Abstract:
A turbine engine composite blade attachment mechanism is provided, including a blade having an airfoil section, a neck section, and a root section. The root section includes a continuous loop integral with the blade and branching out from the neck section to completely surround a root insert. The blade includes a plurality of fibers which extend from the airfoil section through the neck and around the core. The root section and neck are enveloped by a jacket having a thickness which increases from the distal side surface of the loop to the neck portion, and has an inner contour substantially aligned with the proximal side surface contour of the loop and the core. The jacket and root are disposed inside a rotary disk cavity such that when tensile loading is applied, the neck width is maintained and compressive transverse loads may be applied to the fibers to prevent lamination thereof.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     This application is related to and claims priority to U.S. Provisional Patent Application Ser. No. 60/414,060, filed Sep. 27, 2002, entitled TAILORED ATTACHMENT MECHANISM FOR COMPOSITE AIRFOILS, the entirety of which is incorporated herein by reference. 
    
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     n/a 
     FIELD OF THE INVENTION 
     The present invention relates to turbomachinery and gas turbine engines, namely, composite blade attachment mechanisms and rotor assemblies. 
     BACKGROUND OF THE INVENTION 
     Gas turbine engine blades typically have dovetails or roots carried by a slot in a metal rotary disk or drum rotor. A typical blade  1  is displayed in  FIG. 1 , showing an airfoil section  2 , and root section  3 . The root section  3  provides the means by which the blade is attached to the rotor disk or other similar component of a gas turbine engine or comparable piece of turbomachinery. The blade  1  may also include an interface  4  between the airfoil  2  and root  3 , to conform to the rotor disk or other attachment mechanism. 
     Composite blades have many advantages over blades made with other materials, such as current metal alloys. They have a high strength to weight ratio, which allows for the design of low weight parts that can withstand the extreme temperatures and loading of turbomachinery. They can also be designed with parts with design features not possible with other materials (such as extreme forward sweep of compressor blading). A major drawback of composite blades is that their strength is essentially unidirectional. Despite having a relatively high uniaxial tensile strength, the composite materials are fragile and weak under compression or shear. However, in gas turbines, the blades are usually under extremely high tensile loads. Problems usually arise with regard to the transfer of such loads into the disk. Since the blades are often made of a laminated fiber or filament reinforced composite material, and the disks are typically made of metal, the transfer of loads between the two can lead to damage of the fibers, or even worse, delamination of the composite materials. 
       FIGS. 2   a - 2   c  illustrate the problem, where there is shown three separate views of an example of a composite blade root. First there is an unloaded blade  10   a  in  FIG. 2   a . Then, when tensile loading T is applied as shown, there is the case of loaded blade  10   b  in  FIG. 2   b , where shear stresses have caused failure of the root structure. Or, there is the case of loaded blade  10   c  in  FIG. 2   c , where the resulting stresses from tensile load T as applied to the blade from the surrounding disk cavity (not shown) have caused delamination of the blade. The challenge therefore, is to provide an optimum load path between the composite blade structural fibers and the surrounding disk. 
     Previously, one of the technology barriers for high performance composite blades has been designing an attachment scheme that would utilize the strengths of composite materials to prevent the failure illustrated in  FIGS. 2   b - 2   c . As demonstrated in  FIGS. 2   a - 2   b , a critically important area is the blade attachment region or “neck” portion  11  of the blade, where the thicker root transitions out to the relatively thin airfoil section above. It is this portion which tends to delaminate or otherwise fail, when the blade is loaded and the resulting stresses are applied to the root and interface between the root and disk. One reason for such failures is that the disk lugs tend to separate due to both the centrifugal loads acting on the disk and blade.  FIG. 2   d  illustrates another example of a blade  15  inserted into a disk  16 , under no loading. The disk lugs  17  around the neck  18  of the blade  15  define a gap G 0  that conforms to the shape of the blade  15 . In  FIG. 2   e , the blade of  FIG. 2   d  is shown under centrifugal axial loading, where the gap has increased in size to G L . Although this geometrical change in the disk geometry is slight (the dimensions portrayed in  FIGS. 2   d-e  are exaggerated for effect), it no longer conforms to the shape of the blade. The effect of this slight increase in gap induces transverse tension and/or shear stresses. 
     Since composite materials have little ability to handle transverse tension or shear loading, this will result in failure of the composite blade as in blade  10   c  once the intralaminar tension or shear stresses exceed the ultimate intralaminar stress capabilities of the composite. An example would be uni-directional Kevlar composite, having an ultimate intralaminar stress capability of about 6 ksi. 
     Since composite blades are very useful in gas turbine engines, it is desirable to provide a tailored attachment mechanism of composite airfoils, that both takes advantage of the relatively high tensile strength of composites, and minimizes the disadvantage of the relatively low shear and transverse tension strength of composites. 
     SUMMARY OF THE INVENTION 
     In a first embodiment of the present invention, a turbomachinery blade is provided, having a root insert, and an airfoil having a first end and a second end. The second end forms a loop enclosing the root insert. The loop includes two arms in apposition to form a neck distal to the second end from the root insert. A jacket is disposed around the neck to prevent separation of the opposing arms. 
     In another embodiment of the present invention, a gas turbine engine blade assembly is provided, having a composite blade having an airfoil portion, a neck portion and a root portion. The root portion includes a single root insert enveloped by two distinct arms integral to and extending from the neck portion and forming one continuous loop around the core. A jacket is disposed around the root portion and the neck portion. 
     In yet another embodiment of the present invention, a turbomachinery blade assembly, is provided, having a composite blade having a proximal airfoil portion and a distal neck portion. The blade includes a number of fibers which form a continuous loop integral to the blade distal to the neck portion, and a core insert having a distal surface contour and a proximal surface contour. The core insert is circumscribed by the loop. A jacket having an outer surface and an inner surface encloses the neck portion and the loop. The inner surface includes a surface contour substantially aligned with the proximal surface contour of the core insert. 
     And in another embodiment of the present invention, a gas turbine engine blade assembly is provided, including a root insert member and a blade member having a proximal airfoil portion, an intermediate neck portion, and a distal root portion. The distal root portion forms a continuous loop around the root insert member. In addition, a rotor member defines a cavity for receiving the blade member. The assembly also includes a jacket means for coupling the blade member to the rotor member, the jacket means being disposed inside the cavity around the intermediate neck portion and distal root portion of the blade member. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A more complete understanding of the present invention, and the attendant advantages and features thereof, will be more readily understood by reference to the following detailed description when considered in conjunction with the accompanying drawings wherein: 
         FIG. 1  illustrates a typical gas turbine engine blade; 
         FIGS. 2   a - 2   c  illustrate three views of a composite turbine blade root, showing the root under a tensile load; 
         FIG. 2   d  illustrates a blade inserted into a disk, under no loading; 
         FIG. 2   e  illustrates the blade of  FIG. 2   d , under centrifugal axial loading; 
         FIG. 3  is a schematic of the attachment principles employed in the present invention; 
         FIG. 4  is a cross section of a turbine blade root attachment mechanism in accordance with the principles of the present invention; 
         FIG. 4A  is a cross-section of only the turbine blade displayed in the turbine blade root attachment mechanism of  FIG. 4 ; and 
         FIG. 4B  is a cross-section of only the jacket displayed in the turbine blade root attachment mechanism of FIG.  4 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     As used herein, a “jacket” shall mean any outer layer type structure enveloping an inner structure, such as a sheath, cladding, ring, or coating. Also as used herein, a “shim” shall mean any piece of material, used to fill gaps, make something level, or adjust something to fit properly, or any material placed between two parts to make the parts continuously connect or fit. Both a jacket or shim may be made of any material. 
     Also as used herein, a “blade” shall mean any structure used in turbine engines, turbomachinery, aircraft, mobile or stationary powerplants, or other machines or devices which include a structure having surfaces which interact with fluid flows. Blades shall include vanes, airfoils, rudders, and any other aerodynamic, hydrodynamic or control surface structures. 
     Also as used herein, the term “circumscribe” shall refer to one element enclosing, surrounding, or abutting against another element to as to be substantially in contact with the outer surface of the circumscribed element, but not necessarily all of such outer surface. 
       FIG. 3  is a simplified schematic of the attachment principles employed in the present invention. Since composite blades are generally made up of a number of longitudinal layers having a relatively high tensile strength, the present invention takes advantage of this feature in tailoring an attachment mechanism for gas turbine blades.  FIG. 3  shows a schematic cross-section of a blade having a number of layers  20  extending along a longitudinal axis  21 . The layers  20  may be fibers, fiber composites, fiber sheets, or any group of elements which have a layered structure stacked substantially parallel to the longitudinal axis  21  of the blade. The layers are wrapped around an insert  25 . The insert  25  thereby provides the means by which the composite layers  20  are anchored to the rotor disk, such that a tensile load T, as shown, is transmitted throughout the length of the layers  20 , into cylinders  30  as insert  25  cannot pass between cylinders  30 . 
     However, this simplified representation neglects the increase in gap G shown in  FIG. 3 , which is due to the elastic deformation caused by centrifugal loading T on the blade and its surrounding disk (not shown). In practice it is necessary to maintain this gap unchanged in order to prevent the development of transverse tension or intralaminar shear in the composite blade. This gap can be maintained by means of a “jacket” that is designed to move together with the blade as the disk lugs move apart. The jacket can pinch or induce transverse compression (such as along arrows P shown in  FIG. 3 ) into the blade laminate, although this is not required providing gap G is maintained. 
       FIG. 4  illustrates the foregoing principles in practice in a gas turbine blade.  FIG. 4  is a cross section of a blade root attachment mechanism in accordance with the principles of the present invention. The lower portion of blade  100  is shown inserted into a rotor disk  101 . The blade  100  includes an airfoil portion  105 , a neck portion  106  and a root portion  107 , the root portion  107  having a single insert  108  enveloped by two distinct arms  110  integral to and extending from the neck portion  106  and forming one continuous loop  1111  around the core  108  as shown. A jacket  112  is disposed around the root portion  107  and the neck portion  106 . The entire blade is symmetrically oriented about a longitudinal axis  115  along which a tensile load T is applied in the direction shown. 
       FIG. 4A  is a cross-section of only the blade  100  displayed in the blade root attachment mechanism of  FIG. 4 , showing the various sections of the airfoil portion  105 , neck portion  106 , and root portion  107 . As clearly shown in  FIG. 4A , the loop  111  of the root portion  107  completely envelops and circumscribes the inner core  108 , which in this case is tear-drop shaped. The loop  111  includes a distal half  118  distal to the airfoil portion  105  and a proximal half  119  proximal to the airfoil. 
       FIG. 4B  is a cross-section of only the jacket  112  displayed in the blade root attachment mechanism of FIG.  4 . The jacket  112  is substantially U-shaped and includes a central portion  120  in apposition with the distal half  118  of the loop, as shown in FIG.  4 . The central portion  120  has a thickness that is substantially less than two end portions  122  disposed against opposite sides of the neck portion  106 . Each the two end portions  122  of the jacket  112  have a thickness which gradually increases from the thickness of the central portion  120  as the two end portions  122  extend over the proximal half  119  of the loop  111 , as shown in FIG.  4 . 
     In  FIG. 4 , tensile load T is indicated only an approximation of the resulting stresses and loads on the blade  100  due to a number of aerodynamic and centrifugal forces generated by the operation of the turbine engine. It is not meant to present an exact picture of blade loading, but does represent an overall tensile load that is generally applied to blades under these conditions. 
     As shown in  FIG. 4 , the rotary disk  101  defines a blade-receiving cavity  130  and an outer surface  135 . The cavity  130  is bounded by the outer surface  135 , wherein the neck portion  106 , the root portion  107 , and the jacket  112  are all disposed inside the cavity  130 . In one regard, jacket  112  functions as a shim that fills the space between the blade-receiving cavity  130  and the root and neck portions  107  and  106  of the blade  105 . This anchors the blade  105  to the rotor member  101 . For purposes used herein, the rotor disk member  101  need not be a rotary disk per se, but may be any element in a turbine engine or other similar piece of turbomachinery, at the compressor or turbine stage, wherein a blade such as blade  100  is attached. 
     The jacket  112  acts as more than just a shim or gap-filling element. In the embodiment shown in  FIG. 4 , the disk  101  circumscribes the jacket  112  such that when tensile loading T is applied to the blade  100  along longitudinal axis  115 , the jacket  112  maintains the gap G and potentially applies compressive loads to the gap G along the neck portion  106  normal to the longitudinal axis  115  as shown. It can be clearly seen in  FIG. 4  that the jacket  112  applies compressive loads G to the neck portion  106  at a point inside the cavity  130  beneath the outer surface  135  of the disk  101 . Furthermore, the disk  101  circumscribes the jacket  112 , and the jacket  112  circumscribes the root portion  107  and neck portion  106  in such a way that when tensile loading T is applied to the blade  100 , the jacket  112  hinders the expansion of the neck portion  106  normal to the longitudinal axis  115  at the compression point or gap defined by the arrows G as shown. Thus, as centrifugal loads T tend to widen gap G as the disk lugs  140  move apart, the jacket  112  deforms and deflects to conform to the outer contours of the root portion  107  and neck portion  106  of the blade, so as to maintain gap G. By maintaining this gap, the intralaminar shear and transverse tension forces in the blade are minimized or eliminated. 
     The particular dimensions of the jacket  112  and blade  100  are not limited to the ratios and proportions shown in  FIG. 4 , and are provided as an example only.  FIG. 4  illustrates one possible configuration, where jacket  112  has a thickness “F” at its central portion  120 . This thickness need not be very thin or very thick, and does not significantly affect the performance of the present invention. If the arms  110  have a thickness “t”, then the thickness of the jacket  112  will be such that it generally conforms to the contours of the inner surface of cavity  101  and the outer surface of the root portion  107 , where the radius of curvature “r” of the inner face of the upper portion  122  of jacket  112 , proximate the interface of the root portion  107  and neck portion  6 , is about equal to thickness t of the arms. The thickness t will vary depending on the particular composite blade, but radius r will generally be approximate to thickness t. Furthermore, the angle a shown as the slope of the outer surface of the jacket  112  at its thickest point will be in the range of 30 degrees plus or minus 10 degrees. This variation is required to accommodate differing rotor disk materials with differing stress capabilities (such as titanium, steel, etc.). 
     The composite may be made of a fiber reinforced laminar material. The fibers may be made of graphite, fiberglass, or Kevlar® brand synthetic fibers of the kind made by E. I. Du Pont De Nemours and Company. The resins may be epoxy or polyimide. The insert  108  may be round, elliptical, teardrop shaped as shown in  FIG. 4 , or some other suitable shape for restraining the blade in accordance with the present invention. The insert  108  and jacket  112  may be made of metal, metal alloy, plastic or polymer material. 
     The tailored attachment mechanism of the present invention is assembled and manufactured using industry standard techniques. 
     It will be appreciated by persons skilled in the art that the present invention is not limited to what has been particularly shown and described herein above. In addition, unless mention was made above to the contrary, it should be noted that all of the accompanying drawings are not to scale. A variety of modifications and variations are possible in light of the above teachings without departing from the scope and spirit of the invention, which is limited only by the following claims.