Patent Application
20170356297
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.
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.
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.
Referring now to the drawings, in which like numerals refer to like elements throughout the several views,
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.
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.
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.
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.
