Patent Publication Number: US-11047397-B2

Title: Gas turbine engine stator vane mistuning

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to U.S. Provisional Application No. 61/931,414, which was filed on Jan. 24, 2014 and is incorporated herein by reference. 
    
    
     BACKGROUND 
     This disclosure relates to a stator vane array for a gas turbine engine. 
     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 combustor 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 typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
     The compressor and turbine section includes circumferential arrangements of fixed and rotating stages. Structural vibratory coupling between adjacent airfoils can occur during engine operation. For rotating stages of the engine, blade mistuning has been used in which there are two sets of blades are arranged in circumferentially alternating relationship to provide an even numbered blade array. One set of blades has a different characteristic than the other set of blades to provide two different resonant frequencies. For fixed stages, vanes have been mistuned by providing different sets of vanes in adjacent quadrants of the array. 
     SUMMARY 
     In one exemplary embodiment, a gas turbine engine includes a circumferential array of stator vanes that have first and second vanes with different vibrational frequencies than one another. The first vanes are arranged in circumferentially alternating relationship with the second vanes. 
     In a further embodiment of the above, a compressor section is included in which the array is arranged in the compressor section. 
     In a further embodiment of any of the above, the compressor section includes low and high pressure compressors. The array is arranged in the high pressure compressor. 
     In a further embodiment of any of the above, the stator vanes are provided on multiple arcuate segments. Each segment includes the first and second vanes. 
     In a further embodiment of any of the above, the segments are approximately 180°. 
     In a further embodiment of any of the above, one segment includes N blades, and the other segment includes N+3 blades. 
     In a further embodiment of any of the above, the array includes an odd number of vanes. A pair of first vanes is arranged next to one another. 
     In a further embodiment of any of the above, at least one of the first and second vanes are cantilevered. 
     In a further embodiment of any of the above, the first and second vanes are integrated with an outer platform. 
     In a further embodiment of any of the above, the outer platform is supported by hooks relative to an outer case structure. 
     In a further embodiment of any of the above, the second vanes have an increased bending moment frequency of about 60% relative to the first vanes. 
     In a further embodiment of any of the above, the second vanes have an increased stiffness. 
     In a further embodiment of any of the above, the second vanes have a larger airfoil thickness compared to a first vane airfoil thickness. 
     In a further embodiment of any of the above, the first and second vanes have the same leading and trailing edge thicknesses. 
     In a further embodiment of any of the above, the first and second vanes have the same chord, stagger angle and span. 
     In a further embodiment of any of the above, the first vanes have a reduced mass at an inner diameter of the first vanes. 
     In a further embodiment of any of the above, at least one of a leading edge and a trailing edge is clipped at the inner diameter. 
     In a further embodiment of any of the above, a stiffness of the second vanes is increased and a mass of the first vanes is decreased. 
     In a further embodiment of any of the above, the first and second vanes are provided on a fully integrated ring. 
     In another exemplary embodiment, a stator vane stage includes a circumferential array of stator vanes that have first and second vanes. The first vanes are arranged in circumferentially alternating relationship with the second vanes. The second vanes have an increased bending moment frequency of about 6% relative to the first vanes. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
         FIG. 1  schematically illustrates a gas turbine engine embodiment. 
         FIG. 2  is a schematic view through an engine section including a fixed stage and a rotating stage. 
         FIG. 3  is a perspective view of a segment of stator vanes having first and second vanes arranged in circumferentially alternating relationships with one another. 
         FIG. 4  is a schematic view of a full ring of stator vanes. 
         FIG. 5A  illustrates a segment having N blades and another segment having N+3 blades. 
         FIG. 5  is a schematic view of first and second segments of stator vanes. 
         FIG. 6  is a schematic side elevational view of an example stator vane. 
         FIG. 7  is a cross-sectional view of the stator vane shown in  FIG. 6  taken along line  7 - 7 . 
         FIG. 8  is a cross-sectional view of the stator vane shown in  FIG. 6  taken along line  8 - 8 . 
         FIG. 9  is an enlarged view of an inner diameter portion of the stator vane shown in  FIG. 6 . 
     
    
    
     The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible. 
     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 augmenter 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 X 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 X 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 invention 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, 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. 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. 
     Referring to  FIG. 2 , a portion of an engine section is shown, for example, a compressor section. It should be understood, however, that disclosed section also may be provided in a turbine section. 
     The section includes a fixed stage  60  that provides a circumferential array of vanes  64  arranged axially adjacent to a rotating stage  62 . In the example, the vane  64  includes an outer diameter portion  68  having hooks  66  that support the array of vanes  64  with respect to a case structure  74 . An airfoil  70  extends radially from a platform of the outer diameter portion  68 . In the examples that are illustrated, the vanes  64  are of the cantilevered type in which an inner diameter portion  72  of the airfoil  70  is unsupported. It should be understood that the disclosed vane arrangement could be used for vane structures having a platform at the inner diameter portion of the airfoil. 
     The vanes  64  may be arranged in a cluster to provide an arcuate vane segment  76 , as shown in  FIG. 3 . Circumferential ends  78  of adjacent segments  76  are sealed relative to one another. The vane segment  76  includes first vanes A and second vanes B that are arranged in an alternating relationship with the first vanes A. The vanes  64  have chord, stager angle and span parameters. The first and second vanes A, B have different vibration frequencies than one another to mistune the array of vanes and reduce the structural and aerodynamic coupling between adjacent vanes. As a result, the airfoil resonant vibration response, the vibration response after engine compressor stall from aerodynamic separation induced vibration, and the airfoil aero-elastic flutter vibration response all may be reduced. 
     The array of stator vanes may be provided as a full ring  80 , as shown in  FIG. 4 , or ring halves  82 ,  84  as shown in  FIG. 5 . An odd number of vanes may be provided in a segment, such that the same vane of adjacent segments would be arranged next to each other, as shown in  FIG. 5  (see adjacent vanes A-A). In one example, the first half  82  includes N number of blades, and the second half  84  includes N+3 blades. 
     Referring to  FIG. 6 , the vane  64  includes leading and trailing edges  86 ,  88 . The second vanes B may include an increased bending moment frequency of about 6% relative to the first vanes A. In another embodiment, the second vanes B may include an increased bending moment frequency of up to 60% relative to the first vanes A, and in another example about 6% relative to the first vanes. The second vanes B may also have an increased torsional frequency relative to the vanes A. This may be achieved in a variety of suitable manners. For example, the first and second vanes A, B may have different microstructures than one another. In another example, the second vanes may have an airfoil  170  with an increased stiffness, for example, by providing a larger airfoil thickness (dashed lines in  FIG. 7 ) compared to an airfoil thickness of the first vanes (solid line in  FIG. 7 ). However, it is desirable to maintain the same leading and trailing edge thicknesses, chord, stager angle and span of the first and second vanes, A, B. Another way of increasing the relative bending moment frequency of the second vanes is to reduce the mass of airfoil  270  of the first vane A, for example, by reducing the thickness and mass at the inner diameter portion  72  (dashed lines relative to solid lines in  FIG. 8 ). This also may be achieved by clipping off one or more corners  90  at least of one of the leading and trailing edges  86 ,  88  at the inner diameter portion  72 , as shown in  FIG. 9 . A combination of the stiffening or reducing mass may be used to change the relative bending moments of the first and second vanes A, B. 
     It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention. 
     Although the different examples have specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
     Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.