundary line:

(for ease of reference, this relationship will be referred to as formula 1)

where:

hi: measured height above the hub intersection of the location of the maximum IF vibration stress at the leading edge

Ji ₂ : measured height above the hub intersection of the location of the maximum IF

Oscillation voltage (= first swing vibration) at the trailing edge h ₃ : measured height above the hub intersection of the location of the maximum IF oscillation voltage on the suction side

L: total lattice length or axial position of the trailing edge in the middle of the duct with reference to the leading edge in the center of the duct

L ₃ : Axial position of the location of the maximum 1F vibration stress on the suction side relative to the leading edge.

Alternatively or additionally, it can also be provided that two regions are provided on the pressure side which differ in the thickness of their protective layer or in that one protective layer is provided in one of these regions and in the other of these two regions no protective layer is given, for which these two areas are separated by a borderline:

0.88X (UxA ₁ -A ₂ ) 0.22x (V + Λ) _v

(for simplicity of reference, this relationship will be referred to as Formula 2) With regard to the meaning of the parameters h _ls h ₂ and L, reference is made to the above explanation.

With the formulas 1 and 2, exemplary gradients are given for a respective boundary line in the x-y direction. The formula values or the formulas 1 and 2, however, are only preferred examples. Instead of the factor 1.1, z. B. values between 1.0 and 1.5 are used. The formula or formulas 1 and 2 for the parabola or parabolas can also be extended by the coordinate z if required.

Formulas 1 and 2 refer in particular to an xy coordinate system in which the origin is located so that - in a platform design - x = 0 at the platform leading edge (on the leading edge) and on the side facing the airfoil the platform is in the region of the front edge of the platform or the leading edge y = 0.

The 1F oscillation voltage is in particular the oscillation voltage of the first bending oscillation. The channel center is in particular substantially the middle between the surface of the platform facing the blade leaf and the housing located radially outside thereof in the radial direction; the channel center of a plurality of blades held in an aircraft engine on the same rotor or disc defines for the arrangement of these substantially a hollow cylindrical shape.

The borderline may also be related to allowable repair areas for apertures or patches. In addition, this may mean that no stripping is necessary in the course of repair work, if the erosion-prone areas coincide with the permissible repair areas and these are mechanically removed anyway with the used layers. Ie. The recoating is then possible without previous stripping.

The layer is preferably a multilayer layer.

According to a method according to the invention, it is provided that voltages, in particular voltage maxima, are determined to which the blade is subjected in a predetermined operation is exposed to a predetermined aircraft engine or will be exposed to the operation, which can be done with respect to the pressure side and / or the suction side. The determination of the voltages or voltage maxima can, for example, be based on empirical values or on the basis of calculations or empirically or in another way.

Before, subsequently or in parallel, an erosion load is determined which the bucket is likely to be exposed to during operation. This can be done, for example, based on empirical values.

It is further provided that areas of the blade of the blade are determined which are not to be coated or compared to other areas of the blade with a reduced layer thickness, these determinations being carried out depending on the determined voltages or the erosion load.

Subsequently, the blade or the blade is coated, taking into account the determination of the areas of the blade of the blade, which are not coated or with reduced layer thickness.

It should be noted that the method according to the invention can be carried out with respect to the pressure side of an airfoil and / or with respect to the suction side of an airfoil.

According to a particularly preferred embodiment is provided in particular that z. B. multilayer coatings with little effect on the fatigue strength are used and / or the layers only in the transition region from the platform to the airfoil, where there is little erosion attack and / or in areas where voltage maxima of the vibration are omitted. It can be carried out in an advantageous embodiment, the following procedure. First, the voltage maxima can be determined. Subsequently, an overlay with an erosion image on a simulation program for particle erosion, such as CFX5 of the company ANSYS done. Subsequently, areas can be determined which should not be coated or less. It can also be provided that subsequently these areas are shaded, which can be done for example by known to those skilled devices or procedures. It should be noted that instead of particle simulation programs, empirically determined (simulation) images of already known components can be used.

In an advantageous embodiment, it is provided that the stripping step before recoating is dispensed with if permissible repair areas coincide with optimized coating areas and the area of the erosion attack.

In the following, embodiments of the invention will now be explained in more detail with reference to FIGS. It shows:

Fig. Ia a known design;

FIG. 1 b shows an exemplary design according to the invention in a schematic view; FIG.

2 shows a blade in a schematic view with schematically and exemplarily added zones of different levels of erosion stress; and

Fig. 3 shows an exemplary inventive blade from its suction side.

1 a shows a modified design of a blade 1 of an aircraft engine, wherein the design already mentioned in the introduction according to FIG. 1 a has a conventional coating surface and the design according to FIG. 1 b is an exemplary design according to the invention.

As FIG. 1 b shows, the blade 1 has an airfoil 10, a blade root 12 partially shown in this figure, and a platform 14. The platform 14 separates the airfoil 10 from the blade root 12. The airfoil 10 has a coating 16 on its pressure side and / or on its suction side.

The airfoil 10 has a coating 16 on its suction side and / or pressure side. At least a first region 18 and a second region 20 are formed on the addressed suction side and / or pressure side of this airfoil 10, with this first region and this second region abutting one another at a boundary line 22. In the exemplary embodiment according to FIG. 1 b it is provided that in the first region 18 the already mentioned coating or protective layer 16 is provided, and the second region 20 is free of such a protective layer. Alternatively, however, it may also be provided, for example, that the first region 18 and the second region 20 each have a protective layer, wherein these two protective layers or regions 18, 20 differ by the thickness of their protective layer. This may in particular be such that in the first region 18 the protective layer or coating is thicker than in the second region 20.

In contrast to the design according to FIG. 1 a, in which the local boundary line 122 between the first 118 and second region 120 lies completely on a straight line, at the boundary line 22 according to FIG. 1b at least two points of this boundary line 22, whose connecting distance from deviates from the boundary line between the two points.

In other words, the boundary line 22 according to FIG. 1b is not located completely on a straight line.

The boundary line 22 according to FIG. 1b is curved here, in particular parabolic, designed. As clearly shown in FIG. 1b, the curvature is concave there or the parabolic shape in the direction of the blade root 12 or the platform 14 is open.

The parabolic shape here can have a profile according to the formula 1 or according to the formula 2.

Compared to the design according to FIG. 1 a, the erosion attack in the design according to FIG. 1 b is reduced, in particular due to the formation of the protective layer or coating surface there.

FIG. 2 schematically shows a blade as well as an exemplary erosion load over the blade length or blade height. This erosion stress can be determined, for example, from a particle simulation program or empirical experience or the like. The exemplary erosion load shown in FIG. 2 is such that the airfoil 10 is subjected to little to no erosion stress in an area 80 located near the bucket 12, and, if necessary, stepped with increasing distance from the blade root 12 - Is exposed to an increasing erosion load, which can be roughly divided into a region 82 with average erosion load and a region 84 with high to very high erosion load.

Fig. 3 shows schematically an exemplary inventive airfoil 10, in view of the suction side.

Shown schematically on the blade 10 in FIG. 3 is a tension profile which can presumably occur during operation of the blade or of the blade 10 in a compressor or a turbine of an aircraft engine. The stress profile can be determined, for example, empirically or from empirical values, or can be calculated or determined in another suitable manner.

The reference numeral 24 indicates the maximum 1F oscillation voltage at the leading edge. The reference numeral 26 indicates the maximum 1F oscillation voltage at the trailing edge and the reference numeral 28 indicates the maximum 1F oscillation voltage on the suction side.

As can be seen from the view on the suction side of the airfoil 10 in FIG. 3, a coating 16 is provided in the radially outer region or in the first region 18. The second region 20 is uncoated or slightly coated or provided with a thinner coating than the first region 18. The first region 18 is separated from the second region 20 by a boundary line 22 or abut the regions 18 and 20 at this boundary line 22, wherein the boundary line 22 having the parameters indicated in the legend in Fig. 3 has a course which is formed according to the formula 1 given above.

On the pressure side, not shown in Fig. 3, also such a boundary line separating a coated from an uncoated or little coated area may be provided, this limit preferably being in accordance with the formula 2 above is specified runs. Also on the pressure side is then preferably the coated or more coated area radially outward.

The zero point of the graph or of the corresponding coordinate system may lie at the point indicated by IDLE in FIG. 3 or at the point mentioned above.

In particular, blades can be loaded by bending and torsional modes under stresses such as pumps or flutter with the highest intensities. The maxima of these modes (critical stress peaks) are often located in the lower half of the sheet. In these areas, no layer boundary should run or it is expedient if there is no layer boundary in these areas, or at least the layer thickness is reduced or it is expedient if the layer thickness is at least reduced.

There are often higher voltages on the suction side than on the pressure side; d. H. If noticeable wear also occurs in the suction area, it may be necessary to differentiate between the suction and pressure sides in the case of local coating, but this can also be different.

As described above, it is possible to describe, as it were, a boundary line (in particular a parabola or a parabola) which distinguishes areas which can be coated without and with restriction, as is the case, for example, in an advantageous embodiment of the invention.

Referring to Fig. 3, the following legend applies:

Legend:

Zero point for the graphs is IDLE (Inner Diameter Leading Edge see Fig. 3)

hl: Measured height over hub cut (see picture) of the location of the maximum IF swing voltage at the leading edge h2: Gem. height above hub intersection of the location of the maximum 1F vibration stress at the trailing edge

h3: Gem. Height above hub section of the location of the maximum 1 F swing voltage on the suction side

L: Overall lattice length or axial position of the trailing edge in the center of the duct relative to the front edge in the center of the duct

L3: Axial position of the location of the maximum 1 F swing voltage on the suction side relative to the leading edge

Claims

claims

An aircraft engine having a compressor and at least one turbine, said compressor and turbine each having blades (1) each provided with an airfoil (10) forming a suction side and a pressure side, at least a first (10) said blade (10) is coated with a protective layer (16) on at least one side of the blade (10) to reduce erosion or wear such that at least two portions (18, 18) abutting one another at a boundary line (22) , 20), of which a first region (18) is provided with the protective layer (16) such that the protective layer (16) has a substantially constant first thickness in this first region (18), of which a second Is free of the protective layer (16) or provided with the protective layer (16) such that this protective layer (16) in this second region (20) is a substantially constant, has a second thickness deviating from the first thickness, characterized in that at least two points of this boundary line (22) exist whose connecting path deviates from the boundary line course between these two points.

A bucket for a compressor or turbine of an aircraft engine, said bucket having a suction side and a pressure side forming airfoil (10) coated to reduce erosion or wear with a protective layer (16) disposed on at least one side of the airfoil (10) is applied in such a way that at least two regions (18, 20) abutting on a boundary line (22) are formed, of which a first region (18) is provided with a protective layer (16) such that the Protective layer (16) in this first region (18) has a substantially constant first thickness, and of which a second region (20) is free of a protective layer (16) or provided with the protective layer (16) such that the protective layer (16) in this second region (20) has a substantially constant, deviating from the first thickness second thickness, in particular for an aircraft engine according to claim 1, characterized gekennzei chnet that At least two points of this boundary line (22) exist whose connection distance deviates from the boundary line course between these two points.

3. A blade according to claim 2, characterized in that the boundary line (22) is designed parabolic.

4. A blade according to claim 3, characterized in that the boundary line (22) has a blade root (12), and the boundary line (22) is designed parabolic in such a way that the parabolic shape in the direction of the blade root (12) is open.

5. A bucket according to any one of claims 2 to 4, characterized in that the boundary line (22) in response to the occurring during operation of the blade in an aircraft engine maximum 1F oscillation voltage (oscillation voltage of the first bending vibration) at the trailing edge and in dependence on the operation the blade in an aircraft engine maximum 1F vibration stress at the leading edge and depending on whether the boundary line (22) surface areas of the suction side or the pressure side of the blade from each other - depending on the occurring during operation of the blade in an aircraft engine maximum 1F- Oscillation voltage on this page, ie Suction side or pressure side, runs such that these three or three maximum 1F voltages are located on the same side of the boundary line (22).

6. blade according to one of claims 2 to 5, characterized in that on the suction side at least one or exactly one boundary line (22) is given, for which at least one or exactly one boundary line (22) applies:

_{Λ Λ} . I ₅ Ix (Z ² X (Zz ₃ - Zz ₁ ) - ^ X (A ₂ - A ₁ ))

where: hi: measured height above the hub intersection of the location of the maximum IF vibration stress at the leading edge h ₂ : measured altitude above the hub intersection of the location of the maximum IF vibration stress at the trailing edge h ₃ : measured altitude above the hub intersection of the location of the maximum IF vibration stress on the suction side

7. A blade according to any one of claims 2 to 6, characterized in that on the pressure side at least one or exactly one boundary line (22) is given, for which at least one or exactly one boundary line (22) applies:

y

where: hi: measured height above the hub intersection of the location of the maximum IF vibration stress at the leading edge h ₂ : measured altitude above the hub intersection of the location of the maximum IF vibration stress at the trailing edge

8. aircraft engine according to claim 1, characterized in that at least one of the blades is designed according to one of claims 2 to 7.

9. A method for coating a blade of an aircraft engine, in particular turbine or compressor blade, comprising the steps of:

Determining voltages, in particular maximum stresses, to which the blade is subjected in a predetermined operation in a predetermined aircraft engine or is expected to be exposed during operation;

- Determine the erosion load that the bucket is likely to encounter during operation;

- Determining portions of the blade (10) of the blade, which are not or compared to other areas of the airfoil (10) to be coated with a reduced thickness, this determination is carried out in dependence of the determined voltages and the erosion load; and

- Coating the blade, and in particular of the blade (10) taking into account the determination of the portions of the blade (10) of the blade, which are not to be coated or with reduced layer thickness.