AIRFOIL FOR A COMPRESSOR OF A TURBOMACHINE

Abstract

The invention relates to an airfoil for a compressor of a turbomachine, which extends starting from a blade root between a leading edge and a trailing edge to a blade tip, wherein the leading edge has a leading-edge thickness and the airfoil has a maximum profile thickness, the ratio of which to each other represents a relative leading-edge thickness, and the airfoil has a leading-edge wedge angle.

Claims

1. An airfoil for a compressor of a turbomachine, which extends starting from a blade root between a leading edge and a trailing edge to a blade tip, wherein the leading edge has a leading-edge thickness and the airfoil has a maximum profile thickness, the ratio of which to each other represents a relative leading-edge thickness, and the airfoil has a leading-edge wedge angle, wherein a product of the relative leading-edge thickness and the leading-edge wedge angle, in at least one cross-section of the airfoil, forms a leading-edge ratio parameter, the value of which is greater than 5.5.

2. The airfoil according to claim 1, wherein the at least one cross-section of the airfoil lies in a region in which the relative airfoil height is at least 20% of the total airfoil height.

3. The airfoil according to claim 1, wherein the value of the leading-edge ratio parameter, in at least one cross-section of the airfoil, is greater than 6.

4. The airfoil according to claim 1, wherein the at least one cross-section of the leading edge lies in a region in which the relative airfoil height is at least 40% of the total airfoil height.

5. The airfoil according to claim 1, wherein the leading-edge thickness has a value of 0.2 mm to 5 mm.

6. The airfoil according to claim 1, wherein the leading-edge wedge angle has a value of 2° to 45°.

7. An airfoil arrangement for a compressor, comprising at least one airfoil according to claim 1.

8. A compressor for a turbomachine, comprising at least one airfoil according to claim 1.

9. A turbomachine with a compressor, wherein the compressor is configured and arranged with at least one airfoil according to claim 1.

10. A compressor for a turbomachine, comprising an airfoil arrangement according to claim 7.

11. A turbomachine with a compressor, wherein the compressor is configured and arranged with an airfoil arrangement according to claim 7.

12. A turbomachine with a compressor, wherein the compressor is configured and arranged in accordance with claim 8.

Description

BRIEF DESCRIPTION OF THE DRAWING FIGURES

[0025] Further features, advantages, and possible applications of the invention ensue from the following description in connection with the figures. Herein:

[0026] FIG. 1 shows a schematic illustration of an exemplary airfoil according to the invention for a compressor of a turbomachine;

[0027] FIG. 2 shows a schematic illustration of a profile in cross-section of an exemplary airfoil according to the invention for a compressor of a turbomachine;

[0028] FIG. 3 shows a schematic illustration of a cross-section of a leading edge of an exemplary airfoil according to the invention; and

[0029] FIG. 4 shows a diagram in which leading-edge ratio parameters of airfoils of the prior art as well as a region V are depicted.

DESCRIPTION OF THE INVENTION

[0030] FIG. 1 shows a schematic illustration of an exemplary airfoil 10 for a compressor of a turbomachine. The airfoil 10 extends starting from a blade root 31 between a leading edge 11 and a trailing edge 12 to a blade tip 32. Extending between the leading edge 11 and the trailing edge 12 is a suction side 13 and an opposite-lying pressure side 14 of the airfoil 10.

[0031] A relative airfoil height sbh.sub.rel is specified starting from the blade root 31. The cross-section A of the airfoil 10 depicted in FIG. 1 lies in a region in which the relative airfoil height sbh.sub.rel is greater than 20% of the total airfoil height sbh.sub.ges. In a cross-section A of the airfoil 10 (in the flow direction), the leading-edge ratio parameter for a proposed design is greater than 5.5.

[0032] FIG. 2 shows a schematic profile of the airfoil 10 in the flow direction in cross-section A from FIG. 1. In a direction perpendicular to the axis of the drawing, the airfoil 10 extends starting from a blade root 31 to a blade tip 32, which are not depicted here. The airfoil 10 extends between an inflow-side leading edge 11 and a trailing edge 12. The airfoil 10 has essentially a convex suction side 13 and an opposite-lying, essentially concave pressure side 14. At each point to the suction side 13 and to the pressure side 14 of the profile of the airfoil 10, a profile centerline 15 has the same distance, whereby the maximum profile thickness d.sub.max represents the largest possible inscribed diameter of a circle on the profile centerline 15 of the airfoil 10.

[0033] FIG. 3 shows a schematic illustration of the leading edge 11 in cross-section A of the exemplary embodiment of the airfoil 10 from FIG. 2. Depicted in order to highlight the invention is a leading-edge thickness d.sub.LE that corresponds to the diameter of a circle on the profile centerline 15 of the airfoil 10 at the leading edge 11. At a point of the suction side 13 at which the diameter of a circle of the leading-edge thickness d.sub.LE transitions to the airfoil 10, a suction-side tangent 23 is depicted. At a point of the pressure side 14 at which diameter of a circle of the leading-edge thickness d.sub.LE transitions to the airfoil 10, a pressure-side tangent 24 is depicted. At their intersection, the two tangents 23 and 24 form a leading-edge wedge angle α.sub.w.

[0034] From the leading-edge thickness d.sub.LE that is placed in ratio to the maximum profile thickness d.sub.max depicted in FIG. 2, a relative leading-edge thickness vkd.sub.rel of the airfoil 10 is formed. A product obtained from the thus formed relative leading-edge thickness vkd.sub.rel and the leading-edge wedge angle α.sub.w affords a leading-edge ratio parameter ϑ, which characterizes the geometry of the leading edge 11 of the airfoil 10. In the case of the proposed design, the value of this leading-edge ratio parameter ϑ for an airfoil 10 is greater than 5.5.

[0035] FIG. 4 shows an illustration of a diagram that depicts curves a-e of the correlation between a relative airfoil height sbh.sub.rel and a leading-edge ratio parameter for various measured airfoils 10 of the prior art that are not designed in accordance with the invention. The relative airfoil height sbh.sub.rel is thereby specified starting from a blade root 31.

[0036] Depicted in the diagram is an advantageous region V in which the leading-edge ratio parameter ϑ is greater than 5.5. A relative airfoil height sbh.sub.rel of at least 20%, in particular, from the blade root 31 has been demonstrated to be advantageous in the design of airfoils in order to improve the operating performance and service life for an airfoil 10, in particular in regard to damage from a foreign object.