Patent Publication Number: US-2017356297-A1

Title: Lockwire Tab Backcut For Blade Stress Reduction (9E.04)

Description:
TECHNICAL FIELD 
     The present application and the resultant patent relate generally to gas turbine engines and more particularly relate to a modified turbine blade lockwire tab designed to divert the load path of a mounted turbine blade around a stress concentrating feature. 
     BACKGROUND OF THE INVENTION 
     Gas turbine disks may include a number of circumferentially spaced dovetails about the outer periphery of the disk defining dovetail slots therebetween. Each of the dovetail slots may receive a turbine blade axially therein. The turbine blade may have an airfoil portion and a blade dovetail with a shape complementary to the dovetail slots. The turbine blade may be cooled by air entering through a cooling slot in the disk and through grooves or slots formed in the dovetail portions of the blade. Typically, the cooling slots may extend circumferentially therearound through the alternating dovetails and dovetail slots. 
     The interface locations between the blade dovetails and the dovetail slots are potentially life-limiting locations due to overhanging blade loads and stress concentrating geometries. In the past, dovetail backcuts have been used in certain turbine engines to relieve such stresses. These backcuts, however, were minor in nature were not optimized to balance stress reduction on the disk, stress reduction on the turbine blades, and a useful life of the turbine blades. 
     Similarly, the turbine blades may be prevented by moving axially in the dovetail slots by a lockwire passing through circumferentially aligned tabs positioned about the dovetail of the respective turbine blades. These lockwire tabs also may have stress concentrating geometries that may benefit from optimized cutbacks. 
     There is thus a desire for improved turbine blades and/or disks and the interaction therebetween. Such improved turbine blades and/or disks may promote overall stress reduction for an improved turbine blade lifetime and improved system efficiency without negatively impacting the aeromechanical behavior of the turbine blades. 
     SUMMARY OF THE INVENTION 
     The present application and the resultant patent thus provide a method for reducing stress on a turbine blade wherein each of the turbine blades includes a dovetail with lockwire tab. The method may include the steps of (a) determining a starting line for a backcut relative to a lockwire tab end, (b) determining a cut angle for the backcut, and (c) removing material from the lockwire tab according to the starting line and the cut angle to form the dovetail backcut. The starting line may be positioned about 0.6 inches (about 15.24 millimeters), plus or minus 0.065 inches (about 1.65 millimeters) from the lockwire tab end along the dovetail axis. 
     The present application and the resultant patent further provide a turbine blade. The turbine blade may include an airfoil and a blade dovetail, wherein the blade dovetail a lockwire tab with includes a backcut sized and positioned according to optimized blade geometry. A starting line of the backcut, which defines a length of the backcut along a dovetail axis, is about 0.6 inches (about 15.24 millimeters), plus or minus 0.065 inches (about 1.65 millimeters) from a lockwire tab end along the dovetail axis. 
     These and other features and improvements of the present application and the resultant patent will become apparent to one of ordinary skill in the art upon review of the following detailed description when taken in conjunction with the several drawings and the appended claims. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic diagram of a gas turbine engine showing a compressor, a combustor, a turbine, and a load. 
         FIG. 2  is a perspective view of a turbine disk segment with an attached turbine blade. 
         FIG. 3  is a perspective view of the suction side of the turbine blade of  FIG. 2 . 
         FIG. 4  is a perspective view of the pressure side of the turbine blade of  FIG. 2 . 
         FIG. 5  is a partial perspective view of a turbine blade with a lockwire tab as may be described herein. 
         FIG. 6  is a partial sectional view of the turbine blade lockwire tab of  FIG. 5 . 
         FIG. 7  is a partial perspective view of an alternative embodiment of a turbine blade with a lockwire tab as may be described herein. 
     
    
    
     DETAILED DESCRIPTION 
     Referring now to the drawings, in which like numerals refer to like elements throughout the several views,  FIG. 1  shows a schematic view of gas turbine engine  10  as may be used herein. The gas turbine engine  10  may include a compressor  15 . The compressor  15  compresses an incoming flow of air  20 . The compressor  15  delivers the compressed flow of air  20  to a combustor  25 . The combustor  25  mixes the compressed flow of air  20  with a pressurized flow of fuel  30  and ignites the mixture to create a flow of combustion gases  35 . Although only a single combustor  25  is shown, the gas turbine engine  10  may include any number of combustors  25  positioned in a circumferential array and the like. The flow of combustion gases  35  is in turn delivered to a turbine  40 . The flow of combustion gases  35  drives the turbine  40  so as to produce mechanical work. The mechanical work produced in the turbine  40  drives the compressor  15  via a shaft  45  and an external load  50  such as an electrical generator and the like. 
     The gas turbine engine  10  may use natural gas, various types of syngas, liquid fuels, and/or other types of fuels and blends thereof. The gas turbine engine  10  may be any one of a number of different gas turbine engines offered by General Electric Company of Schenectady, N.Y., including, but not limited to, those such as a 7 or a 9 series heavy duty gas turbine engine and the like. The gas turbine engine  10  may have different configurations and may use other types of components. Other types of gas turbine engines also may be used herein. Multiple gas turbine engines, other types of turbines, and other types of power generation equipment also may be used herein together. 
       FIG. 2  is a perspective view of an example of a gas turbine disk segment  55  with a gas turbine blade  60 . The disk segment  55  may include a dovetail slot  65  that receives a correspondingly shaped blade dovetail  70  to secure the turbine blade  60  to the disk  55 .  FIG. 3  and  FIG. 4  show opposite sides of the turbine blade  60  including an airfoil  75  and the blade dovetail  70 .  FIG. 3  illustrates a pressure side of the turbine blade  60  and  FIG. 4  illustrates a suction side of the turbine blade  12 . The dovetail slots  65  typically are termed “axial entry” slots in that the dovetails  70  of the blades  60  may be inserted into the dovetail slots  65  in a generally axial direction, i.e., generally parallel but skewed to the axis of the disk  55 . 
     The interface surfaces between the blade dovetail  70  and the disk dovetail slot  65  may be subject to stress concentrations. An example of a stress concentrating feature may be a cooling slot. As described above, the upstream or downstream face of the turbine blade  60  and the disk  55  may be provided with an annular cooling slot that extends circumferentially there around and passes through a radially inner portion of each dovetail  70  and dovetail slot  65 . Cooling air (e.g., compressor discharge air and the like) may be supplied to the cooling slot which in turn supplies cooling air into the radially inner portions of the dovetail slots  65  for transmittal through grooves or slots (not shown) in the base portions of the blades  60  for cooling the interior of the blade airfoil portions  75 . 
     A second example of a stress concentrating feature may be a blade retention or a lockwire tab  80 . A forward end  85  or an aft end  90  of the blade  60  may be provided with the lockwire tab  80  defining an annular retention slot that extends circumferentially therearound, passing through the radially inner portion of each dovetail  70  and dovetail slot  65 . A blade retention wire may be inserted into the lockwire tab  80  which in turn provides axial retention for the blades. In either of these examples and in similar situations, the stress concentrations potentially may be life-limiting locations of the turbine disk  55  and/or turbine blade  60 . 
       FIGS. 5 and 6  show an example of a turbine blade  100  as may be described herein. In this example, the turbine blade  100  may part of a second stage of a 9E.04 gas turbine engine offered by General Electric Company of Schenectady, N.Y. Other types of gas turbine engines may be used herein. The turbine blade  100  may include an airfoil  105  and a dovetail  110  similar to that described above. The turbine blade  100  may include a lockwire tab  120  positioned about the dovetail  110 . Depending on the turbine class and blade and disk stage, the lockwire tab  120  may be on either the forward end  85  or the aft end  90  end of the dovetail  110 . In this example, the lockwire tab  120  is positioned about the aft end  90 . One or more backcuts  130  may be formed by removing a predetermined amount of material from the lockwire tab  120 . The material may be removed using any suitable process such as a grinding or milling process or the like. Moreover, these processes may be the same as or similar to the corresponding processes used for forming the blade dovetail  110  (and/or disk dovetail slot  65 ). 
     The amount of material to be removed and thus the size of the backcut  130  may be determined by first finding a starting line  150  for the dovetail backcut  130 , i.e., the starting line  150  defining a length  160  therefrom of the backcut  130  along the dovetail axis to a front end  170  or an aft end  175 . A cut angle  180  also may be determined for the backcut  130 . The starting line  150  and the cut angle  180  may be optimized according to blade and disk geometry so as to maximize a balance between stress reduction on the turbine disk  55 , stress reduction of the turbine blade  100 , a useful life of the turbine blade  100 , and maintaining or improving the aeromechanical behavior of the turbine blade  100 . As such, if a backcut  130  is too large, the backcut  130  may have a negative effect on the life span of the turbine blade  100 . If the backcut  130  is too small, although the life of the turbine blade  100  may be maximized, stress concentrations in the interface between the turbine blade and the disk may not be minimized such that the disk may not benefit from the maximized life span. The backcut  130  may be planar or non-planar. In this context, the cut angle  180  may be defined as a starting cut angle. The backcuts  130  may be formed in one or both of the pressure side and suction side of the turbine blade  100 . 
     The starting line  150  and the cut angle  180  for the backcut  130  may be determined by executing finite element analyses on the geometry of the blade and the disk. Virtual thermal and structural loads based on engine data may be applied to finite element grids of the blade  100  and the disk  55  to simulate engine operating conditions. The no-backcut geometry and a series of varying backcut geometries may be analyzed using the finite element model. A transfer function between the backcut geometry and blade and disk stresses may be inferred from the finite element analyses. The predicted stresses then may be correlated to field data using proprietary materials data in order to predict blade and disk lives and blade aeromechanical behavior for each backcut geometry. An optimum backcut geometry and an acceptable backcut geometry range may be determined through consideration of both the blade and disk life and the blade aeromechanical behavior. 
     The optimized starting line  150  and the cut angle  180  for each backcut  130  thus may be determined by using finite element analyses in order to maximize a balance between stress reduction on the turbine disk, stress reduction on the turbine blades, a useful life of the turbine blades, and maintaining or improving the aeromechanical behavior of the gas turbine blade. Although specific dimensions will be described, the turbine blade  100  described herein is not necessarily meant to be limited to such specific dimensions. The maximum dovetail backcut may be measured by the nominal distance between the starting line  150  and the front end  170  or the aft end  175 . Through the finite element analyses, it has been determined that a larger dovetail backcut would result in sacrifices to the acceptable life of the gas turbine blade. 
     Alternatively, the starting line  150  also may be determined using finite element analysis based upon a predetermined the datum line W through the dovetail  110 . The datum line W provides an identifiable reference point for each stage blade and disk of each turbine class for locating the optimized dovetail backcut starting line. In this example, the backcut  130  may be optimized for a second stage of a 9E.04 gas turbine engine offered by General Electric Company of Schenectady, N.Y. 
     The length  160  of the backcut  130  may be about 0.6 inches (about 15.24 millimeters), plus or minus 0.065 inches (about 1.65 millimeters), i.e., from the starting line  150  to the aft end  175 . Different lengths  160 , however, also may be used herein. The cut angle  180  also may be determined for the dovetail backcut  130 . In this example, the cut angle  180  may be about 1.0 degrees, plus or minus about 0.3 degrees. Other cut angles  180  may be used herein. Other suitable sizes, shapes, and configurations may be used herein. 
       FIG. 7  shows a further embodiment of a turbine blade  200  as may be described herein. In this example, the turbine blade  200  may part of a first stage of a 9E.04 gas turbine engine offered by General Electric Company of Schenectady, N.Y. The turbine blade  100  may include an airfoil  105  and a dovetail  110  similar to that described above. The turbine blade  100  may include a lockwire tab  120  positioned about the dovetail  110 . In this example, the lockwire tab  120  is positioned about the forward end  170 . The lockwire tab  120  may have a backcut  130  therein. The backcut  130  may have similar dimensions to those described above. 
     It is anticipated that the backcuts may be formed into a unit during a normal hot gas path inspection process. With this arrangement, the blade load path should be diverted around the high stress region in the disk and/or blade stress concentrating features. The relief cut parameters including an optimized starting line and an optimized cut angle define a backcut that maximizes a balance between stress reduction in the gas turbine disk, stress reduction in the gas turbine blades, a useful life of the gas turbine blades, and maintaining or improving the aeromechanical behavior of the gas turbine blade. The reduced stress concentrations serve to reduce distress in the gas turbine disk, thereby realizing a significant overall disk fatigue life benefit. 
     It should be apparent that the foregoing relates only to certain embodiments of the present application and the resultant patent. Numerous changes and modifications may be made herein by one of ordinary skill in the art without departing from the general spirit and scope of the invention as defined by the following claims and the equivalents thereof.