Patent Abstract:
A stator joint for a gas turbine engine has a center axis, and a shroud having a radial wall facing substantially radially with respect to the center axis. A slot wall defines in-part a slot in the shroud. A relief wall defines a relief area of the slot. The relief wall extends between the radial wall and the slot wall. A vane has an airfoil and a lug extending into the slot. A flowable attachment material is disposed in the relief area for engagement of the vane to the shroud. A vane assembly and a gas turbine engine are also disclosed.

Full Description:
BACKGROUND 
     This disclosure relates to a joint between a vane and shroud, for use in a gas turbine engine. 
     Gas turbine engine vane assemblies typically include vanes having airfoils mounted between two rings or partial rings (shrouds) that form a flowpath for the gas turbine engine. The vanes are typically brazed to the shrouds, and may have lugs at radial ends received in slots in the shrouds. 
     One application for such an assembly is in a compressor. Generally, there are vane assemblies intermediate rotor stages in the compressor. 
     In the prior art, the lugs are inserted into slots in radially inner and outer shrouds. Some flowable attachment material, such as a brazing material, has typically been deposited between the lugs and the slots. There have been two basic types of this structure used. In a first type, the lugs extend radially inward of the outer shroud and radially outwardly of the inner shroud. These enlarged lugs provide a dam preventing the flowable attachment material from extending to locations on the airfoil. However, these enlarged lugs also present an obstruction to a desired air flow cross-sectional area between the airfoils. 
     It is also known to have the lugs not extend radially beyond the shroud walls. With this structure, the flowable attachment material could move beyond the lug and toward surfaces of the airfoil which can cause damage to the vanes. 
     These and other features of this application will be best understood from the following specification and drawings, the following of which is a brief description. 
     SUMMARY 
     In a featured embodiment, a vane and shroud for a gas turbine engine include a center axis with a shroud having a radial wall facing substantially radially with respect to the center axis. A slot wall defines a slot in the shroud. A relief wall defines a relief area of the slot and extends between the radial wall and the slot wall. A vane has an airfoil and a lug extending into the slot. A flowable attachment material is disposed in the relief area for engagement of the vane to the shroud. 
     In another embodiment according to the previous embodiment, the slot is larger than the lug, such that said flowable attachment material is also disposed between the lug and the slot wall. 
     In another embodiment according to any of the previous embodiments, the relief area has a triangular cross-section. 
     In another embodiment according to any of the previous embodiments, the relief area has a curved cross-section. 
     In another embodiment according to any of the previous embodiments, the relief area has a rectangular cross-section. 
     In another embodiment according to any of the previous embodiments, the lug merges into a transition section which curves circumferentially inwardly from the lug to the airfoil. 
     In another embodiment according to any of the previous embodiments, the radial wall is generally radially aligned with a radial extent of the transition section which is most adjacent to the radial wall. 
     In another embodiment according to any of the previous embodiments, a depth of the relief wall is defined to a point most radially distant from a surface of the radial wall facing the center axis. A radial wall thickness is defined for the shroud adjacent to the relief area, and a ratio of the depth to the radial wall thickness is between about 0.2 and 0.6. 
     In another featured embodiment, a vane assembly includes a circumferentially extending outer shroud and a circumferentially extending inner shroud centered on a center axis. A plurality of vanes is positioned radially between the inner and outer shrouds. A joint is between the vanes and at least one of the inner and outer shrouds. The at least one shroud has a radial wall facing substantially radially with respect to the center axis. A plurality of slots is in the at least one shroud. Slot walls define the slots in the at least one shroud. A relief wall defines a relief area of the slots and extends between the radial wall and the slot wall. The vanes have an airfoil and a lug extending into one of the slots. A flowable attachment material is disposed in the relief area for engagement of the vane to at least one of the inner and outer shrouds. 
     In another embodiment according to the previous embodiment, the slot is larger than the lug, such that said flowable attachment material is also disposed between the lug and the wall. 
     In another embodiment according to any of the previous embodiments, the relief area has a triangular cross-section. 
     In another embodiment according to any of the previous embodiments, the relief area has a curved cross-section. 
     In another embodiment according to any of the previous embodiments, the relief area has a rectangular cross-section. 
     In another embodiment according to any of the previous embodiments, the lug merges into a transition section which curves circumferentially from the lug to the airfoil. 
     In another embodiment according to any of the previous embodiments, the radial wall is generally radially aligned with a radial extent of the transition section which is most adjacent to the radial wall. 
     In another embodiment according to any of the previous embodiments, the at least one shroud is the outer shroud. 
     In another embodiment according to any of the previous embodiments, a depth of the relief area is defined to a point most radially distant from a surface of the radial wall facing the center axis. A radial wall thickness is defined for the shroud adjacent to the relief area, and a ratio of the depth to the radial wall thickness is between about 0.2 and 0.6. 
     In another featured embodiment, a gas turbine engine has a compressor section, a combustor section and a turbine section. The compressor section and the turbine section are defined by a plurality of rotor stages and a plurality of vane assemblies positioned between adjacent ones of the rotor stages. At least one of the vane assemblies has a circumferentially extending outer shroud and a circumferentially extending inner shroud centered on a center axis. A plurality of vanes is positioned radially between the inner and outer shrouds. A joint is between the vanes and at least one of the inner and outer shrouds such that the at least one shroud has a radial wall facing substantially radially with respect to the center axis. A plurality of slots is in the at least one shroud. Slot walls define the slots in the at least one shroud. A relief wall defines a relief area of the slots and extends between the radial wall and the slot wall. The vanes have an airfoil and a lug extending into one of the slots. A flowable attachment material is disposed in the relief area for engagement of the vane to at least one of the inner and outer shrouds 
     In another embodiment according to the previous embodiment, the at least one of vane assemblies is in the compressor section. 
     In another embodiment according to any of the previous embodiments, the at least one shroud is the radially outer shroud. 
     These and other features may be best understood from the following drawings and specification. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic, cross-sectional side view of an embodiment of a gas turbine engine. 
         FIG. 2  shows a vane assembly for use in  FIG. 1 . 
         FIG. 3  is an enlarged view of the area inside the box 3 of  FIG. 2 . 
         FIG. 4A  illustrates a first embodiment of a vane and shroud. 
         FIG. 4B  shows a detail of  FIG. 4A . 
         FIG. 5A  is a view of a second embodiment vane and shroud. 
         FIG. 5B  shows a detail of  FIG. 5A . 
         FIG. 6A  is a view of a third embodiment vane and shroud. 
         FIG. 6B  is a detail of  FIG. 6A . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is shown herein is a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air in a bypass flowpath B and also drives air along a core flowpath C for compression and communication into the compressor section  24 , and combustor section  26 , then expansion through the turbine section  28 . Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures, and ground-based power generating engines. 
     The engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure (or first) compressor section  44  and a low pressure (or first) turbine section  46 . The inner shaft  40  is connected to the fan  42  through a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and high pressure (or second) turbine section  54 . A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  supports one or more bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
     The core airflow C is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     As known, the compressor sections  44  and  52  include rotating blade stages  18  and intermediate vane assemblies  19 . Both of these structures are shown schematically. It is known that the blades  18  typically rotate with a rotor. The vanes  19  typically are provided in the form of a ring, with vanes extending radially between an inner shroud and an outer shroud. The turbine sections  44  and  46  also have blades  18  and vane assemblies  19 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture  48  is an epicyclic gear train, such as a star gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. In addition, gas turbine engines for other applications such as land-based power generation turbines may also benefit from the teachings of this application. 
     A vane assembly  150  for use in a compressor section of a gas turbine engine is illustrated in  FIG. 2 . As seen in  FIG. 2 , inner shroud  60  and outer shroud  70  may be segmented for easier installation within the engine  20 . Only a circumferential portion of the vane assembly  150  is shown. As known, a plurality of segments  65  are connected together, and typically form a full ring. Inner shroud  60  has a plurality of slots, and an outer shroud  70  has a plurality of slots  75 . A plurality of vanes  17  are disposed in the slots in the inner shroud  60  and the outer shroud  70 . 
     Referring to  FIG. 3 , an outer portion of the vane  17  is disposed in a slot  75  in outer shroud  70 . Typically, lugs  90  (e.g., see lug  90  in  FIG. 4A ) are used to attach the vanes  17  to the outer shroud  70 . Some flowable attachment material, which is appropriate for securing the respective metals of the vane  17  and the shrouds  60  and  70  may be utilized. Various brazing materials are known, and would be appropriate for the teachings of this application. 
     Referring to  FIG. 4A , lugs  90  are shaped to generally fit into respective slots  75 . Vane  17  has a curved transition section  100  formed to merge an airfoil  80  into lug  90  by curving circumferentially inwardly. A maximum stress area  105  exists where the transition section  100  blends in the airfoil  80 . 
     As is clear, the slot  75  is larger than lug  90 , so there is clearance. A brazing material  120  is disposed in the clearance, and used to secure the lugs  90  to the shroud  70 . Material  120  does not substantially contact area  105  during the brazing because of chamfers or relief areas  101  formed by a relief wall  135  formed in radially inner wall  140 . This will be explained below. This lack of contact prevents fatigue at area  105  and thereby extends the life of the vane assembly  150 . At the same time, the lightweight and aerodynamic configuration does not cause flow obstruction that could otherwise reduce engine efficiency. 
     In this embodiment, an outer extent  102  of the transition section  100  may be in register (i.e., aligned) with inner wall  140  of the outer shroud  70  to not obstruct air flow. Alternatively, the outer extent  102  of the transition section  100  may be radially outwardly of the inner wall  140 , as this would also eliminate obstruction to air flow. The slot  75  is generally defined by the slot walls  145 . As can be seen, the relief walls  135  are formed as chamfers. The relief wall  135  extends in a direction with a radially outer component, and a component in a circumferential direction, such that the resulting shape is triangular, or a chamfer. The relief area  101  provides an area for the brazing material  120  to flow when it is heated, thereby minimizing a possibility that the brazing material  120  might reach the transition section  100  or the maximum stress area  105 . 
       FIG. 4B  shows shroud  70  has a wall thickness t 1 . A radially outermost point  200  of the relief wall  135  extends to a distance d 1  away from the inner wall  140 . In embodiments, t 1  may be between 0.08-0.1″ (0.20-0.25 cm). Notably, t 1  may be the same across the embodiments of  FIGS. 4B and 5B . In such embodiments, d 1  may be between 0.02-0.05″ (0.05-1.3 cm). A ratio of d 1  to t 1 , or a ratio of the deepest portion of the relief area to the wall thickness of the shroud may be between about 0.2 and 0.6. 
       FIG. 5A  shows another embodiment wherein the relief area  201  is formed by a curved relief wall  235 , which in this embodiment may be a circular section. The relief area  201  will function much like the relief area in the  FIG. 4A  embodiment to provide a space for the flowable material to move, such that it does not move onto the transition section  100 . 
       FIG. 5B  shows the wall thickness t 1  of the shroud  70 , and that the depth of the relief area  235  is formed at a radius r 1 . In embodiments, r 1  may be between 0.02-0.05″ (0.05-1.3 cm). Thus, a ratio of r 1  to t 1  may be between about 0.2 and 0.6. 
       FIG. 6A  shows another relief area embodiment  301  wherein the shape of the relief wall  335  is generally rectangular. Again, this shape will provide space to receive the flowable attachment material. 
       FIG. 6B  shows a detail of the relief wall  335 . The distance d 2  to the deepest portion of the relief wall, measured away from the wall  140 , was between 0.02-0.05″ (0.05-1.3 cm). Again, a ratio of d 2  to t 1  may be between about 0.2 and 0.6. 
     The distance t 1  could be defined as the radial wall thickness of the shroud measured adjacent to the relief area. The dimensions d 1 , d 2 , and r 1  could all be defined as a depth of the relief area measured to a point most radially distant from an inner surface of the wall  140 . 
     The relief areas work generically to limit flowable attachment material from flowing into the transition section  100  since the flowable attachment material maintains a relatively high viscosity, even when fluent. The material will tend to move into an area of lesser resistance created by the relief areas, rather than turning the corner, such as at outer extent  102 , and moving onto the transition section  100 . 
     In accordance with the methods of this application, the outer lug  90  is inserted into the outer slot, and an inner lug is inserted into an inner slot. The vane may be tack welded to the shrouds. The flowable attachment material is then deposited between the slots and the lugs, and the assembly is heated to allow the flowable attachment material to move to a final position at which it hardens, and to create the vane assembly  150 . 
     While the disclosure of this application has been directed to the outer shroud, a worker of ordinary skill in the art would recognize that all of these teachings would apply equally to an inner shroud, and may be utilized at both the inner and outer shrouds. 
     Although an example embodiment has been disclosed, a person of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For instance, a relief area may be created within the transition section. For this reason, the following claims should be studied to determine their true scope and content.

Technology Classification (CPC): 8