Abstract:
A blade fixing and blade mounting slot arrangement for a gas turbine engine has a mismatch fit along a portion of the length of the blade fixing and slot where contact stress would otherwise be maximal.

Description:
BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   The present invention relates to gas turbine engines and, more particularly, to blade and disk interfaces of such engines. 
   2. Background Art 
   Fan rotors can be manufactured integrally or as an assembly of blades around a disk. In the case where the rotor is assembled, the fixation between each blade and the disk has to provide retention against extremely high radial loads. This in turn causes high radial stress in the disk retaining the blades. 
   In the case of “swept” fans, the blades are asymmetric with respect to their radial axis. A significant portion of the weight of these blades is cantilevered over the front portion of the fixation, which causes an uneven axial distribution of the radial load on the fixation and disk. This load distribution causes high local radial stress in the front of the disk and high contact forces between the blade and the front of the disk. 
   Although a number of solutions have been provided to even axial distribution of stress in blades, such as grooves in blade platforms to alleviate thermal and/or mechanical stresses, these solutions do not address the problem of high local radial stress in the disk supporting the blades. 
   Some solutions have also been provided to reduce the increase of contact stress resulting in a non-zero broach angle of the blade, including the elimination of diagonally opposite portions of the load transfer interface which are less stressed. However, such solutions are not applicable to reduce the increased local contact stress produced by the asymmetry of “swept” fans. In addition, such solutions do not address the problem of high local radial stress in the disk supporting the blades. 
   Accordingly, there is a need for a blade and disk interface for a gas turbine engine fan producing reduced local contact stress and reduced local radial stress in the disk. 
   SUMMARY OF INVENTION 
   It is a general aim of the present invention to provide an improved blade and disk interface for a gas turbine engine. 
   It is also an aim of the present invention to provide a method for reducing a local contact stress between a disk and a blade. 
   It is a further aim of the present invention to provide a method for reducing a local radial stress in a bladed rotor disk assembly. 
   Therefore, in accordance with a general aspect of the present invention, there is provided a gas turbine engine rotor assembly comprising a rotor disk having a plurality of blade mounting slots circumferentially distributed about a periphery thereof for receiving complementary blade fixing portions of rotor blades, each of said blade mounting slots being bounded by a pair of opposed sidewalls extending longitudinally from a front side to a rear side of the rotor disk, and wherein a localized lateral play is provided between the sidewalls of each slot and the blade fixing portion of a respective one of the rotor blades along a longitudinal portion where contact stress is known to be maximal, said longitudinal portion being smaller than a length of the blade mounting slot and the blade fixing portion. In accordance with another feature of the present invention, the localized lateral play is at least partty provided by a region of reduced width in the blade fixing portion. 
   In accordance with a further general aspect of the present invention, there is provided a gas turbine engine rotor blade mountable in a blade retaining slot of a rotor disk, the rotor blade comprises a platform, an airfoil portion extending upwardly from said platform, a root depending downwardly from said platform and adapted for engagement in the blade retaining slot of the rotor disk, said root having a length extending from a front side to a rear side of the root, and wherein the root has a localized reduced width along a portion of the length thereof in a region where contact stress between the root and the slot is known to be high. 
   In accordance with a further general aspect of the present invention, there is provided a method for reducing high local stress transfer between a gas turbine engine blade fixing and a blade mounting slot of a rotor disk, the method comprising the steps of: a) determining which portion of a full length of the blade fixing and the blade mounting slot is subject to maximal contact stresses, and b) providing a mismatch fit in said portion of maximal stress. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     Reference will now be made to the accompanying drawings, showing by way of illustration a preferred embodiment of the present invention and in which: 
       FIG. 1  is a side view of a gas turbine engine, in partial cross-section; 
       FIG. 2  is a partial perspective view of a fan blade, showing a dovetail according to a preferred embodiment of the present invention; 
       FIG. 3  is a front view of the dovetail of  FIG. 2 , in cross-section, when engaged in a dovetail groove of a fan disk; and 
       FIG. 4  is a top view of the dovetail and dovetail groove of  FIG. 3 , in cross-section. 
   

   DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     FIG. 1  illustrates a gas turbine engine  10  of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan  12  through which ambient air is propelled, a multistage compressor  14  for pressurizing the air, a combustor  16  in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section  18  for extracting energy from the combustion gases. 
   Referring to  FIG. 2 , a part of a blade  32  of the fan  12 , which is a “swept” fan, is illustrated. Although the present invention applies advantageously to such fans, it is to be understood it can also be used with other types of conventional fans, as well as other types of rotating equipment requiring a smoother axial distribution of radial stress in the disk and in a disk to blade interface including, but not limited to, compressor and turbine rotors. 
   Referring to  FIGS. 2–3 , the fan  12  includes a disk  30  supporting a plurality of the blades  32  which are asymmetric with respect to their radial axis. Each blade  32  comprises an airfoil portion  34  including a leading edge  36  in the front and a trailing edge  38  in the back. The airfoil portion  34  extends radially outwardly from a platform  40 . A blade root  42  extends from the platform  40 , opposite the airfoil portion  34 , such as to connect the blade  32  to the disk  10 . The blade root  42  includes an axially extending dovetail  44 , which is designed to engage a corresponding dovetail groove  46  in the disk  10 . The airfoil section  34 , platform  40  and root  42  are preferably integral with one another. 
   As stated above, the asymmetry of the blade  32  causes a significant portion of the blade weight to be cantilevered over the front portion of the dovetail  44 . This creates an uneven axial distribution of the radial load on the dovetail  44  and disk  30 . Such a load distribution produces unacceptably high local radial stress in the front of the disk  30  and contact stress between the dovetail  44  and the front of the dovetail groove  46 . Each airfoil portion  34  has a center of gravity which is offset axially forwardly relative to the center of the blade fixing portion  44 . The blades are forward swept. 
   Referring to  FIGS. 3–4  and according to a preferred embodiment of the present invention, the high local stress in the front of the disk  30  and contact stress between the dovetail  44  and the front of the dovetail groove  46  are minimized or even cancelled by way of a relief mismatch or play  50  between the dovetail  44  and the dovetail groove  46  at the leading edge. The dovetail  44  is narrower at a front portion thereof, while the dovetail groove  46  has a constant section. This creates the mismatch  50  at the front, which minimizes or removes contact between the dovetail  44  and dovetail groove  46  at that point. As shown in  FIG. 3 , the mismatch  50  is preferably only present on the belly portion of the dovetail  44 . The rest of the front portion of the dovetail is at the larger thickness. The minimized contact brought by the mismatch  50  reduces the local contact stress as well as the local radial stress in the disk  30  for the leading edge. The radial stress is thus redistributed along the remainder of the contact surface in the axial direction. 
   In a preferred embodiment, the thickness difference between the narrow front portion of the dovetail  44  and the remainder of the dovetail  44  is approximately 0.010 inches. 
   It understood that the localized mismatch  50  can be created in alternative ways, such as by increasing the width of the dovetail groove  46  at the front while keeping the section of the dovetail  44  constant. The mismatch  50  can also be similarly created in alternative attachments such as bottom root profiles commonly known as “fir tree” engaging a similarly shaped groove in the disk  30 . 
   The mismatch  50  thus eliminates the unacceptably high local radial stress in the front of the disk  30  and contact forces between the dovetail  44  and the front of the dovetail groove  46  by minimizing or avoiding contact between the dovetail  44  and dovetail groove  46  in the region where the stress is maximal. 
   The embodiments of the invention described above are intended to be exemplary. Those skilled in the art will therefore appreciate that the foregoing description is illustrative only, and that various alternatives and modifications can be devised without departing from the spirit of the present invention. Accordingly, the present is intended to embrace all such alternatives, modifications and variances which fall within the scope of the appended claims.