Patent Application
20140348661
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.
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.
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
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
According to
As shown in
As shown in
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
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
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.
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
The lateral wall contouring of the lateral wall 6 according to the exemplary embodiment shown in
However, the blade 2 shown in
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
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.
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
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
In substantial contrast with the exemplary embodiments according to
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
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
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.
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.
| Number | Date | Country | Kind |
|---|---|---|---|
| 13169058.8 | May 2013 | EP | regional |
