Abstract:
A metallic component is by at least one peripheral edge. The component includes at least one elongated treated zone having a length substantially greater than its width. This treated zone is spaced away from and disposed generally parallel to the peripheral edge of the component and the entire thickness of the component within the treated zone is in a state of residual compressive stress. Crack growth from the edge due to fatigue or damage is resisted.

Description:
BACKGROUND OF THE INVENTION 
   This invention relates generally to metallic components and more particularly for a method of producing fatigue-resistant and damage-tolerant metallic components. 
   Various metallic components, such as gas turbine engine fan and compressor blades, are susceptible to cracking from fatigue and damage (e.g. from foreign object impacts). This damage reduces the life of the part, requiring repair or replacement. 
   It is known to protect components from crack propagation by inducing residual compressive stresses therein. Methods of imparting this stress include shot peening, laser shock peening (LSP), pinch peening, and low plasticity burnishing (LPB). These methods are typically employed by applying a “patch” of residual compressive stresses over an area to be protected from crack propagation, for example a leading edge of a gas turbine engine compressor blade. However, this process is relatively slow and expensive. This procedure also modifies areas of the blade in which the aerodynamic performance is quite sensitive to dimensional variations. 
   Accordingly, there is a need for a method of protecting metallic components from crack propagation without altering the performance characteristics thereof. 
   BRIEF SUMMARY OF THE INVENTION 
   The above-mentioned need is met by the present invention, which according to one aspect provides a metallic component bounded by at least one peripheral edge. The component includes at least one elongated treated zone which is spaced away at an offset distance from and the peripheral edge. The entire thickness of the component within the treated zone is in a state of residual compressive stress, such that crack growth from the peripheral edge is resisted. 
   According to another aspect of the invention, a metallic airfoil for a gas turbine engine includes a root spaced apart from a tip, spaced-apart leading and trailing edges, a suction side extending from the leading edge to the trailing edge, and an opposed pressure side extending from the leading edge to the trailing edge, wherein a thickness of the airfoil is defined between the pressure side and the suction side. The airfoil includes at least a first elongated treated zone within which the entire thickness of the airfoil is under residual compressive stress. The treated zone is spaced apart at a first offset distance from a selected one of the leading edge, the trailing edge, or the tip. The treated zone resists the growth of cracks propagating from the selected leading edge, trailing edge, or tip, respectively. 
   According to another aspect of the invention, a method of reducing crack propagation in metallic components includes providing a metallic component which is bounded by at least one peripheral edge, the component including at least one crack-prone area bounded in part by the peripheral edge. Pressure is applied to the component in at least one elongated treated zone, such that the entire thickness of the component within the treated zone is left a state of residual compressive stress, such that cracks are prevented from propagating out of the crack-prone zone. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
       FIG. 1  is a schematic perspective view of a prior art gas turbine engine compressor blade; 
       FIG. 2  is a side view of the compressor blade of  FIG. 2 ; 
       FIG. 3  is a side view of another compressor blade treated in accordance with the method of the present invention; and 
       FIG. 4  is a side view of another compressor blade treated in accordance with the present invention. 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  illustrates an exemplary gas turbine engine compressor blade  10 . This component is used merely as an example, and the present invention is equally applicable to other types of metallic components susceptible to cracking from fatigue or damage, such as compressor stator vanes, fan blades, turbine blades, shafts and rotors, stationary frames, actuator hardware and the like. The compressor blade  10  comprises an airfoil  12 , a platform  14 , and a shank  16 . In this particular example the shank  16  includes a dovetail  18  for being received in a slot of a rotating disk (not shown). The airfoil  12  has a leading edge  20 , a trailing edge  22 , a tip  24 , a root  26 , a suction side  28 , and a pressure side  30 . 
     FIG. 2  shows an enlarged portion of the compressor blade  10 . Certain areas of the compressor blade  10  are “crack-prone” or particularly subject to crack initiation, for example because of fatigue or damage from foreign object impacts. These areas include the leading edge  20 , the trailing edge  22 , and the tip  24 . Exemplary cracks “C” are shown in these areas of the compressor blade  10 . If the compressor blade  10  is left in service, the cracks C will propagate further into the compressor blade  10 , eventually rendering it unfit for further service or even causing component failure. 
   It is known to apply areas of residual compressive stress to metallic components to prevent or delay cracking. In the prior art, relatively large areas of residual compressive stress are applied to vulnerable areas such as the above-mentioned leading edge  20 , trailing edge  22 , and tip  24 . However, these surface area patches applied at the edge of a component are inefficient. First, they are expensive and time consuming because of the relatively large areas involved. Second, certain areas of the compressor blade  10  are critical to the aerodynamic performance, for example the leading and trailing edges, and are sensitive to small dimensional changes that the surface area patch can induce. 
     FIG. 3  shows a compressor blade  110  which has been treated in accordance with the present invention. The compressor blade  110  is substantially similar in construction to the compressor blade  10  and includes leading edge  120 , trailing edge  122 , and tip  124 . The compressor blade  110  incorporates thin treated zones therein. Within each treated zone, the material of the compressor blade  110  is in a condition of residual compressive stress throughout its thickness. A first treated zone  32  extends in a generally radial direction at a first offset distance D 1  from the leading edge  120 . A second treated zone  34  extends in a generally chordwise direction at a second offset distance D 2  from the tip  124 . A third treated zone  36  extends in a generally radial direction at a third offset distance D 3  from the trailing edge  122 . In the illustrated example the treated zones  32 ,  34 , and  36  are depicted as straight lines. The offset distances D 1 , D 2 , D 3  will vary with the particular application but may generally be selected small enough to stop a crack before it exceeds the maximum length allowable in service, and large enough that the treated zone can be positioned in a reproducible manner (i.e. away from highly curved features). The offset distances D 1 , D 2 , D 3  are also preferably selected to place the treated zones outside any areas which are aerodynamically sensitive to dimensional changes. For example, in a compressor blade  110  constructed of a Ti 4-4-2 alloy, the offset distance D 1  may be about 0.25 mm (0.01 in.) to about 1.3 mm (0.05 in.) The treated zones  32 ,  34 , and  36  each have a width “W” which is selected to be wide enough to effectively stop crack propagation while minimizing the number of burnishing or peening passes required. In the illustrated example, the width W may be about 0.13 mm (0.005 in.) to about 0.25 mm (0.01 in.) The lengths of the treated zones D 1 , D 2 , and D 3 , denoted L 1 , L 2 , and L 3 , respectively, may be substantially greater than the width W in order to simplify the treatment process. 
   The treated zones  32 ,  34 , and  36  need not be linear, and other patterns such as curves may be used to suit a particular application. For example, as shown in  FIG. 4 , a compressor blade  210  may include a treated zone  132  having a shape comprising a series of angled segments or “chevrons”. The treated zone could also comprise a matrix of densely spaced point or dots. 
   One or more additional treated zones of residual compressive stress may be applied to the compressor blade  110  to further reduce crack propagation. For example, a supplemental treated zone  38  (see  FIG. 3 ) extends in a generally radial direction at a supplemental offset distance D 4  from the leading edge  120 , which is greater that the first offset distance D 1 . The length-to-width ratio of this supplemental treated zone  38  is similar to that of the treated zones described above. This stops the growth of any cracks which may extend past the first treated zone  36 . If desired, further treated zones (not shown) could be applied behind the supplemental treated zone  38 . 
   The zones of residual compressive stress may be applied by a number of methods. Examples of known suitable methods include laser shock peening (LSP), pinch peening, shot peening, low plasticity burnishing (LPB), or the use of a textured forming die. One preferred method is low plasticity burnishing, in which a normal force is applied to the compressor blade  110  using a stylus of a known type (not shown). The stylus is translated along the surface of the compressor blade  110  form the intended treated zone of residual compressive stress. The amount of cold-working applied to the compressor blade  110  during this process is of relatively little concern given the anticipated operating conditions. 
   In operation, the compressor blade  110  will be subjected to fatigue and damage that tends to cause cracking. The cracks initiate in “crack-prone” areas which may be relatively thin, which are exposed to debris impact, or which contain edges or geometric features that cause stress risers. Examples of such crack-prone areas include the leading edge  120 , tip  124 , and the trailing edge  122  as discussed above. Unchecked, these cracks would grow and extend further into the compressor blade  110 , until eventually the compressor blade failed in service or could not be economically repaired. 
   However, with the treatment as described herein, any cracks which initiate in the peripheral areas of the compressor blade  110  will eventually intersect one of the treated zones which resists crack propagation. The crack size will thus be limited to an acceptable value, which prevents catastrophic failure and improves the chances of successful repair of the compressor blade  110 . In particular, the presence of the treated zones as described above allows the service limits for crack propagation to be “opened up”—in other words, because the limit of crack propagation is rendered more predictable, a smaller margin of safety is required and the crack may safely be allowed to become larger than in an untreated component, before the treated component is removed from service. 
   The treatment described in this application may be used with newly made components. However, there is also significant repair activity on components such as compressor blades. The methods described in this application can enhance the performance of repaired blades. Repaired blades typically have their dimensions restored by welding, which often has inferior fatigue behavior as compared to the original components. By applying one or more treated zones to a repaired component as described herein, their fatigue and life resistance to crack propagation can be enhanced. 
   The foregoing has described fatigue- and damage-resistant components and methods for making such components. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation, the invention being defined by the claims.