US20100166562A1

US20100166562A1 - Turbine blade root configurations

Info

Publication number: US20100166562A1
Application number: US12/346,334
Authority: US
Inventors: Bradley T. Boyer
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2008-12-30
Filing date: 2008-12-30
Publication date: 2010-07-01
Also published as: KR20100080452A; CN101782000A; DE102009059319A1; JP2010156338A

Abstract

A rotor blade for a turbine engine comprising a root and, extending in a radial direction from the root, an airfoil, wherein: the root comprises a shank and a dovetail; the shank extends from the dovetail and comprises a platform at an radial outward surface; the dovetail includes one or more tangs; the platform comprises an axially and circumferentially oriented surface that defines, at least in part, the inner most radial boundary of the flow path through the turbine; the dovetail is linear; and the platform is curved.

Description

BACKGROUND OF THE INVENTION

This present application relates generally to apparatus, methods and/or systems concerning improved turbine blade root configurations. More particularly, but not by way of limitation, the present application relates to apparatus, methods and/or systems pertaining to turbine blades that combine axial entry, linear dovetails with curved platforms.
The conventional configuration and design of turbine blades that have large root chords and cambers generally result in the airfoils of the blades becoming “nested.” As one of ordinary skill year will appreciate, “nested” is a common term that refers to a condition wherein the curvature of neighboring airfoils overlaps. This overlap generally means that the turbine blades, if aligned as they might be when installed in a rotor wheel of a conventional turbine engine, cannot be separated with an axial or a linear movement of one of the blades because of the interference between the nested airfoils, i.e., the airfoils would make contact and prevent separation in this manner.
To address this issue, conventional turbine blades often are designed with curved platforms and dovetails. This allows neighboring turbine blades whose airfoils are nested to be separated because, during separation, the turbine blade follows a curved route and, thereby, avoids the neighboring airfoil. However, as one of ordinary skill in the art will appreciate, turbine blades with platforms and dovetails that are curved present operational issues of their own, including, for example, increased difficulty and complexity of manufacture. In addition, as one of ordinary skill in the art will appreciate, with turbine blades that have platforms and dovetails that are curved, it is difficult or impossible to remove sets of neighboring blades from the turbine wheel at the same time because of the interference that necessarily occurs between the curved platforms and roots of neighboring blades. As a result, there remains a need for an improved turbine blade, and particularly an improved design for the root (i.e., the dovetail, shank and/or platform components) of the turbine blade that allows for more efficient manufacture, assembly, and/or operation.

BRIEF DESCRIPTION OF THE INVENTION

The present application thus describes a rotor blade for a turbine engine comprising a root and, extending in a radial direction from the root, an airfoil, wherein the root includes a linear dovetail and a curved platform.
The present application further describes a rotor blade for a turbine engine comprising a root and, extending in a radial direction from the root, an airfoil, wherein: the root comprises a shank and a dovetail; the shank extends from the dovetail and comprises a platform at an radial outward surface; the dovetail includes one or more tangs; the platform comprises an axially and circumferentially oriented surface that defines, at least in part, the inner most radial boundary of the flow path through the turbine; the dovetail is linear; and the platform is curved.
These and other features of the present application will become apparent upon review of the following detailed description of the preferred embodiments when taken in conjunction with the drawings and the appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other objects and advantages of this invention will be more completely understood and appreciated by careful study of the following more detailed description of exemplary embodiments of the invention taken in conjunction with the accompanying drawings, in which:

FIG. 1 is a schematic representation of an exemplary turbine engine in which certain embodiments of the present invention may be used;

FIG. 2 is a sectional view of the compressor section of the gas turbine engine of FIG. 1;

FIG. 3 is a sectional view of the turbine section of the gas turbine engine of FIG. 1;

FIG. 4 is a perspective view of a turbine assembly of a gas turbine engine in which certain embodiments of the present invention may be used;

FIG. 5 is a view of a turbine blade that includes a dovetail and a platform configuration according to conventional design;

FIG. 6 is a view of a turbine blade that includes a dovetail and a platform configuration according to another conventional design;

FIG. 7 is a view of a turbine blade that includes a dovetail and a platform configuration according to an exemplary embodiment of the present application.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to the figures, FIG. 1 illustrates a schematic representation of a gas turbine engine 100. In general, gas turbine engines operate by extracting energy from a pressurized flow of hot gas that is produced by the combustion of a fuel in a stream of compressed air. As illustrated in FIG. 1, gas turbine engine 100 may be configured with an axial compressor 106 that is mechanically coupled by a common shaft to a downstream turbine section or turbine 110, and a combustor 112 positioned between the compressor 106 and the turbine 110. Note that the following invention may be used in all types of turbine engines, including, for example, gas turbine engines, steam turbine engines, and aircraft engines. Hereinafter, the invention will be described in relation to a gas turbine engine, though this description is exemplary only and not intended to be limiting in any way.
FIG. 2 illustrates a view of an exemplary multi-staged axial compressor 118 that may be used in a gas turbine engine. As shown, the compressor 118 may include a plurality of stages. Each stage may include a row of compressor rotor blades 120 followed by a row of compressor stator blades 122. Thus, a first stage may include a row of compressor rotor blades 120, which rotate about a central shaft, followed by a row of compressor stator blades 122, which remain stationary during operation. The compressor stator blades 122 generally are circumferentially spaced one from the other and fixed about the axis of rotation. The compressor rotor blades 120 are circumferentially spaced and attached to the shaft such that, when the shaft rotates during operation, the compressor rotor blades 120 rotates about it. As one of ordinary skill in the art will appreciate, the compressor rotor blades 120 are configured such that, when spun about the shaft, they impart kinetic energy to the air or working fluid flowing through the compressor 118. The compressor 118 may have many other stages beyond the stages that are illustrated in FIG. 2. Additional stages may include a plurality of circumferential spaced compressor rotor blades 120 followed by a plurality of circumferentially spaced compressor stator blades 122.
FIG. 3 illustrates a partial view of an exemplary turbine section or turbine 124 that may be used in the gas turbine engine. The turbine 124 also may include a plurality of stages. Three exemplary stages are illustrated, but more or less stages may present in the turbine 124. Each stage may include a plurality of turbine buckets or turbine rotor blades 126, which rotate about the shaft during operation, and a plurality of nozzles or turbine stator blades 128, which remain stationary during operation. The turbine stator blades 128 generally are circumferentially spaced one from the other and fixed about the axis of rotation. The turbine rotor blades 126 may be mounted on a turbine wheel (not shown) for rotation about the shaft (not shown). The direction of flow of the hot gases through the hot gas path is indicated by the arrow. As one of ordinary skill in the art will appreciate, the turbine 124 may have many other stages beyond the stages that are illustrated in FIG. 3. Each additional stage may include a row of turbine stator blades 128 followed by a row of turbine rotor blades 126.
Note that as used herein, reference, without further specificity, to “rotor blades” is a reference to the rotating blades of either the compressor 118 or the turbine 124, which include both compressor rotor blades 120 and turbine rotor blades 126. Reference, without further specificity, to “stator blades” is a reference to the stationary blades of either the compressor 118 or the turbine 124, which include both compressor stator blades 122 and turbine stator blades 128. The term “blades” will be used herein to refer to either type of blade. Thus, without further specificity, the term “blades” is inclusive to all type of turbine engine blades, including compressor rotor blades 120, compressor stator blades 122, turbine rotor blades 126, and turbine stator blades 128.
In use, the rotation of compressor rotor blades 120 within the axial compressor 118 may compress a flow of air. In the combustor 112, energy may be released when the compressed air is mixed with a fuel and ignited. The resulting flow of hot gases from the combustor 112 then may be directed over the turbine rotor blades 126, which may induce the rotation of the turbine rotor blades 126 about the shaft, thus transforming the energy of the hot flow of gases into the mechanical energy of the rotating blades and, because of the connection between the rotor blades in the shaft, the rotating shaft. The mechanical energy of the shaft may then be used to drive the rotation of the compressor rotor blades 120, such that the necessary supply of compressed air is produced, and also, for example, a generator to produce electricity.
FIG. 4 depicts a portion of a turbine assembly 130 of the gas turbine engine 100. The turbine assembly 130 may be mounted downstream from the combustor (not shown in FIG. 4) for receiving hot combustion gases 131 therefrom. The turbine assembly 130 generally comprises a disk 132 having a plurality of turbine rotor blades 126 securely attached thereto. Typically, the turbine rotor blade 126 comprises an airfoil 136 that extends radially from a root 138, which it generally is integral therewith. A platform 140 is disposed at the base of the airfoil 136 and generally is also integral therewith. The turbine assembly 130 is axisymmetrical about an axial centerline axis 141. An annular shroud 142 surrounds the blades 126 and is suitably joined to a stationary stator casing (not shown). The shroud 142 provides a relatively small clearance or gap between it and the rotor blades 126, which limits the leakage of combustion gases 131 over the blades 126 during operation.
The airfoil 136 generally includes a concave pressure sidewall or pressure side 143 and a circumferentially or laterally opposite, convex suction sidewall or suction side 144. Both the pressure sidewall 143 and the suction sidewall 144 extend axially between a leading edge 146 and a trailing edge 148. The pressure sidewall 143 and the suction sidewall 144 further extend in the radial direction between the radially inner root 138 at the platform 140 and a radially outer blade tip 150.
As one of ordinary skill in the art will appreciate, the root 138 generally includes a shank 152, the outer radial surface of which is the platform 140, and a dovetail 154. The dovetail 154 is the inner radial section of the root 138, while the shank 152 is the section that connects the dovetail 154 to the airfoil 136. As illustrated, the dovetail 154 has a side entry type configuration that includes a plurality of tangs 156, which generally provides the root 138 with a serrated cross-section. The shank 152 extends from the outer radial portion of the dovetail 154 to the outer radial surface of the shank 152, which, as stated, is the platform 140. Like the airfoil 136, the root 138 may be described as having a trailing edge or face 158 and a leading edge or face 160, and, as illustrated, the root 138 may extend in a linear direction from the trailing face 158 to the leading face 160. In addition, the root 138 may be described as having a pressure face 162 and a suction face 164, which correspond, respectively, with the pressure side 143 and the suction side 144 of the airfoil 136.
The disc 132 may have a plurality of dovetail grooves 166 formed around its circumference. Each of the dovetail grooves 166 may be formed as a mate to the dovetails 154 of the rotor blades 126 such that each of the dovetails 154 may be axially inserted into the dovetail groove 162. It will be appreciated that the configuration of the dovetail 154/dovetail groove 166 connects the rotor blades 126 to the disc 132 and prevents the radial displacement of the rotor blades 126 during operation. As illustrated, the dovetail 154 may be linear, i.e., have a linear orientation from the trailing face 158 to the leading face 160, and the dovetail groove 162 may be linearly oriented as well. Formed in this manner, the rotor blades 126 may be axially inserted into the dovetail grooves 162 a linear fashion. As discussed in more detail below, a curved configuration for the root is also possible.
Note that the present invention is discussed in relation to its usage in turbine rotor blades 126. Turbine rotor blades, as stated, are the rotating blades within the turbine section of the turbine engine. This description is exemplary only, as embodiments of the invention described herein are not limited to usage with only turbine rotor blades. As one of ordinary skill in the art will appreciate, the present invention also may be applied to compressor rotor blades 120, which, generally, are the rotating blades within the compressor section of the turbine engine. Accordingly, reference herein to “rotor blades,” without further specificity, is meant to be inclusive of both turbine rotor blades and compressor rotor blades. And, examples that are applied to turbine rotor blades are not meant to exclude usage of the present invention in compressor rotor blades.
Similar to that shown in FIG. 4, FIG. 5 depicts a rotor blade with a conventional linear root 138. The linear root 138 includes a platform 140 and a dovetail 154 that have a linear orientation from the trailing face 158 to the leading face 160 of the root 138. More particularly, the pressure face 162 and the suction face 164 of the root 138 are not curved and generally run in a straight manner from the trailing face 158 to the leading face 160. It will be appreciated that the linearly oriented platform 140 is approximately rectilinear in shape. Each edge of the platform 140 may be identified by its relationship to the trailing face 158, leading face 160, the pressure face 162, and the suction face 164. Accordingly, the platform 140 may be described to include a trailing edge 170, a leading edge 172, a pressure edge 174, and a suction edge 176. Per conventional linear design, the pressure edge 174 is generally linear or straight. Similarly, the suction edge 176 is generally linear or straight. As stated, the dovetail 154 also may extend from the trailing face 158 to the leading face 160 in an approximate linear manner. Other portions of the shank 152 also may be linear. As described, performance criteria for airfoil design may require that airfoils become “nested” when positioned in an assembled configuration. When this is the case, removing blades linearly (which is what would be the case with linear configurations similar to FIG. 5) becomes impossible.
FIG. 6 depicts a rotor blade with a conventional curved root 138. The curved root may include a curved platform 140 and a curved dovetail 154. In this case, the pressure face 162 and the suction face 164 of the root 138 are curved. The pressure edge 174 of the platform 140 may form a concave curve. The suction edge 176 of the platform 140 may form a similar curve, though it may be a convex curve. As stated, the dovetail 154 also may form a similar curve. Other portions of the shank 152 may form a similar curve. The curvature for all of these components may be similar and, generally, is an arc of a circle.
FIG. 7 depicts a rotor blade with a curved platform 140 and a linear dovetail 154 according to exemplary embodiments of the present invention. As illustrated, the dovetail 154 may be substantially similar to the dovetail 154 of FIG. 5. That is, the dovetail 154 may be substantially linear and be configured to mate with a substantially linear dovetail groove 166. In some embodiments, the linear dovetail 154 and dovetail groove 166 may be aligned such that, on installation, each runs parallel with the centerline axis 141. In other embodiments, the linear dovetail 154 and the dovetail groove 166 may be skewed in relation to the direction of the centerline axis 141. While the dovetail 154 is linear, the platform 140, according to exemplary embodiments of the present invention, may be curved, i.e., substantially similar to the platform 140 configuration of FIG. 6. Specifically, as illustrated, the pressure edge 174 of the platform 140 may form a curve, which in preferred embodiments is a concave curve. Similarly, the suction edge 176 of the platform 140 may form a similar curve, though the suction edge 176 may form a convex curve. In preferred embodiments, the curvature of the suction edge 176 and the pressure edge 174 may be substantially the same, though offset by the width of the platform 140. In this manner, the pressure edge 174 of one blade may engage the suction edge 176 of a neighboring blade so that the platform 140 of the neighboring blades forms a smooth substantially continuous surface.
As illustrated, the trailing edge 170 and the leading edge 172 of the platform 140 may remain linear, though this is not required. The portions of the shank 152 below the platform generally may form a transition between the curved platform 140 and the linear dovetail 154. As stated, in some preferred embodiments, the curvature of the pressure edge 174 and the suction edge 176 may be approximately the same. In addition, in some preferred embodiments, the curve of the pressure edge 174 and the suction edge 176 may form the arc of an approximate circle. As one of ordinary skill in the art will appreciate, root configurations consistent with the present invention may provide advantages associated linear root configurations, such as the one illustrated in FIG. 5, while also providing advantages associated with curved root configurations, such as the one illustrated in FIG. 6.
From the above description of preferred embodiments of the invention, those skilled in the art will perceive improvements, changes and modifications. Such improvements, changes and modifications within the skill of the art are intended to be covered by the appended claims. Further, it should be apparent that the foregoing relates only to the described embodiments of the present application and that numerous changes and modifications may be made herein without departing from the spirit and scope of the application as defined by the following claims and the equivalents thereof.

Claims

1. A rotor blade for a turbine engine comprising a root and, extending in a radial direction from the root, an airfoil, wherein the root includes a linear dovetail and a curved platform.

2. The rotor blade according to claim 1, wherein the linear dovetail comprises one or more tangs and is configured to engage a linear dovetail groove.

3. The rotor blade according to claim 1, wherein the root comprises a shank, the shank extending from the dovetail and comprising the platform at an radial outward surface;

wherein:

the platform comprises an axially and circumferentially oriented surface that defines, at least in part, the inner most radial boundary of the flow path through the turbine; and

the airfoil extends in an outward radial direction from the platform.

4. The rotor blade according to claim 1, wherein the linear dovetail is configured to engage a linear dovetail groove that is one of parallel in relation to the direction of the centerline axis and skewed in relation to the direction of the centerline axis.

5. The rotor blade according to claim 1, wherein:

the platform comprises a pressure edge that coincides with a pressure side of the airfoil and a suction edge that coincides with a suction side of the airfoil; and

both the pressure edge and the suction edge are curved.

6. The rotor blade according to claim 5, wherein the pressure edge comprises a concave curve and the suction edge comprises a convex curve.

7. The rotor blade according to claim 6, wherein the curvature of the concave curve of the pressure edge and the curvature of the convex curve of the suction edge is substantially the same.

8. The rotor blade according to claim 6, wherein the curvature of the pressure edge and the curvature of the suction edge is configured such that, upon the proper installation of the rotor blade and a second similarly formed rotor blade that is installed in an adjacent position to the rotor blade, the suction edge of the rotor blade engages the pressure edge of the second rotor blade so that the platforms of the adjacent rotor blades form a substantially continuous surface.

9. The rotor blade according to claim 6, wherein the curvature of the pressure edge and the curvature of the suction edge is configured such that, upon the proper installation of the rotor blade and a second similarly formed rotor blade that is installed in an adjacent position to the rotor blade, the pressure edge of the rotor blade engages the suction edge of the second rotor blade so that the platforms of the adjacent rotor blades form a substantially continuous surface.

10. The rotor blade according to claim 6, wherein the curvature of the concave curve of the pressure edge and the curvature of the convex curve of the suction edge comprises the arc of an approximate circle.

11. The rotor blade according to claim 1, wherein:

the root comprises a shank, the shank extending from the dovetail and comprising the platform at an radial outward surface; and

the shank forms a transition between the curved platform and the linear dovetail.

12. The rotor blade according to claim 1, wherein the rotor blade is configured to operate as one of a rotor blade in the turbine of the turbine engine and a rotor blade in the compressor of the turbine engine.

13. A rotor blade for a turbine engine comprising a root and, extending in a radial direction from the root, an airfoil, wherein:

the root comprises a shank and a dovetail;

the shank extends from the dovetail and comprises a platform at an radial outward surface;

the dovetail includes one or more tangs;

the platform comprises an axially and circumferentially oriented surface that defines, at least in part, the inner most radial boundary of the flow path through the turbine;

the dovetail is linear; and

the platform is curved.

14. The rotor blade according to claim 13, wherein the linear dovetail is configured to engage a linear dovetail groove that is one of parallel in relation to the direction of the centerline axis and skewed in relation to the direction of the centerline axis.

15. The rotor blade according to claim 13, wherein:

both the pressure edge and the suction edge are curved.

16. The rotor blade according to claim 15, wherein the pressure edge comprises a concave curve and the suction edge comprises a convex curve.

17. The rotor blade according to claim 16, wherein the curvature of the pressure edge and the curvature of the suction edge is configured such that, upon the proper installation of the rotor blade and a second similarly formed rotor blade that is installed in an adjacent position to the rotor blade, the suction edge of the rotor blade engages the pressure edge of the second rotor blade so that the platforms of the adjacent rotor blades form a substantially continuous surface.

18. The rotor blade according to claim 16, wherein the curvature of the pressure edge and the curvature of the suction edge is configured such that, upon the proper installation of the rotor blade and a second similarly formed rotor blade that is installed in an adjacent position to the rotor blade, the pressure edge of the rotor blade engages the suction edge of the second rotor blade so that the platforms of the adjacent rotor blades form a substantially continuous surface.

19. The rotor blade according to claim 16, wherein the curvature of the concave curve of the pressure edge and the curvature of the convex curve of the suction edge comprises the arc of an approximate circle.

20. The rotor blade according to claim 13, wherein the rotor blade is configured to operate as a rotor blade in the turbine of the turbine engine.