Patent Publication Number: US-10329931-B2

Title: Stator assembly for a gas turbine engine

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to U.S. Provisional Application No. 62/058,389, which was filed on Oct. 1, 2014 and is incorporated herein by reference. 
    
    
     BACKGROUND 
     A gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. 
     The compressor section for the gas turbine engine generally includes a rotor assembly and a stator vane assembly. The rotor assembly includes rows or arrays of rotor blades. The arrays of rotor blades extend radially outward across a gas path. The stator vane assembly includes arrays of stator vanes axially separating each of the arrays of rotor blades. The arrays of stator vanes extend inward from a radially outward case across the gas path into proximity with the rotor assembly. The arrays of stator vanes guide a working flow medium through the gas path as the working flow medium is discharged from each of the arrays of rotor blades. 
     A significant amount of effort is placed on increasing the efficiency of the gas turbine engine. One way to increase the efficiency of the gas turbine engine is to decrease the amount of compressor air that leaks from the compressor section. In order to reduce unwanted air leaks from the compressor section, various seals are incorporated into the compressor section to prevent the compressed air from leaking out. One type of seal used is a knife edge seal. Knife edge seals create a region with a pressure drop to deter compressed air from leaking past the seal. However, leakage occurs in other locations, such as between vanes. Therefore, there is a need for a compressor section with that reduces the loss of compressed air. 
     SUMMARY 
     In one exemplary embodiment, a stator assembly includes a platform located on a radially inner end of a plurality of vanes that connects a first vane to a second vane. There is a platform groove on a radially inner side of the platform between the first vane and the second vane. 
     In a further embodiment of the above, a radially outer side of the platform is continuous between the first vane and the second vane. 
     In a further embodiment of any of the above, a bridge portion extends along a distal end of the platform groove and includes a crack. 
     In a further embodiment of any of the above, the crack extends between a radially inner side of the bridge portion and a radially outer side of the bridge portion. 
     In a further embodiment of any of the above, the bridge portion extends along at least one of a leading edge and a trialing edge of the platform. 
     In a further embodiment of any of the above, the platform groove extends between approximately 5% and 20% of the thickness of the platform. 
     In a further embodiment of any of the above, the platform includes a leading edge and a trailing edge. The platform groove is spaced axially inward from the leading edge and the trailing edge. 
     In a further embodiment of any of the above, the platform groove includes a component that extends in an axial direction and a circumferential direction. 
     In a further embodiment of any of the above, a damper extends around the platform. 
     In another exemplary embodiment, a stator assembly for a gas turbine engine includes a platform that is located on a radially inner end of a plurality of vanes. A platform groove is on a radially inner side of the platform between a first vane and a second vane. A bridge portion extends along a distal end of the platform groove and includes a crack. 
     In a further embodiment of any of the above, the crack extends between a radially inner side of the bridge portion and a radially outer side of the bridge portion. 
     In a further embodiment of any of the above, the groove extends between approximately 5% and 20% of the thickness of the platform. 
     In a further embodiment of any of the above, the platform includes a leading edge and a trailing edge. The platform groove is spaced axially inward from the leading edge and the trailing edge. 
     In a further embodiment of any of the above, a damper extends around the platform. 
     In a further embodiment of any of the above, the bridge portion extends along a leading edge and a trialing edge of the platform. 
     In one exemplary embodiment, a method of forming a stator assembly includes forming a plurality of vanes with a platform located on a radially inner end, forming a platform groove between a first vane and a second vane and forming a bridge portion that extends along a distal end of the platform groove. 
     In a further embodiment of the above, the method includes cracking the bridge portion. 
     In a further embodiment of any of the above, the platform groove is located on a radially inner side of the platform. A radially outer side of the platform is continuous between the first vane and the second vane. 
     In a further embodiment of any of the above, the platform groove is formed by electro-discharge machining. 
     In a further embodiment of any of the above, the bridge portion extends along a leading edge and a trialing edge of the platform. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic view of an example gas turbine engine. 
         FIG. 2  is an enlarged schematic cross-section of a high pressure compressor section for the gas turbine engine of  FIG. 1 . 
         FIG. 3  is an enlarged view of a vane platform of  FIG. 2 . 
         FIG. 4  is a schematic view of a vane segment. 
         FIG. 5  is another enlarged view of the vane platform. 
         FIG. 6  is a cross-section taken along line  6 - 6  of  FIG. 5 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as 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 along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool 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 two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The exemplary 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, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as 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 second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  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  further supports 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. 
     The core airflow 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 C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     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 about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary 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 five. 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 five 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. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system  58 . The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second). 
       FIG. 2  illustrates an enlarged schematic view of the high pressure compressor  52 , however, other sections of the gas turbine engine  20  could benefit from this disclosure. The high pressure compressor  52  includes multiple stages, however, only a first rotor assembly  60  and a second rotor assembly  62  are shown in the illustrated example. The first rotor assembly  60  and the second rotor assembly  62  are attached to the outer shaft  50  ( FIG. 1 ). 
     The first rotor assembly  60  includes a first array of rotor blades  64  circumferentially spaced around a first disk  68  and the second rotor assembly  62  includes a second array of rotor blades  66  circumferentially spaced around a second disk  70 . Each of the first and second array of rotor blades  64 ,  66  include a respective first and second root portion  72 ,  74 , a first and second platform  76 ,  78 , and a first and a second airfoil  80 ,  82 . Each of the first and second root portions  72 ,  74  is received within a respective one of the first and second disks  68 ,  70 . The first airfoil  80  and the second airfoil  82  extend radially outward toward a first and second blade outer air seal (BOAS) assembly  84 ,  86 , respectively. 
     Alternatively, the first rotor assembly  60  or the second rotor assembly  62  could be an integrally bladed rotor assembly with the first and second airfoils  80 ,  82  formed integrally with the respective first and second disks  68 ,  70 , without a separate first and second root portion  72 ,  74  or a separate first and second platform  76 ,  78 , respectively. 
     A shroud assembly  88  within the engine case structure  36  between the first rotor assembly  60  and the second rotor assembly  62  directs the core airflow in the core flow path from the first array of rotor blades  64  to the second array of rotor blades  66 . The shroud assembly  88  may at least partially support the first and second blade outer air seals  84 ,  86  and include an array of vanes  90  that extend between a respective inner vane platform  92  and an outer vane platform  94 . The outer vane platform  94  may be supported by the engine case structure  36  and the inner vane platform  92  supports abradable annular seals  96 , such as a honeycomb, to seal the core airflow in the axial direction with respect to knife edges  98  on a seal assembly  100 . 
       FIG. 3  shows an enlarged view of the inner vane platform  92  along with a portion of the vane  90 . The inner vane platform  92  includes a pair of protrusions  102  that retain an inner diameter air seal carrier  104  that supports the abradable annular seals  96 . An inner diameter platform spring  106  radially loads the inner diameter air seal carrier  104  against the pair of protrusions  102  to control vibratory response of the vane  90  with frictional damping. In the illustrated example, the inner diameter platform spring  106  includes a mid-portion  108  that abuts the inner vane platform  92  and flexible ends  110  that bend over the mid-portion  108  and abut the inner diameter air seal carrier  104  to provide a biasing force that damps vibrations. In another example, the inner diameter platform spring  106  could include only a single flexible end  110 . 
       FIG. 4  illustrates a vane segment  112  with a plurality of the vanes  90  forming a portion of a stator ring  113  (shown in dashed lines). The outer vane platforms  94  of the vanes  90  are attached together circumferentially and form an outer diameter shroud  114 . The outer diameter shroud  114  extends continuously such that at least a portion of the outer vane platform  94  between adjacent vanes  90  is free of gaps. The inner vane platforms  92  of the vanes  90  are attached together circumferentially and form an inner diameter shroud  116 . The inner diameter shroud  116  extends continuously such that at least a portion of the inner vane platform  92  between adjacent vanes  90  is free of gaps. 
     As shown in  FIG. 5 , a groove  120  is formed in the inner vane platform  92 . The groove  120  extends through a substantial portion of the inner vane platform  92 . In the illustrated example, the groove  120  extends to a leading edge  122  and a trailing edge of the inner vane platform  92 . A bridge portion  126  extends along a radially outer portion of the inner vane platform  92  and onto the leading edge  122  and the trailing edge  124 . The bridge portion  126  includes an example non-limiting thickness D 1  of approximately 0.010 inches to 0.020 inches (0.254 mm to 0.508 mm). 
     The vanes  90  can be cast, fabricated, or machined as a single ring or segments of a ring as shown in  FIG. 4 . The groove  120  is formed in the inner vane platform  92  between adjacent vanes  90  through a machining process, such as electro-discharge machining (EDM) with a thin plate electrode in the shape of the groove  120 . Alternatively, the groove  120  could be formed without additional machining if the vanes  90  were produced with an additive manufacturing process. In the illustrated example, the groove  120  extends at least 50% through a thickness of the inner vane platform  92 . In another example, the groove  120  extends between approximately 5% and 20% of the thickness of the inner vane platform  92 . 
     As shown in  FIG. 6 , a crack  128  can form in the bridge portion  126  in the inner vane platform  92  adjacent a distal end of the groove  120 . A radius of the distal end of the groove  120  can function as a crack initiation site so that the crack  128  will form from the distal end and extend radially outward until the crack  128  reaches a radially outer diameter of the inner vane platform  92 . The crack  128  could be parallel to the engine axis “A” or skewed with some circumferential component relative to the engine axis “A.” 
     The cracks  128  are caused by static or vibratory loads that occur in the vanes  90  under typical operation of the gas turbine engine  20 . The thickness D 1  of the bridge  126  is designed so as not to be able to withstand these loads without forming the cracks  128 . 
     The crack  128  will allow for relative movement between adjacent vanes  90  while providing the smallest possible circumferential gap because opposing surfaces of the crack  128  form nearly perfect matching faces. Because the crack  128  will allow for the smallest possible circumferential gap in the inner vane platform  92 , less compressed air will leak past the inner vane platform  92  and increase the performance of the gas turbine engine  20 . 
     Additionally, by forming the groove  120  with an EDM having a draft angle along the leading edge  122  and trialing edge  124  that forms a sharp point at the radially inner end of the bridge portion  126 , crack propagation along the bridge portion  126  is promoted. 
     The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.