Abstract:
A method for use in repairing gas turbine engine components includes applying a stress to a first gas turbine engine component to cause surface cracking on the first gas turbine engine component and establishing a location of an elevated stress region of a second gas turbine engine component based upon the location of the surface cracking on the first gas turbine engine component.

Description:
[0001]    This invention was made for government support under Contract No. F33657-99-D-2051 awarded by the United States Air Force. The government therefore has certain rights in this invention. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    This disclosure relates to repairing gas turbine engine components and, more particularly, to determining a location of an elevated stress region of a gas turbine engine component. 
         [0003]    Gas turbine engine components, such as turbine blades, turbine vanes, compressor blades, compressor vanes, or other components typically operate in a relatively high stress and high temperature environment. The stresses and temperature may result in damage to the component from corrosion, erosion, deformation, or the like. Depending on the type and severity of the damage, the components may be repaired and reused. 
         [0004]    The type of repair process depends on the type of damage. For example, relatively elevated stresses and temperatures within the engine may cause deformation of a blade, vane, or other component. For a blade that is effectively cantilevered from one end, the thermal and mechanical loads may result in a twisting deformation of the blade about its axis. The blade may be restored to near its original shape by twisting the blade in the opposite direction that caused the deformation. Other stresses may cause bending or other types of deformation. 
         [0005]    Typically, after a repair process, one or more representative components are metallurgically analyzed to determine whether the stresses applied during the repair process damaged the component. For example, if the stress exceeds the yield strength of the component, cracks may form. A typical metallurgical analysis requires that the representative component be sectioned into a relatively large number of pieces. The pieces are then analyzed through known metallurgical methods for cracking or other damage. 
         [0006]    To reduce the number of sections that are required, it is desirable to predict a location of a region of highest stress on the component using computer analysis and then sectioning only that region. This would provide analysis of the region of the component that is most vulnerable to cracking. However, one possible drawback of using computer analysis is that it is based on simulation, modeling, and experimental data that may deviate from actual conditions and lead to an inaccurate prediction of the location of the highest stress region. 
         [0007]    Additionally, the computer analysis may be used to determine a maximum amount of stress that can be applied to a component during a repair process without causing cracking. However, since the computer analysis results can deviate from actual conditions that cause cracking, the stress used in the repair process may not be reliable for avoiding cracking. 
         [0008]    Accordingly, there is a need for a method to verify that a predicted location of a high stress region on a gas turbine engine component is accurate and to accurately determine a maximum amount of stress that can be used in the repair process without causing cracking. 
       SUMMARY OF THE INVENTION 
       [0009]    An example method for use in repairing a gas turbine engine component includes applying a stress to a first gas turbine engine component to cause surface cracking on the first gas turbine engine component. A location of an elevated stress region of a second gas turbine engine component is established based upon the location of the surface cracking on the first gas turbine engine component. 
         [0010]    In another aspect, the method includes establishing a predicted location of an elevated stress region on a gas turbine engine component and applying a stress to the gas turbine engine component to determine a location of an actual location of the elevated stress region. The predicted location and the actual location are compared to verify the accuracy of the predicted location. 
         [0011]    In another aspect, the method includes determining a magnitude of a first stress that causes surface cracking of a first gas turbine engine component and establishing a maximum magnitude of a second stress that is applied in a repair process to a second gas turbine engine component based upon the magnitude of a first stress. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS  
         [0012]    The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows. 
           [0013]      FIG. 1  illustrates an example gas turbine engine. 
           [0014]      FIG. 2  illustrates an example turbine blade of the gas turbine engine. 
           [0015]      FIG. 3  illustrates an example predicted location of an elevated stress region of an example turbine blade. 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
       [0016]      FIG. 1  illustrates selected portions of an example gas turbine engine  10 , such as a gas turbine engine  10  used for propulsion. In this example, the turbine engine  10  is circumferentially disposed about an engine centerline  12 . The turbine engine  10  includes a fan  14 , a compressor section  16 , a combustion section  18 , and a turbine section  20 . The combustion section  18  and the turbine section  20  include corresponding blades  22  and vanes  24 . As is known, air compressed in the compressor section  16  is mixed with fuel and burned in the combustion section  18  to produce hot gasses that are expanded in the turbine section  20 .  FIG. 1  is a somewhat schematic presentation for illustrative purposes only and is not a limitation on the disclosed examples. Additionally, there are various types of gas turbine engines, many of which could benefit from the examples disclosed herein and are not limited to the designs shown. 
         [0017]      FIG. 2  illustrates an example of one of the blades  22  from the turbine section  20  of the gas turbine engine  10 . In this example, the blade  22  includes a platform section  30  and an airfoil section  32  that extends outwards from the platform section  30 . 
         [0018]    The blade  22  is formed from a nickel alloy that is generally resistant to elevated temperatures and maintains a desired degree of strength at the elevated temperatures, for example. The blade  22  is also coated with a protective coating  34  that protects the blade  22  from corrosion, erosion, and the like. In one example, the protective coating  34  includes aluminum, such as an aluminide coating, that is at least partially diffused into the nickel alloy of the blade  22 . In one example, the protective coating  34  is MCrAlY, where the M includes at least one of nickel, cobalt, iron, or a combination thereof, Cr is chromium, Al is aluminum and Y is yttrium. Given this description, one of ordinary skill in the art will recognize that other types of protective coatings  34  may also be used. 
         [0019]    In the disclosed example, the protective coating  34  and the underlying nickel alloy of the blade  22  have different mechanical properties that are used to identify a location of an elevated stress region of the blade  22 , such as a maximum stress region. For example, the protective coating  34  is brittle relative to the underlying nickel alloy of the blade  22 . Thus, when a stress is applied to the blade  22 , the protective coating  34  tends to crack at a lower stress than the underlying nickel alloy. 
         [0020]    In the disclosed example, a stress is applied to the blade  22  to identify an elevated stress region of the blade  22 . The type of stress that is used may depend on the type of stress that the blade  22  is subjected to in the engine  10  and the type of stress that may be subsequently applied in a repair process, such as bending stress, torque stress, etc. 
         [0021]    In the disclosed example, the platform section  30  is held within a first fixture  36   a  that is secured on a support  38 . The end of the airfoil section  32  is received within a second fixture  36   b . The second fixture  36   b  may be adapted to receive a handle or other connection for applying a torque to the blade  22 . The second fixture  36   b , handle, or other connection may also include a measuring device, such as a known type of torque sensor, to determine a magnitude of the stress that is applied to the blade  22 . 
         [0022]    In one example, the second fixture  36   b  is rotated manually, as indicated by the rotational arrow, relative to the first fixture  36   a  to apply a torque stress to the blade  22 . When the torque stress exceeds a ultimate strength of the protective coating  34 , the protective coating  34  cracks. The ultimate strength of the protective coating  34  is less than a yield and ultimate strengths of the underlying nickel alloy of the turbine blade  22 . 
         [0023]    When the applied torque stress exceeds the ultimate strength of the protective coating  34 , surface cracks  40  form in the protective coating  34 . In the disclosed example, once the applied torque stress exceeds the ultimate strength of the protective coating  34 , the stress is released such that the ultimate strength of the underlying nickel alloy is not exceeded. In some examples, the formation of the surface cracking  34  is audible and thereby provides an indication that the stress should be released. 
         [0024]    The location of the surface cracks  40  on the blade  22  corresponds to a location of an elevated stress region  42  of the turbine blade  22 . In one example, the elevated stress region  42  represents a maximum stress region, where the blade  22  experienced the greatest magnitude of stress from the torque applied to the second fixture  36   b . If a measuring device is used with the second fixture  36   b , the magnitude of the torque stress that caused the surface cracking  40  can be determined. 
         [0025]    The location of the surface cracking  40 , and hence the location of the elevated stress region  42 , may be identified through the use of a dye penetrant such as a fluorescent dye. In other examples, the size of the surfacing cracking  40  may be visually discernable such that the dye is not required to identify the location. 
         [0026]    In the disclosed example, the location of the elevated stress region  42  can be used to verify a predicted location of the elevated stress region  42 . Referring to  FIG. 3 , a computer analysis, such as finite element analysis, is used to establish a predicted location  44  of the elevated stress region  42  of the blade  22 . In some examples, the predicted location  44  may include contour lines that identify varying magnitudes of stress. For example, the finite element analysis may be based upon the geometry of the turbine blade  22 , mechanical properties of the underlying nickel alloy and/or the protective coating  34 , experimental data, or other inputs. 
         [0027]    The surface cracking  40  represents the actual location of the elevated surface region  42 . To verify that the predicted location  44  is accurate, the location of the surface cracking  40  is compared to the predicted location  44 . For example, the comparison can take any suitable form, such as visual comparison or overlaying the finite element analysis result with the turbine blade  22 . If the predicted location  44  aligns with the surface cracking  40 , the predicted location  44  is accurate. However, if the predicted location  44  varies from the location of the surface cracking  40 , the predicted location  44  may not be accurate. Thus, applying a stress to cause the surface cracking  40  on the turbine blade  22  provides the benefit of verifying the accuracy of the finite element analysis used to predict the location of the elevated surface stress region  42 . 
         [0028]    Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments. 
         [0029]    The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.