Patent Publication Number: US-2010126018-A1

Title: Method of manufacturing a vane with reduced stress

Description:
BACKGROUND OF THE INVENTION 
     This invention relates generally to compressors, and more specifically to blades and vanes used in compression systems. As used herein, the term “fluid” includes gases and liquids. 
     In a gas turbine engine, air is pressurized in a compression module during operation. The air channeled through the compression module is mixed with fuel in a combustor and ignited, generating hot combustion gases which flow through turbine stages that extract energy therefrom for powering the fan and compressor rotors and generate engine thrust to propel an aircraft in flight or to power a load, such as an electrical generator. 
     The compressor includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades extending radially outward from a disk. More specifically, each rotor blade extends radially between a platform adjacent the disk, to a tip. A gas flowpath through the rotor assembly is bound radially inward by the rotor blade platforms, and radially outward by a plurality of shrouds. 
     The stator assembly includes a plurality of circumferentially spaced apart stator vanes or airfoils that direct the compressed gas entering the compressor to the rotor blades. The stator vanes extend radially between an inner band and an outer band. A gas flowpath through the stator assembly is bound radially inward by the inner bands, and radially outward by outer bands. The vanes are typically made in arcuate segments with arcuate outer and inner band segments each having one or more vane airfoils per segment. The segments are conventionally joined together to collectively form a complete 360 Degree stator vane stage. The stator vane segments are mounted within a compressor casing. A vane stage comprises a plurality of circumferentially arranged vane segments, with each segment having a plurality of vane airfoils extending between an arcuate inner band and an arcuate outer band. 
     In some designs, the vane segments are supported solely at its outer band since a conventional annular seal member is disposed between rotor stages, preventing stationary support of the inner band as well. Accordingly, the vane airfoils in these vane segments are cantilevered from the outer band support which creates bending moments due to the fluid flowing between the vanes which must be suitably reacted or accommodated through the outer band. The bending moments in some of these airfoils may be significant since they are supported solely at their outer band, with the inner band thereof being unsupported. 
     During engine operation, the gas flow through the flow path induces mechanical, thermal, and aerodynamic loads on the airfoils. Some of these loads are transmitted by the airfoils to support structures in the engine through the outer bands that are coupled to the airfoil and reacted by the support structures. In some designs, the inner bands may also, similarly, transmit some of the loads applied on the airfoils by the gas flow and reacted by other support structures. Within at least some conventional gas turbine engines, the stresses in the airfoil near the interface with outer band and near the support structures may become large enough to cause distress in the airfoil. Under sufficiently large stresses, cracking may occur in the airfoils within the vane sector near support structure locations that react the loading applied to the vane airfoils in the vane sector by the gas flow. Designing with additional thickness at these high locations may not be possible for several reasons, such as, aerodynamic considerations, flow modifications, additional weight and changed dynamic characteristics of the vanes and/or excessive leakages in the vane sectors. 
     Accordingly, it is would be desirable to have an airfoil assembly having features that reduce the stresses in the airfoil near the interface with the bands that support the airfoil while minimizing leakages in the vane sector. It would be desirable to have a method of manufacturing an airfoil assembly having features that reduce the stresses in the airfoil and having an interface with the bands that support the airfoil. 
     BRIEF DESCRIPTION OF THE INVENTION 
     The above-mentioned needs may be met by exemplary embodiments which provide a method of manufacturing an airfoil assembly, the method comprising the steps of supplying an airfoil and a band, creating a support aperture in the band capable of receiving a portion of the airfoil, creating a slot in the band for reducing stress in the airfoil, and bonding the airfoil with the band. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The subject matter which is regarded as the invention is particularly pointed out and distinctly claimed in the concluding part of the specification. The invention, however, may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
         FIG. 1  is a cross-sectional view of an exemplary gas turbine engine assembly comprising a compression system according to an exemplary embodiment of the present invention. 
         FIG. 2  is an axial cross-sectional view of a portion of a compression system having an airfoil assembly according to an exemplary embodiment of the present invention. 
         FIG. 3  is a perspective view of an airfoil assembly according an exemplary embodiment of the present invention. 
         FIG. 4  is a perspective view (looking radially inward) of a portion of the exemplary airfoil assembly shown in  FIG. 3 . 
         FIG. 5  is a perspective view (looking radially outward) of a portion of the exemplary airfoil assembly shown in  FIG. 3 . 
         FIG. 6  is a schematic view of an exemplary slot in the outer band near the trailing edge of an airfoil in an airfoil assembly according to an exemplary embodiment of the present invention. 
         FIG. 7  is a flow chart showing an exemplary embodiment of a method of manufacturing an airfoil assembly. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  shows a cross-sectional view of a gas turbine engine assembly  10  having a longitudinal axis  11 . The gas turbine engine assembly  10  includes a core gas turbine engine  12  that includes a high-pressure compressor  14 , a combustor  16 , and a high-pressure turbine  18 . In the exemplary embodiment shown in  FIG. 1 , the gas turbine engine assembly  10  also includes a low-pressure turbine  20  that is coupled axially downstream from core gas turbine engine  12 , and a fan assembly  22  that is coupled axially upstream from core gas turbine engine  12 . Fan assembly  22  includes an array of fan blades  24  that extend radially outward from a rotor disk  26 . In the exemplary embodiment shown in  FIG. 1 , engine  10  has an intake side  28  and an exhaust side  30 . In the exemplary embodiment, gas turbine engine assembly  10  is a turbofan gas turbine engine that is available from General Electric Company, Cincinnati, Ohio. Core gas turbine engine  12 , fan assembly  22 , and low-pressure turbine  20  are coupled together by a first rotor shaft  31 , and compressor  14  and high-pressure turbine  18  are coupled together by a second rotor shaft  32 . 
     In operation, air flows through fan assembly blades  24  and compressed air is supplied to a compression system  90  that includes a high-pressure compressor  14 . The air discharged from fan assembly  22  is channeled to compressor  14  wherein the airflow is further compressed and channeled to combustor  16 . Products of combustion from combustor  16  are utilized to drive turbines  18  and  20 , and turbine  20  drives fan assembly  22  via shaft  31 . Engine  10  is operable at a range of operating conditions between design operating conditions and off-design operating conditions. 
       FIG. 2  is an axial cross-sectional view of a portion of the compression system  90  having a high-pressure compressor  14  including an exemplary embodiment of an airfoil assembly  48  according to an exemplary embodiment of the present invention. Core airflow, shown as item  15  in  FIG. 2 , flows in the annular flow path  17  of the compression system  90 . The compression system  90  includes a plurality of compression stages  40  wherein each stage  40  includes a row of circumferentially spaced rotor blades  27  and a stator assembly  42 . The stator assembly  42  includes a row of circumferentially spaced arcuate stator vane sectors  44  that are arranged circumferentially around the longitudinal axis  11 . Each stator vane sector  44  comprises a vane airfoil  50  coupled to a band  61  that supports the airfoil  50 . In the exemplary embodiment shown in  FIG. 2 , the vane airfoil  50  is coupled to an arcuate outer band  60  located near the radially outer portion of the airfoil  50 , and an arcuate inner band  80  located at the radially inner portion of the airfoil  50 , referred to herein with respect to the longitudinal axis  11 . 
     In the exemplary embodiment shown in  FIG. 2 , the airfoil  50  is supported only at the outer band  50  wherein the inner band  80  supports a seal  81  which interfaces with a conventional annular seal  83  located between two adjacent rotors. Conventional sealing shrouds or segments  81  are suitably attached to the radially inner bands  80  of the vane sectors  44  to cooperate with labyrinth teeth extending from annular seal  83  of the compressor rotor for effecting interstage seals. In this configuration, the individual vane sectors  44  are mounted to the outer casing  25  solely by their outer bands, with the vane airfoils  50  and inner bands  80  being suspended therefrom. The vane sector  44  comprises a contoured slot  70  (see  FIG. 3 ), located in the outer band  60 , capable of reducing a stress in the airfoil  50  and lowering leakages, as described subsequently herein. The compression system  90  further comprises a compressor rotor  23  having rotor blades  27  typically supported by rotor disks  26 , and are coupled to rotor shaft  32 . 
     The compression system  90  further comprises a casing  25  that surrounds the compressor  14  and supports the vane sectors  44  in the stator assemblies  42 . Each vane sector  44  comprises an arcuate forward rail  62  located axially forward from the airfoil  50 , and an arcuate aft rail  64  located axially aft from the airfoil  50 , that engage with a corresponding rails in adjacent components or the casing  25 . The loads that are experienced by each vane sector  44  are reacted with the support structures in the casing  25  through lugs  34 ,  35 ,  36  (see  FIG. 3 ) located on each vane sector  44 . During assembly, the individual vane sectors  44  are axially inserted into respective ones of the casing halves by engaging the forward and aft rails  62 ,  64  with the corresponding forward and aft grooves in the support structures or adjacent components. In the exemplary embodiment shown in  FIG. 2 , there are 10 arcuate vane sectors  44 , each having an arcuate sector angle of 36 degrees. Each vane segment or sector  44  is slid circumferentially in turn into the casing slot until all of the vane sectors in each casing half are assembled. The two casing halves are then assembled together so that the vane sectors  44  in each casing slot define a respective annular row of adjoining vane sectors  44  for each compression stage  40 . 
       FIG. 3  shows a vane sector  44  having an exemplary embodiment of the present invention of an airfoil assembly  48  comprising a plurality of airfoils  50  supported by a band  61 . In the exemplary embodiment shown in  FIG. 3 , the airfoil assembly has fourteen airfoils  50  arranged in a circumferential row and supported by an arcuate outer band  60  and an arcuate inner band  80 . The arcuate vane sector  44  shown in  FIG. 3  has a sector angle of 36 degrees. In alternate embodiments, the airfoil assembly may have a different number of airfoils and may be supported at only one end of the airfoil  50 , by an outer band  60  or by an inner band  80 . The vane sector  44  has a forward rail  62  located at an axially forward end, and an aft rail  64  located at an axially aft end of the vane sector  44 . The fwd rail  62  and aft  64  have an arcuate shape and engage with adjacent components as shown in  FIG. 2  to provide a portion of the support for the vane sector  44 . The mechanical, thermal and aerodynamic loads and moments from the airfoil assembly  48  are transmitted to the outer band  61  and these loads are reacted with the support structures, such as, for example, casing  25  (See  FIG. 2 ) through support lugs. In the exemplary embodiment shown in  FIG. 3 , three support lugs are used, a first-end lug  35 , a second-end lug  36  and a center lug  34 . These support lugs  34 ,  35 ,  36  engage with the support structures, such as the casing  25 , and react the loads and moments from the airfoil assembly  50 . The airfoil  50 , outer band  60  and inner band  80  may be made from a known material, such as, for example, Titanium alloys, Nickel and Cobalt based alloys. In the exemplary embodiment shown in  FIG. 3 , the airfoil  50  is made from Inconel 718, the outer band  60  is made from Inconel 718 and the inner band  80  is made from Inconel 625. The airfoil  50 , outer band  60  and inner band  80  are made by a known processes, such as for example, casting, forging or forming. 
     The mechanical, aerodynamic and thermal loads experienced by the individual airfoils  50  in the airfoil assembly  48  create stresses in the airfoils  50 . The peak stresses experienced by individual airfoils  50  in the vane sector  44  are not usually the same. Conventional analysis using known analytical methods has shown that the peak stresses in each airfoil  50  varies, depending on the circumferential location of the airfoil  50  the array and its location with respect to the support lugs  34 ,  35 ,  36 . Peak stress variations in the airfoils may be as high as 25%. Further, it has been seen that, the peak stress occurs in each airfoil  50  at the trailing edge, near the interface with the supporting band  61 , such as the outer band  60 . Full sector analysis of a conventional vane sector using known finite element analytical methods has confirmed that high mean stresses occur in the airfoils located in the center of the sector at the airfoil trailing edge near the interface with the outer band. To meet aerodynamic requirements in compression systems, airfoils  50  are relatively thin at the leading edge  51  and the trailing edge  52 , and peak stresses are preferably to be avoided at these locations to prevent cracking. 
     In the exemplary embodiment of the present invention shown in  FIGS. 3-6 , peak stress at the airfoil  50  trailing edge  62  at the interface with the outer band  60  is eliminated by removing the load path through the trailing edge  52  into the outer band  60 . This is accomplished by having a slot  70  in the outer band  60  near the trailing edge  52 . Cutting the load path at the thin trailing edge  52  of the airfoil  50  forces the airfoil loads and moments to be reacted further forward towards the middle of the airfoil  50  where the airfoil thickness is greater, thereby reducing peak stresses. Full sector analysis, using known finite element methods, of the exemplary embodiment of the present invention shown in  FIGS. 3-6  having a slot  70  in the outer band  60  near each airfoil  50  trailing edge  52  has confirmed that the peak stresses in each airfoil  50  is reduced by about between 35% and 45%. Such large reductions in peak stresses in airfoils significantly reduce the possibility of any cracking at the trailing edges and improves the durability of the vanes. It will be apparent to those skilled in the art that, in other embodiments of the present invention, similar approach of having a slot, such as slot  70  in  FIGS. 3-6 , may also be used near the leading edge  51  of the airfoil  50 , and the slot  70  may be located on the outer band  60 , inner band  80  or both. These and other modifications in implementation are within the scope of the present invention. 
     It will be appreciated that the fluid flowing through the flow path  17  has a relatively higher pressure, and there will be some leakage of that fluid through the slot  70 . In another aspect of the present invention, fluid leakage from the flow path  17  through the slot  70  in the band  61  is minimized. This is accomplished by having a slot-contour shape  72  for the slot  70  that generally corresponds to the airfoil-contour shape  56  of the airfoil near the location of the slot, as shown in  FIGS. 4-6 .  FIG. 4  is a perspective view looking radially inward, and  FIG. 5  is a perspective view looking radially outward, of a portion of the exemplary embodiment of the present invention of airfoil assembly  48  shown in  FIG. 3 . 
     Referring to  FIG. 6 , the geometry of a slot  70  according to an exemplary embodiment of the present invention is described. The airfoil  50  has a first end  57  (see  FIGS. 2 ,  6 ) which extends through a first aperture  78  in the outer band  60 . The first aperture corresponds generally to the shape of the airfoil  50  at its first end  57 , and further comprises a slot  70  near an airfoil edge, such as, for example, the trailing edge  56 . There is a small braze gap (not shown) between the first aperture  78  and the airfoil  50  that is later filled by braze material during manufacturing described subsequently herein. The braze gap is typically in the range of between about 0.000 inches and 0.010 inches. In one aspect of the present invention, the slot-contour shape  72  is designed to generally correspond to the airfoil edge contour  56  thereby minimizing leakage of fluid through the slot  70 . For example, as shown in  FIG. 6 , the slot-contour has a radius ‘P’ (item  75 ) that locates the slot from the airfoil edge. Further, the slot-contour  72  extends along the chordal direction of the airfoil and transitions with a radius ‘R’ (item  74 ) to the first aperture  78  near the airfoil. The resulting slot-contour shape  72  has a width ‘W’ (item  73 ) from the airfoil surface. It has been found that such a contour minimizes the fluid leakage from the flow path and also reduces the airfoil peak stresses by eliminating the load path at the airfoil edges. In the preferred embodiment of the present invention, the radius ‘P’ (item  75 ) has a value of about 0.046 inches, radius a ‘R’ (item  74 ) has a value of about 0.015 inches, and the width ‘W’ (item  73 ) has a value of about 0.040 inches. It has been found that the slot  70  having a slot-contour shape  72  as described above has a 32% reduction in the recirculation flow leakage area as compared to a simple circular hole, and effecting a significant reduction in fluid leakage through the slot  70 . It will be apparent to those skilled in the art that, in other embodiments of the present invention, it is possible to have other variations or modifications in the slot-contour geometry described above. For example, it may be possible to have an offset for the slot-contour  72  with respect to the airfoil-contour shape  56 . These and other modifications in the contour geometries and shapes are within the scope of the present invention. 
       FIG. 7  shows a flow chart schematically showing an exemplary embodiment of a method  500  for manufacturing an airfoil assembly  48  such as, for example, shown in  FIG. 3 . The method  500  comprises the step  510  of supplying the components, such as for example, a band  61  and an airfoil  50 . In some applications, the components may comprise a plurality of airfoils  50 , an outer band  60  and an inner band  80 . The airfoil  50 , outer band  60  and inner band  80  may be made from a known material, such as, for example, Titanium alloys, Nickel and Cobalt based alloys. In the exemplary embodiment shown in  FIG. 3 , the airfoil  50  is made from Inconel 718, the outer band  60  is made from Inconel 718 and the inner band  80  is made from Inconel 625. The airfoil  50 , outer band  60  and inner band  80  are made by a known processes, such as for example, casting, forging or forming. 
     The method  500  further comprises the step  515  of creating a first aperture  78  in the outer band  60  that is capable of receiving a portion of the airfoil  50 . The aperture  78  has a contour that corresponds with the contour of the portion of the airfoil  50  that will extend through it during assembly as described herein. For example, in the exemplary embodiment shown in  FIG. 3 , the aperture  78  has an airfoil shaped contour that corresponds with the first end  57  of the airfoil  50 . The aperture  78  is slightly larger than the airfoil first end  57  such that it has a small gap between the airfoil  50  and the outer band  60  for receiving a braze material. This braze gap is preferably between 0.000 inches and 0.010 inches. The aperture  78  can be created in the outer band  60  using known methods, such as by laser machining and electro discharge machining (EDM). 
     The method  500  further comprises the step  525  of creating a slot  70  in the outer band  60  for reducing stress in the airfoil  50  during operation of the airfoil assembly  48 . The location and contour shape of the slot  70  is selected based on stress analysis of the airfoil assembly  50  using known analytical methods. An exemplary embodiment of the slot  70  is shown in  FIGS. 3-6  and described previously herein. The slot-contour shape  72  of the slot  70  in the outer band  60  is created using known methods, such as by laser machining and electro discharge machining (EDM). In some applications, it is possible to combine step  515  and step  525  described above to perform the machining operations for creating the aperture  78  and slot  70 . 
     The method  500  further comprises the optional step  530  of applying a bond preparation to the components to be brazed subsequently. For example, the optional bond preparation comprises applying a known surface preparation to the airfoil  50 , the outer band  60  and the inner band  80 . In the exemplary embodiment shown in  FIG. 3 , the bonding surface preparation comprises nickel plating of the airfoil  50 , outer band  60  including the slot  70  and the inner band  80  using known materials, such as for example, AMS 2403, and using known plating methods. 
     The method  500  further comprises the step  535  of locating the airfoil  50  in the aperture  78 . As described previously, a portion of the airfoil  50 , such as the first end  57  (see  FIGS. 3-5 ), extends through the aperture  78 . A portion of the airfoil  50 , such as the first end  57 , is located in the aperture  78  such that there is a small gap between the airfoil and the outer band  60 . Preferably, this gap (“braze gap”) between the airfoil and the outer band is uniform and has a value between 0.000 inches and 0.010 inches. In some applications, the step  535  may comprise optionally tack welding some of the components forming the airfoil assembly  48 . The airfoil  50  and the outer band  60  and the inner band  80  may be optionally tack welded in position prior to brazing using known welding methods, such as for example, resistance welding, spot welding, seam welding and projection welding. 
     The method  500  further comprises the step  540  of applying a brazing material in the areas of interface between the airfoil  50  and the outer band  60  and, if applicable, between the airfoil  50  and the inner band  80 . A suitable braze alloy is selected using known methods, depending on the material compositions of the airfoil  50 , outer band  60  and the inner band  80 . In the exemplary embodiment shown in  FIG. 3 , the airfoil  50  and the outer band  60  are made from Inconel 718 and the inner band is made from Inconel 625. The braze material used is a commercially available braze alloy, AMS 4777. 
     The method  500  further comprises the step  545  of brazing the airfoil  50 , outer band  60  and the inner band  80  to form the airfoil assembly  48 . The components of the airfoil assembly  48  is conventionally fixtured so that they can be conventionally brazed together using known brazing materials such as AMS4777. During brazing, the assembly is heated using known methods. The brazing material is suitably melted and spread by capillary action within the braze gap between the airfoil  50  and the outer band  50  and between airfoil  50  and the inner band  80 . During brazing, care is taken to keep the slot  70  substantially free of brazing material. This can be accomplished, for example, by suitably orienting the assembly such that braze material does not flow into the slot  70 . Upon cooling of the assembly, the braze material solidifies and rigidly joins the airfoil  50  with the outer band  60  and the inner band  80  to form the airfoil assembly  48 . In the exemplary embodiment shown in  FIG. 3 , the brazing operations are preferably performed in a vacuum furnace known in the art. After brazing is complete, the method  500  may further optionally comprise the step of heat treating the airfoil assembly  48  using known methods. For example, known age-heat treatment may be applied the airfoil assembly  48  to restore metallurgical properties. 
     As used herein, an element or step recited in the singular and proceeded with the word “a” or “an” should be understood as not excluding plural said elements or steps, unless such exclusion is explicitly recited. When introducing elements/components/steps etc. of designing and/or manufacturing airfoil assembly  48 , vane sector  44  or compression system  90  described and/or illustrated herein, the articles “a”, “an”, “the” and “said” are intended to mean that there are one or more of the element(s)/component(s)/etc. The terms “comprising”, “including” and “having” are intended to be inclusive and mean that there may be additional element(s)/component(s)/etc. other than the listed element(s)/component(s)/etc. Furthermore, references to “one embodiment” of the present invention are not intended to be interpreted as excluding the existence of additional embodiments that also incorporate the recited features. 
     Although the methods and articles such as vanes, outer bands, inner bands and vane segments described herein are described in the context of a compressor used in a turbine engine, it is understood that the vanes and vane segments and methods of their manufacture or repair described herein are not limited to compressors or turbine engines. The vanes and vane segments illustrated in the figures included herein are not limited to the specific embodiments described herein, but rather, these can be utilized independently and separately from other components described herein. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to make and use the invention. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims. 
     While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein, and it is, therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention.