JP2008106776A - Blade/disk dovetail backcut for blade/disk stress reduction (9fa, stage 1) - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To maximize balance between stress reduction on a disk, stress reduction on a blade and a useful life of the blade. <P>SOLUTION: A turbine blade may include an airfoil; and a blade dovetail corresponding to a dovetail slot in a turbine disk. The blade dovetail may include a dovetail backcut 22 sized and positioned according to blade geometry to maximize a balance between the stress reduction on the disk, the stress reduction on the blade, a useful life of the turbine blade, and maintenance or improvement of the aerodynamical behavior of the turbine blade. A start point of the dovetail backcut 22 which defines a length of the dovetail backcut along a dovetail axis, may be determined relative to a datum line positioned a fixed distance from a front surface of the blade dovetail along a center line of the dovetail axis. The start point of the dovetail backcut 22 may be at least approximately 1.539 inches in an aft direction from the datum line. <P>COPYRIGHT: (C)2008,JPO&INPIT

Description

  The present invention relates to gas turbine technology, and more particularly, to a blade load path around a stress concentration feature in a disk on which the blade is mounted and / or around a stress concentration feature in the blade itself. The present invention relates to an improved blade and / or disk dovetail designed to warp.

  Some gas turbine disks include a plurality of circumferentially spaced dovetails around the outer periphery of the disk, the plurality of dovetails forming a dovetail slot therebetween. Each of the dovetail slots axially receives a blade formed with an airfoil portion and a blade dovetail having a shape complementary to the dovetail slot.

  The blade can be cooled by air flowing through cooling slots in the disk and through grooves or slots formed in the dovetail portion of the blade. Generally, the cooling slots extend 360 ° circumferentially through alternating dovetails and dovetail slots.

  It has been found that the position of contact between the blade dovetail and the dovetail slot can be a life-limited position due to the overhanging blade load and stress concentration geometry. In the past, dovetail backcuts have been used to relieve stress in some turbine engines. However, these backcuts are inherently small and have nothing to do with the problem to be addressed here. Further, the position and amount of material to be removed has not been optimized to maximize the balance between stress reduction on the disk, stress reduction on the blade, and the useful life of the blade.

  Accordingly, the present application provides at least a turbine disk and a turbine blade having a plurality of turbine blades attached to the disk and each turbine blade including a blade dovetail engageable within a correspondingly shaped dovetail slot of the disk. A method for reducing the stress above one is described. The method includes (a) determining a start point of the dovetail backcut that defines a length of the dovetail backcut along the dovetail axis with respect to a reference line; and (b) determining a cut angle of the dovetail backcut. And (c) removing material from at least one of the blade dovetail or disk dovetail slot according to the starting point and the cut angle to form a dovetail backcut. The starting point and angle of cut depends on the geometry of the blade and disk: the stress reduction on the disk, the stress reduction on the blade, the useful life of the turbine blade, and the maintenance or improvement of the aerodynamic characteristics of the turbine blade. Can be optimized to maximize the balance between them. The reference line can be located at a constant distance from the front face of the blade dovetail along the center line of the dovetail axis, and step (a) includes at least about 1 in the backward direction from the reference line where the start point of the dovetail backcut is .539 inches.

  In some embodiments, each of the turbine blades can be configured to operate within the first stage of a 9FA turbine. Step (b) can be performed such that the cut angle is a maximum of about 3 °. In other embodiments, step (b) can be performed such that the cut angle is about 0.7 °. The constant distance from the front face of the blade dovetail at the reference line is about 2.964 inches.

  In some embodiments, optimization of the starting point and cut angle can be performed by performing finite element analysis on the blade and disk geometry. Step (b) can be performed by determining a plurality of cut angles to form a dovetail backcut having a non-planar surface. Step (c) can be performed by removing material from the blade dovetail. In other embodiments, step (c) can be performed by removing material from the disk dovetail slot. Step (c) can also be performed by removing material from the blade dovetail and the disk dovetail slot. Step (c) can further be performed such that the angle obtained based on the material removed from the blade dovetail and the disk dovetail slot does not exceed the cut angle.

  The present application further describes a turbine blade that may include an airfoil and a blade dovetail shaped to correspond to a dovetail slot in the turbine disk. The blade dovetail maximizes the balance between stress reduction on the disk, stress reduction on the blade, the useful life of the turbine blade, and maintenance or improvement of the aerodynamic properties of the turbine blade, according to the blade geometry. Dovetail backcuts that are sized and arranged to include The starting point of the dovetail backcut, which defines the length of the dovetail backcut along the dovetail axis, is determined with respect to a reference line located at a constant distance from the front face of the blade dovetail along the centerline of the dovetail axis Can do. The starting point of the dovetail backcut may be at least about 1.539 inches backward from the baseline.

  In some embodiments, each of the turbine blades can be configured to operate within the first stage of a 9FA turbine. The cut angle of the dovetail backcut can be up to about 3 °. In some embodiments, the dovetail backcut cut angle may be about 0.7 °. The constant distance from the front face of the blade dovetail at the reference line may be about 2.964 inches. In some embodiments, the dovetail backcut can have a non-planar surface.

  The present application further describes a turbine blade that includes an airfoil and a blade dovetail shaped to correspond to a dovetail slot in the turbine disk. Each of the turbine blades may be configured to operate within the first stage of the 9FA turbine. The starting point of the dovetail backcut, which defines the length of the dovetail backcut along the dovetail axis, is determined with respect to a reference line located at a constant distance from the front face of the blade dovetail along the centerline of the dovetail axis Can do. The starting point of the dovetail backcut may be at least about 1.539 inches backward from the baseline. The cut angle of the dovetail backcut can be up to about 3 °.

  In some embodiments, the dovetail backcut cut angle may be about 0.7 °. The constant distance from the front face of the blade dovetail at the reference line may be about 2.964 inches. These and other features of the present application will become apparent upon reading the following detailed description of the preferred embodiments in connection with the drawings and the claims.

  FIG. 1 is a perspective view of an exemplary gas turbine disk segment 10 having a gas turbine blade 12 secured therein. The gas turbine disk 10 includes a dovetail slot 14 that receives a correspondingly shaped blade dovetail 16 to secure the gas turbine blade 12 to the disk 10. FIGS. 2 and 3 show the opposite sides of the bottom section of the gas turbine blade 12 including the airfoil 18 and the blade dovetail 16. FIG. 2 shows the so-called pressure side of the gas turbine blade 12, and FIG. 3 shows the so-called negative pressure side of the gas turbine blade 12.

  The dovetail slot 14 is generally referred to as an “axially inserted” slot, and the dovetail 16 of the blade 12 is generally axial, that is, oblique to the axis of the disk 10 but substantially parallel to the axis. It is inserted into the slot 14.

  An example of a stress concentration feature of a gas turbine disk is a cooling slot. An upstream or downstream surface of the blade and disk 10 may be provided with an annular cooling slot that extends completely 360 ° circumferentially through the radially inner portion of each dovetail 16 and dovetail slot 14. When the blade is mounted on the rotor disk 10, cooling air (eg, compressor discharge air) is supplied to the cooling slot, which then supplies cooling air into the radially inner portion of the dovetail slot 14. This feeds the interior of the feed blade airfoil 18 through a groove or slot (not shown) that opens through the base of the blade 12.

  A second example of a gas turbine disk stress concentration feature is a blade retention wire slot. An upstream or downstream surface of the blade 12 and disk 10 can be provided with an annular retaining slot that extends through the radially inner portion of each dovetail 16 and dovetail slot 14 and extends completely 360 ° circumferentially. It will be appreciated that when the blade is mounted on the rotor disk 10, the blade retention wire is inserted into the retention wire slot and the blade retention wire then provides axial retention to the blade.

  The features described herein are generally applicable to any airfoil and disk contact surface. The structure shown in FIGS. 1-3 is merely representative of many different disk and blade designs across different classes of turbines. For example, at least three classes of gas turbines, including disks and blades of different sizes and configurations, are manufactured by General Electric Company (GE), Schenectady, NY. These include (1) the first turbine class, ie GE 6FA and 6FA + e turbine, (2) the second turbine class, ie GE 7FA and 7FA + e turbine, and (3) the third turbine class, ie GE. 9FA and 9FA + e turbines are included. In addition, each turbine includes multiple stages within the turbine having varying blade and disk geometries.

  It has been found that the contact surface between the blade dovetail 16 and the disk dovetail slot 14 is subject to stress concentrations, which can be the life limit position of the turbine disk 10 and / or turbine blade 12. . It may be desirable to reduce such stress concentrations to maximize disk and / or blade life without adversely affecting gas turbine blade life or aerodynamic characteristics.

  4-7, the gas turbine blade dovetail 16 includes several pressure surfaces or tangs 20 on the dovetail pressure side and several pressure surfaces or tangs 20 on the dovetail suction side. Depending on the turbine class and blade and disk stage, a backcut 22 is formed on either or both of the suction side rear end and the pressure side front end of the blade dovetail tongue 20 or the disk dovetail tongue 21 (see FIG. 1). can do. With particular reference to FIGS. 6 and 7, the backcut 22 is formed by removing material from the pressure surface 20 of the blade dovetail 16 or disk dovetail slot 14. The material can be removed using any suitable processing method, such as grinding, milling or the like, which processing method corresponds to the blade dovetail 16 or disk dovetail slot 14 used. It can be the same as or similar to the processing method.

  The amount of material removed, and thus the size of the backcut 22, is determined by first determining the start of the dovetail backcut relative to the baseline defining the length of the dovetail backcut along the dovetail axis. The For the dovetail backcut, the cut angle is also determined, with the exemplary angle shown in FIGS. 6 and 7 being a maximum of 3 °. The starting point and infeed angle depend on the geometry of the blade and disk, the stress reduction on the gas turbine disk 10, the stress reduction of the gas turbine blade 12, the useful life of the gas turbine blade 12, and the aerodynamics of the gas turbine blade. Optimized to maximize the balance between maintaining or improving physical characteristics. Accordingly, if the dovetail backcut 22 is too large, the backcut will adversely affect the life of the turbine blade 12. If the dovetail backcut is too small, the life of the turbine blade will be maximized, but the stress concentration at the interface between the turbine blade and the disk will not be minimized, and the disk will You will not benefit from maximization.

  The backcut 22 can be planar, or alternatively, the backcut 22 'can be non-planar, as shown by the dashed line in FIG. In this connection, the cutting angle is defined as the starting cutting angle. For some turbine classes, the cut angle is reasonable from the start until the backcuts 22, 22 ′ are deep enough that the blade loading surface of the blade dovetail 16 does not contact the disk dovetail slot 14. It has become. If contact with the disk slot 14 is lost, any cut of any depth or shape outside the defined envelope will be acceptable.

  As described above, if the blade dovetail 16 and the disk dovetail slot 14 include several tangs 20, the starting point and / or incision angle of the dovetail backcut is determined separately for each of the several tangs. can do. In this connection, as also mentioned above, the dovetail backcut can be formed in one or both of the pressure side and the suction side of the turbine blade and / or disk.

  Optimization of the dovetail backcut start point and cut angle is determined by performing finite element analysis on the blade and disk geometry. Virtual thermal and structural loads based on engine data are applied to the blade and disk finite element grids to simulate engine operating conditions. The geometry without backcuts and a series of changing backcut geometries are analyzed using a finite element model. The transfer function between backcut geometry and blade and disk stress is inferred from finite element analysis. The proprietary material data is then used to correlate the predicted stress to the field data to predict blade and disk life and blade aerodynamic characteristics for each backcut geometry. The optimal backcut geometry and acceptable backcut geometry range are determined by considering both blade and disk life and blade aerodynamic characteristics.

  The reference line W also varies according to the geometry of the blade or disk. The reference line W is located at a certain distance from the front surface of the blade or disk dovetail along the center line of the dovetail axis. FIGS. 21-30 illustrate the determination of the reference line W for each of the General Electric turbine classes described above and for each blade and disk stage. For example, FIG. 21 illustrates the determination of the baseline W for the first stage blades and disks in the first type of first turbine class (6FA turbine), where the baseline W is the dovetail. It is located 1.704 inches from the front face of the blade and disk dovetail along the axis centerline (reference line S). FIG. 22 shows the determination of the reference line W for the first stage blades and disks in the first turbine class (6FA + e turbine) of the second type, where the reference line W is the dovetail axis. Located 1.698 inches from the front of the blade and disk dovetail along the center line (reference line S). FIG. 23 shows the determination of the reference line W for the second stage blades and discs in the first turbine class (6FA + e turbine) of the second type, where the reference line W is the dovetail axis. Located 1.936 inches from the front of the blade and disk dovetail along the center line (reference line S).

  FIG. 24 shows a dimension of 2.470 inches for the first stage blades and disks in the second turbine class of the first type (7FA + e turbine). FIG. 25 shows a dimension of 2.817 inches in the second stage blade and disk in the second type of second turbine class (7FA + e turbine).

  FIG. 26 shows a dimension of 2.946 inches for first stage blades and disks in a third turbine class of the first type (9FA + e turbine). FIG. 27 shows a 3.379 inch dimension for the second stage blade and disk in the third turbine class of the first type (9FA + e turbine). FIG. 28 shows a dimension of 2.964 inches for the first stage blade in the second type of third turbine class (9FA turbine).

  FIG. 29 shows a dimension of 2.470 inches for the first stage blade and disk in the second type of second turbine class (7FA turbine). FIG. 30 shows a dimension of 2.817 inches for the second stage blade and disk in the second type of second turbine class (7FA turbine). The reference line W provides an identifiable reference point for each stage blade and disk of each turbine class for locating the optimized dovetail backcut starting point.

  Details of the optimized starting point and cut angle for each turbine class at each respective blade and disk stage will be described with reference to FIGS. 8-20 and 31-36. As described above, the optimum starting point and cut angle for each dovetail backcut is the stress reduction on the gas turbine disk, the stress reduction on the gas turbine blade, the useful life of the gas turbine blade, and the aerodynamics of the gas turbine blade. It has been determined using finite element analysis to optimize to maximize the balance between maintenance or improvement of mechanical properties. Although specific dimensions are described, the present invention is not necessarily intended to be limited to such specific dimensions. The maximum dovetail backcut is measured by the nominal dimension from the reference line W to the starting point shown. By finite element analysis, it was determined that a larger dovetail backcut would sacrifice the acceptable life of the gas turbine blade. In describing the optimal dimensions, distinct values can be determined for several tangs 20 of the blade dovetail 16 and / or disk dovetail slot 14.

  FIGS. 8 and 9 show first stage blades in a first type of first turbine class (6FA turbine), including three sets of dovetail tangs, here identified by the overall width between sets of tangs. In this case, the starting point of the dovetail backcut is at least 1.619 inches backward from the reference line W for the wide tongue, and backward from the reference line W for the intermediate tongue. It is at least 1.552 inches, and at least 1.419 inches backward from the reference line W for narrow tangs. The cutting angle is a maximum of 3 °.

  FIGS. 10 and 11 show first stage blades in a second type of first turbine class (6FA + e turbines), including three sets of dovetail tangs, here identified by the overall width between sets of tangs. In this case, the starting point of the dovetail backcut is at least 1.549 inches backward from the reference line W for wide and intermediate tongues, and the reference line W for narrow tongues. At least 1.466 inches in the rearward direction. The cutting angle is a maximum of 3 °. FIG. 12 shows the second stage blades and disks in a second type of first turbine class (6FA + e turbine), including three sets of dovetail tangs, here identified by the overall width between sets of tangs. Show. FIG. 12 shows the starting point of a dovetail backcut that is at least 0.923 inches backward from the reference line W for the wide tongue and at least 1.654 inches backward from the reference line W for the middle tongue. Yes.

  FIGS. 13 and 14 show values for first stage blades and disks in a first type of second turbine class (7FA + e turbine) that includes three sets of dovetail tangs. The starting point of the dovetail backcut is at least 1.945 inches in the rearward direction from the reference line, and the cut angle is a maximum of 3 °. Here, the pressure side of the second stage blade and disk in the second turbine class of the first type (7FA + e turbine) including three sets of dovetail tangs identified by the overall width between the sets of tongues 15 is at least 1.574 inches forward from the reference line W for the wide tongue, at least 1.400 inches forward from the reference line W for the middle tongue, and forward from the reference line W for the narrow tongue. Shows the starting point of the dovetail backcut at least 1.226 inches. The cutting angle is a maximum of 5 °. As shown in FIG. 16, for the suction side of the second stage blade and disk in the first type second turbine class (7FA + e turbine) including three sets of dovetail tongues, the starting point of the dovetail backcut is the reference Located at least 1.725 inches backward from the line. The cutting angle is a maximum of 5 °. 17 and 18 show first stage blades and disks in a first type of third turbine class (9FA + e turbine) that includes three sets of dovetail tangs, in this case the dovetail backcut The starting point is at least 1.839 inches backward from the reference line W. The cutting angle is a maximum of 3 °. FIG. 19 shows the pressure side of the second stage blade in a first type of third turbine class (9FA + e turbine) that includes three sets of dovetail tangs. The starting point of the dovetail backcut is at least 1.848 inches forward from the reference line W. The cutting angle is a maximum of 5 °. FIG. 20 shows the suction side of the second stage blade and disk in a third turbine class (9FA + e turbine) of the first type including three sets of dovetail tangs. The starting point of the dovetail backcut is at least 2.153 inches in the rearward direction from the reference line W, and the cut angle is a maximum of 5 °.

  FIGS. 31 and 32 show a second type of third turbine class (9FA) first stage blades and disks, according to an exemplary embodiment of the present application. FIG. 31 shows the backcut removal area located on the suction side rear end of the dovetail as shown. The backcut can be at least about 1.539 inches in the rearward direction from the reference line W for each of the three dovetail pressure surfaces (ie, tongues). The cutting angle can be varied between about 0 ° and 3 °. In some embodiments as shown in FIG. 32, the cut angle can be about 0.7 °. Thus, for example, in some embodiments, the backcut enters each pressure surface at an angle of 0.7 ° in the above position, and then proceeds through the remainder of the dovetail at an angle of 0.7 °. Can do.

  FIGS. 33 and 34 show second stage second turbine class (7FA) first stage blades and disks, according to an exemplary embodiment of the present application. FIG. 33 shows the backcut removal area located on the suction side rear end of the dovetail as shown. The backcut can be at least about 1.645 inches backward from the reference line W for each of the three dovetail pressure surfaces (ie, tongues). The cutting angle can be varied between about 0 ° and 3 °. In some embodiments as shown in FIG. 34, the cut angle may be about 0.7 °. Thus, for example, in some embodiments, the backcut enters each pressure surface at an angle of 0.7 ° in the above position, and then proceeds through the remainder of the dovetail at an angle of 0.7 °. Can do.

  FIGS. 35 and 36 illustrate a second turbine class (7FA) second stage blade and disk of the second type, according to an exemplary embodiment of the present application. FIG. 35 shows the backcut removal area located on the suction side rear end of the dovetail as shown. The backcut can begin at least about 1.215 inches rearward from the reference line W for each of the three dovetail pressure surfaces (ie, tongues). The cutting angle can be varied between about 0 ° and 3 °. In some embodiments as shown in FIG. 36, the cut angle may be about 2.0 °. Thus, for example, in some embodiments, the backcut enters each pressure surface at an angle of 2.0 ° in the above position and then proceeds through the rest of the dovetail at an angle of 2.0 °. Can do.

  It is anticipated that the dovetail backcut can be formed in the unit during the normal hot gas flow path inspection process. For this configuration, the blade load path should be counter to around the high stress area in the disk and / or blade stress concentration feature. Recess cut parameters, including optimal starting point and optimal insertion angle relative to the baseline, reduce stress in the gas turbine disk, reduce stress in the gas turbine blade, useful life of the gas turbine blade, and aerodynamics of the gas turbine blade A dovetail backcut is created that maximizes the balance between maintaining or improving properties. The reduction in stress concentration helps reduce fatigue in gas turbine disks, thereby realizing a significant advantage in overall disk fatigue life.

  Although the present invention has been described with respect to what are presently considered to be the most practical and preferred embodiments, the invention is not limited to the disclosed embodiments, but rather, the spirit and technology of the appended claims It should be understood that various changes and equivalent arrangements included within the scope are intended to be protected.

1 is a perspective view of an exemplary gas turbine disk segment with gas turbine blades attached. FIG. FIG. 3 is a perspective view of the pressure side of this exemplary gas turbine blade. FIG. 3 is a perspective view of the negative side of this exemplary gas turbine blade. Figure 3 is an enlarged view of a blade or disk dovetail area where material is to be removed. Figure 3 is an enlarged view of a blade or disk dovetail area where material is to be removed. Figure 3 is an enlarged view of a blade or disk dovetail area where material is to be removed. Figure 3 is an enlarged view of a blade or disk dovetail area where material is to be removed. FIG. 2 shows a material removal area in a first stage blade or disk in a first type of first turbine class (6FA turbine). FIG. 2 shows a material removal area in a first stage blade or disk in a first type of first turbine class (6FA turbine). FIG. 3 shows a material removal area in a first stage blade or disk in a second type of first turbine class (6FA + e turbine). FIG. 3 shows a material removal area in a first stage blade or disk in a second type of first turbine class (6FA + e turbine). FIG. 4 shows a material removal area in a second stage blade or disk in a first turbine class (6FA + e turbine) of the second type. FIG. 5 shows a material removal area in a first stage blade or disk in a first type of second turbine class (7FA + e turbine). FIG. 5 shows a material removal area in a first stage blade or disk in a first type of second turbine class (7FA + e turbine). FIG. 3 shows a material removal zone on the pressure side of a second stage blade or disk in a second turbine class of first type (7FA + e turbine). FIG. 3 shows a material removal area on the suction side of a second stage blade or disk in a second turbine class of first type (7FA + e turbine). FIG. 3 shows a material removal area in a first stage blade or disk in a first type of third turbine class (9FA + e turbine). FIG. 3 shows a material removal area in a first stage blade or disk in a first type of third turbine class (9FA + e turbine). FIG. 4 shows a material removal area on the pressure side of a second stage blade or disk in a third turbine class of the first type (9FA + e turbine). FIG. 4 shows a material removal area on the suction side of a second stage blade or disk in a third turbine class of the first type (9FA + e turbine). The figure which shows the determination of the reference line W in each stage blade or disk of each turbine class and type. The figure which shows the determination of the reference line W in each stage blade or disk of each turbine class and type. The figure which shows the determination of the reference line W in each stage blade or disk of each turbine class and type. The figure which shows the determination of the reference line W in each stage blade or disk of each turbine class and type. The figure which shows the determination of the reference line W in each stage blade or disk of each turbine class and type. The figure which shows the determination of the reference line W in each stage blade or disk of each turbine class and type. The figure which shows the determination of the reference line W in each stage blade or disk of each turbine class and type. The figure which shows the determination of the reference line W in each stage blade or disk of each turbine class and type. The figure which shows the determination of the reference line W in each stage blade or disk of each turbine class and type. The figure which shows the determination of the reference line W in each stage blade or disk of each turbine class and type. FIG. 3 shows a material removal area in a first stage blade or disk in a second type of third turbine class (9FA turbine). FIG. 3 shows a material removal area in a first stage blade or disk in a second type of third turbine class (9FA turbine). FIG. 3 shows a material removal area in a first stage blade or disk in a second type of second turbine class (7FA turbine). FIG. 3 shows a material removal area in a first stage blade or disk in a second type of second turbine class (7FA turbine). FIG. 4 shows a material removal area in a second stage blade or disk in a second type of second turbine class (7FA turbine). FIG. 4 shows a material removal area in a second stage blade or disk in a second type of second turbine class (7FA turbine).

Explanation of symbols

10 Gas turbine disc segment 12 Gas turbine blade 14 Dovetail slot 16 Blade dovetail 18 Airfoil 20 Blade dovetail tongue 21 Disc dovetail tongue 22, 22 'Back cut W Reference line

Claims (10)

A plurality of turbine blades (12) are attached to the disk (10) and each of the turbine blades (12) has a blade dovetail (16) engageable within a correspondingly shaped dovetail slot (14) of the disk (10). A method for reducing stress on at least one of a turbine disk (10) and a turbine blade (12) adapted to comprise:
(A) determining a start point of the dovetail backcut (22) relative to a reference line defining a length of the dovetail backcut (22) along the dovetail axis;
(B) determining a cut angle of the dovetail backcut (22);
(C) removing material from at least one of the blade dovetail (16) or disk dovetail slot (14) according to the starting point and cut angle to form the dovetail backcut (22);
The starting point and cut angle are determined according to the geometry of the blade (12) and disk (10), the stress reduction on the disk (10), the stress reduction on the blade (12), and the turbine blade ( Optimized to maximize the balance between the useful life of 12) and the maintenance or improvement of the aerodynamic characteristics of the turbine blade (12);
The reference line is positioned at a certain distance from the front surface of the blade dovetail (16) along the center line of the dovetail axis, and the step (a) includes the start point of the dovetail backcut (22) as the reference. Executed to be at least 1.539 inches backward from the line,
Method.   The method of any preceding claim, wherein each of the turbine blades (12) is configured to operate within a first stage of a 9FA turbine.   The method of claim 1, wherein step (b) is performed such that the cut angle is a maximum of 3 °.   The method of claim 1, wherein step (b) is performed such that the cut angle is 0.7 °.   The method of claim 1, wherein the constant distance from the front surface of the blade dovetail (16) at the reference line is 2.964 inches. A turbine blade (12),
An airfoil (18) and a blade dovetail (16) shaped to correspond to a dovetail slot (14) in the turbine disk (10);
The blade dovetail (16), according to blade geometry, reduces stress on the disk (10), reduces stress on the blade (12), useful life of the turbine blade (12), and the turbine blade A dovetail backcut (22) sized and arranged to maximize the balance between maintaining or improving the aerodynamic properties of (12);
The starting point of the dovetail backcut (22) defining the length of the dovetail backcut (22) along the dovetail axis is a fixed distance from the front of the blade dovetail (16) along the centerline of the dovetail axis Determined with respect to the reference line located at
The starting point of the dovetail backcut (22) is at least 1.539 inches in the rearward direction from the reference line;
Turbine blade (12).   The turbine blade (12) of claim 6, wherein each of the turbine blades (12) is configured to operate within a first stage of a 9FA turbine.   The turbine blade (12) according to claim 6, wherein the cut angle of the dovetail backcut (22) is a maximum of 3 °.   The turbine blade (12) according to claim 6, wherein the cut angle of the dovetail backcut (22) is 0.7 °.   The turbine blade (12) of claim 6, wherein the constant distance from the front surface of the blade dovetail (16) at the reference line is 2.964 inches.

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