BLADE ROW, BLADE AND TURBOMACHINE

Abstract

A blade row for a turbomachine is disclosed. The blade row has an inner lateral wall and an outer lateral wall for bordering a hot gas channel in the turbomachine through which hot gas passes. At least one of the lateral walls has a rounded lateral wall front edge, which is provided with a circumferentially asymmetrical edge contouring that is elliptical or is comprised of one or more segments of a circle.

Description

This application claims the priority of European Patent Application No. EP 13169058.8, filed May 24, 2013, the disclosure of which is expressly incorporated by reference herein.

BACKGROUND AND SUMMARY OF THE INVENTION

The invention relates to a blade row of a turbomachine, a blade and a turbomachine.

To minimize secondary flows, blade rows in turbomachines, such as gas turbines and aircraft engines in particular, are often equipped with a lateral wall contouring on the housing side and/or on the hub side bordering a hot gas channel of the turbomachine through which a hot gas flows. A lateral wall front edge is traditionally designed with circumferential symmetry in order to maintain an axial gap from an upstream blade row and therefore it is not uncontoured. U.S. Patent Application Publication 2010/0172749 A1, however, describes a blade row having a circumferentially asymmetrical lateral wall front edge, provided with wave-shaped edge contouring, which is comprised of a variety of rounded or adjusted surfaces in the direction of flow of the hot gas and/or the axial direction of the turbomachine in general.

The object of the invention is to create a blade row of a turbomachine that has an alternative lateral wall front edge contouring. In addition, another object of the present invention is to create a blade for such a blade row and a turbomachine having such a blade row.

A blade row of a turbomachine according to the invention has an internal lateral wall and an external lateral wall for bordering a hot gas channel of the turbomachine through which a hot gas passes. At least one of the lateral walls has a rounded lateral wall front edge, which is provided with a circumferentially asymmetrical edge contouring, which is elliptical according to the invention or is comprised of one or more segments of a circle.

The elliptical design of the edge contouring and/or the formation of the circumferentially asymmetrical edge contouring from a plurality of segments of a circle, as seen in the direction of flow of the hot gas and/or in the axial direction of the turbomachine, causes a reduction in secondary flows in the area of the lateral wall front edge in comparison with a blade row having a circumferentially symmetrical lateral wall front edge. In addition, the variation in a static pressure in the circumferential direction in the hot gas channel and in the internal cooling channel is reduced due to the edge contouring on an internal lateral wall front edge and/or a hub-side lateral wall front edge, thus leading to a reduction in local influx and efflux intensities, thereby reducing the mixing losses between the hot gas and the cool air. Furthermore, a leakage flow with its flow structures can be coordinated with a secondary flow in the blade row in a targeted manner.

In a preferred exemplary embodiment, the edge contouring is wave-shaped, as seen in the circumferential direction. The wave contour may be formed by two elevations and a recess situated between the elevations, for example, such that the elevations extend over the lateral edges of the blades, considered respectively, so that two elevation sections are formed on each blade in its lateral edge area, each section together with a corresponding elevation section on the neighboring blade forming one elevation. The elevations and recesses each relate to an ideal circumferentially symmetrical lateral wall front edge and yield a further reduction in secondary flows.

An oncoming flow against the root side of the turbine blades of the blade row can be improved if root fillets of the turbine blades are inserted into the edge contouring.

The secondary flows can be reduced in the direction of flow over the entire lateral wall if the lateral wall is also contoured, and in particular if the lateral wall contouring is then integrated into the edge contouring.

To reduce the structural complexity and manufacturing complexity, one exemplary embodiment provides for the lateral wall to be designed without contour, such that the edge contouring tapers out into it.

To minimize a front axial gap despite the edge contouring, a front lateral wall overhang is preferably free of contour and thus is circumferentially symmetrical.

To reduce the structural complexity and manufacturing complexity, a rear lateral wall overhang may be free of contour and thus circumferentially symmetrical.

A blade according to the invention has a platform, the platform front edge of which is rounded and has an edge contouring, which is elliptical or is comprised of several segments of circles. A number of such blades make it possible to form a blade row of a turbomachine according to the invention, which makes it possible to reduce secondary flows in the hot gas channel on the lateral wall end, static pressure differences and mixing losses. Furthermore, leakage flows can be passed through the blade row in a targeted manner.

A turbomachine according to the invention has at least one blade row according to the invention. A turbomachine, such as a gas turbine and an aircraft engine in particular, is characterized by a superior efficiency, because it permits reductions in lateral wall-side secondary flows in the hot channel, static pressure differences and mixing losses. Furthermore, leakage flows can be passed through the blade row in a targeted manner. The at least one blade row is preferably situated on the turbine end and in a low-pressure turbine of the turbomachine, for example.

Preferred exemplary embodiments of the invention are explained in greater detail below on the basis of schematic diagrams.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a perspective diagram of a partial region of a blade row according to the invention with a first exemplary embodiment of a blade according to the invention;

FIG. 2 shows a circumferential section of a blade row according to the invention with a second exemplary embodiment of the blade according to the invention;

FIG. 3 shows a side view of a third exemplary embodiment of the blade according to the invention;

FIG. 4 shows a circumferential section of a blade row according to the invention with a fourth exemplary embodiment of the blade according to the invention;

FIG. 5 shows a perspective diagram of a fifth exemplary embodiment of the blade according to the invention; and

FIG. 6 shows a side view of the fifth exemplary embodiment of the blade according to the invention from FIG. 5.

DETAILED DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a perspective diagram of a subsection of a blade row 1 of a turbomachine according to the invention, with a first exemplary embodiment of the blades 2, 4 according to the invention. The turbomachine is a gas turbine and is an aircraft engine in particular. The blade row 1 is preferably situated on the turbine end, for example, in a low-pressure turbine of the turbomachine. In the exemplary embodiment shown here, the blades 2, 4 are rotor blades in particular, but they may also be guide blades, as illustrated in FIGS. 3 and 4 in particular. The blades 2, 4 are turbine blades in particular.

The blade row 1 has a hub-side lateral wall 6 and a housing-side lateral wall, radially bordering a hot gas channel with a hot gas flowing through it essentially in the axial direction of the turbomachine. For reasons of simplicity, only the hub-side and/or internal lateral wall 6 is/are shown in FIG. 1 and the following FIGS. 2 to 6.

The hub-side lateral wall 6 is formed by internal and/or hub-side platforms 8, 10 of the blades 2, 4, which are numbered in FIG. 2 as an example and are positioned side by side in the circumferential direction and from which turbine blades 12, 14 extend into the hot gas channel. The hub-side lateral wall 6 has a lateral wall front edge 16 and a lateral wall rear edge 18, as seen in the direction of flow of the hot gas.

According to FIG. 1, the lateral wall front edge 16 forms a transitional region between the lateral wall 6 and a lateral wall overhang 20, which is positioned in the front and/or upstream on the inside radially to the hub-side lateral wall 6. In the exemplary embodiment shown in FIG. 1, the lateral wall rear edge 18 of the lateral wall 6 forms a transitional region to a lateral wall overhang 22, which is positioned at the rear and/or downstream, on the inside radially to the front lateral wall overhang 20.

As shown in FIG. 2, according to the formation of the lateral wall 6, the lateral wall front edge 16, the lateral wall rear edge 18, the front lateral wall overhang 20 and the rear lateral wall overhang 22 are formed by individual platform front edges 24, 26, platform rear edges 28, 30, front platform overhangs 32, 34 and/or rear platform overhangs 36, 37 positioned side by side.

As shown in FIG. 1, the lateral wall front edge 16 is designed to be rounded and is provided with edge contouring. The edge contouring is such that the lateral wall front edge 16 is circumferentially asymmetrical. The edge contouring is preferably uniformly wave-shaped, based on an ideal, circumferentially symmetrical lateral wall front edge, wherein an elevation 38, 40 extending onto both platforms 8, 10 and a recess 42 provided between the elevations 38, 40 are both arranged in the lateral edge area of the blades 2, 4. Thus two corresponding elevation sections 44, 46, which, taken separately and isolated from the lateral wall front edge 16, form the elevations 38, 40 are situated in each lateral edge region of the blades 2, 4. The turbine blades 12, 14 are guided into one of the elevation sections 46, each with an oncoming flow edge region 48 of their rounded root in the exemplary embodiment shown here.

The recess 42 is formed between the elevation sections 44, 46 and is preferably situated centrally with the respective platform front edge 24, 26 (see FIG. 2). The recess 42 is offset radially toward the inside in comparison with an ideal lateral wall edge line and/or hot gas channel line, so that in its lowest section, it comes in contact with a foot 50 on the outside radially of the front lateral wall overhang 20. The front lateral wall overhang 20, which has circumferential symmetry, thus develops directly into the recess 42 with its outer foot 50 in at least some sections. However, the elevations 38, 40 and/or elevation sections 44, 46 are at a distance from the outer foot 50, such that free areas 52, 54 extending radially are not contoured between the outer foot 50 and the elevations 38, 40 and/or elevation sections 44, 46.

In the axial direction of the turbomachine, and thus in simplified terms, in the direction of flow and/or as seen from the front lateral wall overhang 20 to the lateral wall 6, the edge contouring is elliptical or is comprised of a plurality of segments of circles with different radii.

The edge contouring develops into a circumferentially asymmetrical lateral wall 6 and/or a circumferentially asymmetrical lateral wall contouring. In the exemplary embodiment shown in FIG. 1, the lateral wall contouring has two peaks, which extend over the lateral edge area of the blades 2, 4 to the neighboring platform 8, 10. These peaks are arranged one after the other in succession, and each is shown only as a peak section 56, 58 due to the perspective. The front peak is at a distance from a suction-side area 60 of the root fillet by way of a channel, which has not been assigned a number. The rear peak extends from an uncontoured lateral wall section 62.

The lateral wall contouring tapers out in the direction of the lateral wall rear edge 18. The lateral wall rear edge 18 is not contoured and therefore has circumferential symmetry. Likewise the rear lateral wall overhang 22 is not contoured and therefore has circumferential symmetry.

FIG. 2 shows a hub-side circumferential section of a turbine blade 1, which is formed from a plurality of blades 2, 4 according to the invention, such as turbine rotor blades according to a second exemplary embodiment. The blade row 1 has a circumferentially symmetrical front lateral wall overhang 20, a circumferentially asymmetrical lateral wall front edge 16, a circumferentially asymmetrical lateral wall 6, a circumferentially symmetrical lateral wall rear edge 18 and a circumferentially symmetrical rear lateral wall overhang 22.

The circumferentially asymmetrical lateral wall front edge 16 is rounded and has a wave-shaped edge contouring in the circumferential direction, which appears elliptical in the direction of flow and/or as seen from the front lateral wall overhang 20 to the lateral wall 6 or is comprised of several segments of circles with different radii. As in the exemplary embodiment according to FIG. 1, the edge contouring develops into a circumferentially asymmetrical lateral wall contouring of the lateral wall 6.

The lateral wall contouring of the lateral wall 6 according to the exemplary embodiment shown in FIG. 2 has two peaks 64, 66 near a suction side 67 of each turbine blade 12, 14. The peaks 64, 66 are spaced a distance apart from one another in the direction of flow and each extends beyond the lateral edge regions of platforms 8, 10 of the neighboring blades 2, 4. A region of the lateral wall 6 and/or of the platforms 8, 10 not provided with peaks 64, 66 is provided with a valley 68 extending beyond the lateral edge regions. Thus each platform has two suction-side peak sections 70, 72 and two more-or-less pressure-side peak sections 74, 76, which together with the corresponding peak sections of the neighboring blades 2, 4, form the peaks 64, 66. The peaks 64, 66 are each spaced a distance apart from a suction-side region 60 of a root fillet through a channel. Furthermore, the root fillet is guided onto the respective neighboring platform 8, 10 on the suction side.

FIG. 3 shows a section through an exemplary embodiment of the blade 2 according to the invention, which is designed as a guide blade, such as a turbine guide blade of a turbomachine. In significant contrast to the exemplary embodiments illustrated in FIGS. 1 and 2, this blade 2 does not have any front lateral wall overhang or any rear lateral wall overhang.

However, the blade 2 shown in FIG. 3, like the preceding exemplary embodiments, does have a hub-side circumferentially asymmetrical platform front edge 24, a hub-side circumferentially asymmetrical platform 8 and a hub-side circumferentially symmetrical platform rear edge 28.

The circumferentially asymmetrical platform front edge 24 is rounded and has a wave-shaped edge contouring, as seen in the circumferential direction, which appears elliptical, as seen in the direction of flow, or is comprised of multiple segments of circles with different radii. As in the exemplary embodiments according to FIGS. 1 and 2, the edge contouring develops into the platform 8 and/or a platform contouring.

With its platform contouring, the platform 8 forms a partial region of a circumferentially asymmetrical lateral wall contouring. Starting from an ideal hot gas channel line 78, the platform contouring is such that the lateral wall contouring has a front peak 64 and a rear peak 66 between two turbine blades 12, the rear peak being designed to be lower in comparison with the front peak 64 and being connected to the front peak 64 by means of a section 80 with a reduced height in relation to the peaks 64, 66. The platform contouring tapers out in the direction of the circumferentially symmetrical platform rear edge 28, which is preferably angular in the exemplary embodiment shown here.

FIG. 4 shows a hub-side circumferential section of a blade row 1 formed from a plurality of the blades 2, 4 according to the invention, such as turbine guide blades according to a fourth exemplary embodiment. According to the preferred embodiment of the blades 2, 4 as turbine guide blades, the blade row 1 does not have a front lateral wall overhang or a rear lateral wall overhang, but it does have a circumferentially asymmetrical lateral wall front edge 16, a circumferentially asymmetrical lateral wall 6 and a circumferentially symmetrical lateral wall rear edge 18.

The circumferentially asymmetrical lateral wall front edge 16 is rounded and has a wave-shaped edge contouring as seen in the circumferential direction, appearing elliptical, as seen in the direction of flow and/or from the front lateral wall overhang to the lateral wall, or is comprised of multiple segments of circles with different radii. In the exemplary embodiments according to the preceding FIGS. 1 to 3, the edge contouring develops into the circumferentially asymmetrical lateral wall 6 and/or its lateral wall contouring.

The lateral wall contouring has a peak 64 which extends between root fillets 82, 84 of the respective neighboring turbine blades 12, 14 and develops into the root fillets 82, 84 at the height of an ideal hot gas channel line.

In the area of the suction-side root fillet 84, the peak 64 extends around an uncontoured lateral wall section 88, from which the suction-side root fillet 84 emerges smoothly. Furthermore, the lateral wall contouring has a front valley 68 and a rear valley 86, each of which is between two turbine blades 12, 14, separated from one another by the peak 64.

As also shown in FIG. 4, the lateral wall contouring tapers out from the outflow edges 90 of the turbine blades 12, 14 in an approximate axial position, so that a noncontoured lateral wall section 62 is formed between the outflow edges 90 and the lateral wall rear edge 18.

FIGS. 5 and 6 show an exemplary embodiment of the blade 2 according to the invention, which is designed as a rotor blade, such as a turbine rotor blade of a turbomachine. The blade 2 has a hub-side front platform overhang 32, a hub-side platform front edge 24, a hub-side platform 8, a hub-side platform rear edge 28 and a hub-side rear platform overhang 36.

In substantial contrast with the exemplary embodiments according to FIGS. 1 to 4, the platform 8 is not contoured but instead is designed to be circumferentially symmetrical.

In further substantial contrast, the rear platform overhang 36 continues an ideal hot gas channel line 78 of the platform 8, so that in contrast with the exemplary embodiments according to FIGS. 1, 2 and 3, it is not offset radially toward the inside in comparison with the platform 8 but instead it forms the platform rear edge 28. In the exemplary embodiments according to FIGS. 1, 2 and 3, the front platform overhang 32, the rear platform overhang 36 and the platform rear edge 28 are all circumferentially symmetrical.

The platform front edge 24 is rounded and has a circumferentially asymmetrical edge contouring, as seen in the circumferential direction, appearing elliptical in the direction of flow and/or as seen from the front platform overhang 32 to the platform 8, or is comprised of a plurality of segments of circles with different radii. As in the preceding exemplary embodiments, the edge contouring may be wave-shaped in the circumferential direction. The edge contouring develops into the circumferentially symmetrical platform contouring and/or tapers out into that.

It should be pointed out that although FIGS. 1 to 6 are based only on the hub side and/or inner lateral wall 6/platforms 8, 10, any considerations and/or contouring and/or noncontourings may of course also be applied in the case of an outer and/or housing-side lateral wall/platform. In particular, both the hub-side lateral wall 6 and the housing-side lateral wall as well as the blade platforms 8, 10 required for this purpose may be contoured accordingly for bordering a hot gas channel of a turbomachine.

A blade row for a turbomachine is disclosed, having an inner lateral wall and an outer lateral wall for bordering a hot gas channel of the turbomachine through which hot gas flows, wherein at least one of the lateral walls has a rounded lateral wall front edge, which is provided with a circumferentially asymmetrical edge contouring that is elliptical or is comprised of a plurality of segments of circles, a blade for such a blade row and a turbomachine.

LIST OF REFERENCE CHARACTERS

• 1 Blade row
• 2 Blade
• 4 Blade
• 6 Inner/hub-side lateral wall
• 8 Platform/blade platform
• 10 Platform/blade platform
• 12 Turbine blade
• 14 Turbine blade
• 16 Lateral wall front edge
• 18 Lateral wall rear edge
• 20 Front lateral wall overhang
• 22 Rear lateral wall overhang
• 24 Platform front edge
• 26 Platform front edge
• 28 Platform rear edge
• 30 Platform rear edge
• 32 Front platform overhang
• 34 Front platform overhang
• 36 Rear platform overhang
• 37 Rear platform overhang
• 38 Elevation
• 40 Elevation
• 42 Recess
• 44 Elevation section
• 46 Elevation section
• 48 Region of a root fillet on the side of the oncoming flow edge
• 50 Outer foot
• 52 Free area
• 54 Free area
• 56 Peak section
• 58 Peak section
• 60 Suction-side region of a root fillet
• 62 Noncontoured lateral wall section
• 64 Peak
• 66 Peak
• 67 Suction side
• 68 Valley
• 70 Peak section
• 72 Peak section
• 74 Peak section
• 76 Peak section
• 78 Ideal hot gas channel line/ideal platform line
• 80 Section with a reduced height
• 82 Root fillet
• 84 Root fillet
• 86 Valley
• 88 Noncontoured lateral wall section
• 90 Outflow edge

As also discussed above, the foregoing disclosure has been set forth merely to illustrate the invention and is not intended to be limiting. Since modifications of the disclosed embodiments incorporating the spirit and substance of the invention may occur to persons skilled in the art, the invention should be construed to include everything within the scope of the appended claims and equivalents thereof.

Claims

1. A blade row of a turbomachine, comprising:

an inner lateral wall;

an outer lateral wall;

wherein the inner lateral wall and the outer lateral wall border a hot gas channel of the turbomachine through which hot gas flows; and

a plurality of turbine blades disposed between the inner lateral wall and the outer lateral wall;

wherein at least one of the inner lateral wall and the outer lateral wall has a rounded lateral wall front edge which includes a circumferentially asymmetrical edge contouring;

and wherein the edge contouring is elliptical or is comprised of one or more segments of a circle.

2. The blade row according to claim 1, wherein the edge contouring is wave-shaped in a circumferential direction.

3. The blade row according to claim 1, wherein respective root fillets of the plurality of turbine blades are disposed in the edge contouring.

4. The blade row according to claim 1, wherein the edge contouring is integrated into a lateral wall contouring of the at least one of the inner lateral wall and the outer lateral wall.

5. The blade row according to claim 1, wherein the at least one of the inner lateral wall and the outer lateral wall is free of a contour and the edge contouring tapers into the at least one of the inner lateral wall and the outer lateral wall.

6. The blade row according to claim 1, wherein a lateral wall rear edge of the at least one of the inner lateral wall and the outer lateral wall is free of a contour.

7. The blade row according to claim 1, wherein a front lateral wall overhang of the at least one of the inner lateral wall and the outer lateral wall is free of a contour.

8. The blade row according to claim 1, wherein a rear lateral wall overhang of the at least one of the inner lateral wall and the outer lateral wall is free of a contour.

9. A turbomachine having a blade row according to claim 1.