AIRFOIL TRAILING EDGE

Abstract

An airfoil includes a radially inner edge extending in a radial direction of the airfoil to a radially outer edge. Also included is a leading edge extending in an axial direction of the airfoil to a trailing edge. Further included is a trailing edge geometry comprising at least one wave segment having simultaneous curvature in at least two directions.

Description

BACKGROUND OF THE INVENTION

The subject matter disclosed herein relates to airfoils, and more particularly to trailing edge regions of airfoils for use in gas turbine engines.

In gas turbine engines, such as industrial and aircraft systems, compressor blade failure is an important concern. One reason for blade failure relates to wake shed by upstream struts and stator vanes on the downstream blades. The wake creates unsteady pressure load on the blades and if the frequency of the wake matches with the natural frequency of the blades, the failure can be significant. Therefore, wake strength reduction is a common goal in gas turbine engine industries.

BRIEF DESCRIPTION OF THE INVENTION

According to one aspect of the invention, an airfoil includes a radially inner edge extending in a radial direction of the airfoil to a radially outer edge. Also included is a leading edge extending in an axial direction of the airfoil to a trailing edge. Further included is a trailing edge geometry comprising at least one wave segment having simultaneous curvature in at least two directions.

According to another aspect of the invention, a compressor includes an airfoil. Also included is a trailing edge of the airfoil. Further included is a trailing edge geometry comprising a plurality of wave segments including a first wave segment having a degree of curvature in an axial direction and a second wave segment having a degree of curvature in a circumferential direction.

According to yet another aspect of the invention, a gas turbine engine includes a compressor section. Also included is a turbine section. Further included is an airfoil disposed in at least one of the compressor section and the turbine section, the airfoil having a trailing edge comprising a geometry having at least one wave segment including simultaneous curvature in an axial direction and in a circumferential direction.

These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:

FIG. 1 is a schematic illustration of a gas turbine engine;

FIG. 2 is a partial perspective view of an inlet of a compressor section of the gas turbine engine;

FIG. 3 is a schematic illustration of the inlet of the compressor section;

FIG. 4 is a side, elevational view of the compressor section;

FIG. 5 is a perspective view of an airfoil according to a first embodiment;

FIG. 6 is a side, elevational view of the airfoil according to the embodiment of FIG. 5;

FIG. 7 is a side, elevational view of the airfoil according to a second embodiment; and

FIG. 8 is a schematic illustration of various geometries of a portion of the airfoil.

The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings.

DETAILED DESCRIPTION OF THE INVENTION

The terms “axial” and “axially” as used in this application refer to directions and orientations extending substantially parallel to a center longitudinal axis of a turbine system. The terms “radial” and “radially” as used in this application refer to directions and orientations extending substantially orthogonally to the center longitudinal axis of the turbine system. The terms “upstream” and “downstream” as used in this application refer to directions and orientations relative to an axial flow direction with respect to the center longitudinal axis of the turbine system.

Referring to FIG. 1, a turbine system, such as a gas turbine engine 10, constructed in accordance with an exemplary embodiment of the invention, is schematically illustrated. The gas turbine engine 10 includes a compressor section 12, a combustor section 14, a turbine section 16, a shaft 18 and a fuel nozzle 20. It is to be appreciated that one embodiment of the gas turbine system 10 may include a plurality of compressor sections 12, combustor sections 14, turbine sections 16, shafts 18 and fuel nozzles 20. The compressor section 12 and the turbine section 16 are coupled by the shaft 18. The shaft 18 may be a single shaft or a plurality of shaft segments coupled together to form the shaft 18.

Referring now to FIG. 2, a partial, cut-away view illustrates an inlet 22 of the compressor section 12. FIG. 3 schematically illustrates the inlet 22 of the compressor section 12. The inlet 22 generally refers to a region configured to route an incoming airflow to the compressor section 12 and comprises a compressor bell mouth 24. Half of the compressor bell mouth 24 has been removed in FIG. 2 to illustrate various vanes and blades disposed at an interior region of the compressor section 12, relative to a compressor section casing 26. The compressor bell mouth 24 includes an outer surface 28 and an inner surface 30, with the incoming airflow passing between these two surfaces. Typically, a plurality of support members 32 are operatively coupled to the outer surface 28 and the inner surface 30 for support. As shown in FIG. 3, the inlet 22 includes a strut 33 for supporting the structures and guiding the incoming airflow prior to passing over a plurality of inlet guide vanes (IGVs) 34.

Referring to FIG. 4, the plurality of IGVs 34 is arranged in a circumferentially spaced manner in what is referred to as a stage. Downstream of one or more stages of the plurality of IGVs 34 are a plurality of rotor blades and a plurality of stator vanes. The stator vanes are generally fixed to a stator or a compressor section casing 26, while the rotor blades are connected to the shaft 18. The plurality of IGVs 34 is generally fixed as well, but may pitch around a radial axis to vary the direction or amount of incoming flow. The plurality of IGVs 34 is followed by a first stage of rotor blades 36, which is in turn followed by a first stage of stator vanes 38. Disposed downstream of the first stage of stator vanes 38 is a second stage of rotor blades 40, which is followed by a second stage of stator vanes 42. It can be appreciated that the compressor section 12 may include varying numbers of stages of rotor blades and stator vanes, depending on the particular application.

Referring to FIGS. 5 and 6, an airfoil 50 is shown and represents any of the above-described compressor section airfoils. In particular, the airfoil 50 may be the strut 33, one of the plurality of IGVs 34, and/or the stator vanes. Although illustrated and described in accordance with airfoils of the compressor section 12 and the inlet 22, it is to be appreciated that airfoils in other parts of the gas turbine engine 10, such as the turbine section 16, may benefit from the embodiments of the airfoil 50 described below.

The airfoil 50 extends predominantly in an axial direction 52 from a leading edge 54 to a trailing edge 56, although curvature of the airfoil 50 is common The airfoil 50 is defined in a radial direction 58 by a radially inner edge 60 and a radially outer edge 62. In order to mitigate wake strength proximate regions downstream of the trailing edge 56, a trailing edge geometry 64 is formed along the trailing edge 56. The trailing edge geometry 64 comprises a multi-dimensional wave geometry that includes waves having curvature in multiple directions. In particular, the trailing edge geometry 64 is formed of at least one, but typically a plurality of wave segments 68 having simultaneous curvature in at least two directions. In the illustrated embodiment, the plurality of wave segments 68 includes simultaneous curvature in both the axial direction 52 and a circumferential direction 70. In other words, as each of the plurality of wave segments 68 curve in one direction (i.e., axially or circumferentially), simultaneous curvature in another direction is made.

Varying degrees of curvature may be employed in different embodiments, depending on the particular flow characteristics of the particular application. As shown in FIG. 8, the axial angle of curvature θ or indentation of the plurality of wave segments 68 may vary. In one embodiment, the angle of curvature θ ranges from about 45° to about 80°. The plurality of wave segments 68 shown in the embodiment of FIGS. 5 and 6 are radially oriented toward the radially outer edge 62. However, in an alternative embodiment, a portion of the plurality of wave segments 68 are radially oriented toward the radially outer edge 62, while a portion of the plurality of wave segments 68 are radially oriented toward the radially inner edge 60 (FIG. 7). In yet another embodiment, all of the plurality of wave segments 68 are radially oriented toward the radially inner edge 60. As shown, the number of wave segments may vary. Regardless of whether all or a portion of the plurality of wave segments 68 are radially oriented in a direction, the radial angle α of orientation may vary (FIG. 8). In one embodiment, the radial angle α ranges from about 0° to about 35°.

In an alternative embodiment, the plurality of wave segments 68 comprises an alternating arrangement of wave segments. The alternating arrangement refers to a circumferentially curved wave segment followed by an axially curved wave segment, or vice versa. This arrangement is repeated along all or a portion of the trailing edge 56.

Regardless of the precise configuration of the trailing edge geometry 64, the plurality of wave segments 68 enhance flow mixing prior to routing of the airflow to regions downstream of the airfoil 50. Efficient flow mixing reduces the effect of wake shed by the airfoil 50, thereby reducing an unsteady pressure load on downstream blades. Exemplary airflow patterns facilitating flow mixing are illustrated in FIGS. 6 and 7 and are referenced with numeral 80. In addition to enhancing blade life, the wake strength reduction assists in having a reduced axial gap between various components, thereby reducing the overall axial length of the gas turbine engine 10, and in particular the inlet 22 of the compressor section 12 for embodiments of airfoils disposed in the inlet 22 and/or the compressor section 12. For example, as shown in FIG. 3, an inlet casing length 82, an inlet outer diameter casing length 84, a strut length 86, a strut to IGV length 88, and a flange to the first rotor stage 90 may all benefit from a reduction in axial length with the embodiments of the airfoil 50 describe above.

While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.

Claims

1. An airfoil comprising:

a radially inner edge extending in a radial direction of the airfoil to a radially outer edge;

a leading edge extending in an axial direction of the airfoil to a trailing edge; and

a trailing edge geometry comprising at least one wave segment having simultaneous curvature in at least two directions.

2. The airfoil of claim 1, wherein the at least two directions comprises the axial direction and a circumferential direction.

3. The airfoil of claim 2, wherein the at least one wave segment comprises a plurality of wave segments.

4. The airfoil of claim 3, wherein at least one of the plurality of wave segments is radially oriented toward the radially inner edge and at least one of the plurality of wave segments is radially oriented toward the radially outer edge.

5. The airfoil of claim 3, wherein each of the plurality of wave segments are radially oriented toward the radially outer edge.

6. The airfoil of claim 3, wherein each of the plurality of wave segments are radially oriented toward the radially inner edge.

7. The airfoil of claim 2, wherein the airfoil comprises a strut disposed in a compressor inlet region of a gas turbine engine.

8. The airfoil of claim 2, wherein the airfoil comprises a stator vane disposed in a compressor section of a gas turbine engine.

9. A compressor comprising:

an airfoil;

a trailing edge of the airfoil; and

a trailing edge geometry comprising a plurality of wave segments including a first wave segment having a degree of curvature in an axial direction and a second wave segment having a degree of curvature in a circumferential direction.

10. The compressor of claim 9, wherein the plurality of wave segments are disposed in an alternating arrangement of curvature in the axial direction and in the circumferential direction.

11. The compressor of claim 9, wherein at least one of the plurality of wave segments is radially oriented toward a radially inner edge of the airfoil and at least one of the plurality of wave segments is radially oriented toward a radially outer edge of the airfoil.

12. The compressor of claim 9, wherein each of the plurality of wave segments are radially oriented toward a radially outer edge of the airfoil.

13. The compressor of claim 9, wherein each of the plurality of wave segments are radially oriented toward a radially inner edge of the airfoil.

14. The compressor of claim 9, wherein the airfoil comprises a strut disposed in an inlet region of the compressor.

15. The compressor of claim 9, wherein the airfoil comprises a stator vane disposed in the compressor.

16. A gas turbine engine comprising:

a compressor section;

a turbine section; and

an airfoil disposed in at least one of the compressor section and the turbine section, the airfoil having a trailing edge comprising a geometry having at least one wave segment including simultaneous curvature in an axial direction and in a circumferential direction.

17. The gas turbine engine of claim 16, wherein the at least one wave segment comprises a plurality of wave segments.

18. The gas turbine engine of claim 17, wherein at least one of the plurality of wave segments is radially oriented toward a radially inner edge of the airfoil and at least one of the plurality of wave segments is radially oriented toward a radially outer edge of the airfoil.

19. The gas turbine engine of claim 17, wherein each of the plurality of wave segments are radially oriented toward a radially outer edge of the airfoil.

20. The gas turbine engine of claim 17, wherein each of the plurality of wave segments are radially oriented toward a radially inner edge of the airfoil.