Patent 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.

Full Description:
BACKGROUND OF THE INVENTION 
       [0001]    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. 
         [0002]    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. 
         [0003]    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 
       [0004]    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. 
         [0005]    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. 
         [0006]    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 
         [0007]    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: 
           [0008]      FIG. 1  is a schematic representation of an exemplary turbine engine in which certain embodiments of the present invention may be used; 
           [0009]      FIG. 2  is a sectional view of the compressor section of the gas turbine engine of  FIG. 1 ; 
           [0010]      FIG. 3  is a sectional view of the turbine section of the gas turbine engine of  FIG. 1 ; 
           [0011]      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; 
           [0012]      FIG. 5  is a view of a turbine blade that includes a dovetail and a platform configuration according to conventional design; 
           [0013]      FIG. 6  is a view of a turbine blade that includes a dovetail and a platform configuration according to another conventional design; 
           [0014]      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 
       [0015]    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. 
         [0016]      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 . 
         [0017]      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 . 
         [0018]    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 . 
         [0019]    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. 
         [0020]      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. 
         [0021]    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 . 
         [0022]    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 . 
         [0023]    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. 
         [0024]    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. 
         [0025]    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. 
         [0026]      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. 
         [0027]      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. 
         [0028]    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 . 
         [0029]    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.

Technology Classification (CPC): 5