Patent Publication Number: US-10774650-B2

Title: Gas turbine engine airfoil

Description:
BACKGROUND 
     This disclosure relates to a gas turbine engine airfoil. More particularly, this disclosure relates to a gas turbine engine airfoil having an improved design. 
     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 shape of an airfoil designed for turbomachinery applications is an important characteristic. It is often a result of multidisciplinary considerations including aerodynamics, durability, structure and design. However, recent advances in the design of aerodynamically high-performing, high-pressure turbine blades, particularly at the tip, have caused increased difficulties in the design of blades. 
     SUMMARY 
     In one exemplary embodiment, a component for a gas turbine engine includes a platform that has a radially inner side and a radially outer side. A root portion extends from the radially inner side of the platform. An airfoil extends from the radially outer side of the platform. The airfoil includes a pressure side that extends between a leading edge and a trailing edge. A suction side extends between the leading edge and the trailing edge. A curvature inflection point is located between 30% and 70% of an axial chord length of the airfoil. 
     In a further embodiment of any of the above, the curvature inflection point at 0% span occurs at between 60% and 70% of the axial chord length. 
     In a further embodiment of any of the above, the curvature inflection point at 25% span occurs at between 50% and 60% of the axial chord length. 
     In a further embodiment of any of the above, the curvature inflection point at 50% span occurs at between 40% and 50% of the axial chord length. 
     In a further embodiment of any of the above, the curvature inflection point at 75% span occurs at between 30% and 40% of the axial chord length. 
     In a further embodiment of any of the above, the curvature inflection point at 90% span occurs at between 30 and 40% of the axial chord length. 
     In a further embodiment of any of the above, the curvature inflection point occurs at a span location with the smallest radius of curvature. 
     In a further embodiment of any of the above, the curvature inflection point at 0% span is axially downstream of the curvature inflection point at 50% span. 
     In a further embodiment of any of the above, the curvature inflection point at 0% span is axially downstream of the curvature inflection point at 50% span. 
     In a further embodiment of any of the above, the airfoil is a blade and extends from the platform to an unshrouded end and the curvature inflection point occurs at a span location with the smallest radius of curvature. 
     In another exemplary embodiment, a gas turbine engine includes a compressor section and a turbine section. A circumferential array of airfoils are located in one of the compressor section and the turbine section. At least one of the airfoils include a pressure side that extends between a leading edge and a trailing edge. A suction side extends between the leading edge and the trailing edge and a curvature inflection point is located between 30% and 70% of an axial chord length of the airfoil. 
     In a further embodiment of any of the above, the curvature inflection point at 0% span occurs at between 60% and 70% of the axial chord length. 
     In a further embodiment of any of the above, the curvature inflection point at 25% span occurs at between 50% and 60% of the axial chord length. 
     In a further embodiment of any of the above, the curvature inflection point at 50% span occurs at between 40% and 50% of the axial chord length. 
     In a further embodiment of any of the above, the curvature inflection point at 75% span occurs at between 30% and 40% of the axial chord length. 
     In a further embodiment of any of the above, the curvature inflection point at 90% span occurs at between 30% and 40% of the axial chord length. 
     In a further embodiment of any of the above, the curvature inflection point occurs at a span location with the smallest radius of curvature. 
     In a further embodiment of any of the above, the curvature inflection point at 0% span is axially downstream of the curvature inflection point at 50% span. 
     In a further embodiment of any of the above, the curvature inflection point at 0% span is axially downstream of the curvature inflection point at 50% span. 
     In a further embodiment of any of the above, the airfoil is a blade and extends from a platform to an unshrouded end and the curvature inflection point occurs at a span location with the smallest radius of curvature. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic view of an example gas turbine engine according to a first non-limiting embodiment. 
         FIG. 2  is a schematic view of a section of the gas turbine engine of  FIG. 1 , such as a turbine section. 
         FIG. 3  is a schematic view of a turbine blade. 
         FIG. 4  is a cross-sectional view through an airfoil according to this disclosure. 
         FIG. 5  is a graphical representation of a curvature inflection point on the airfoil at 0% span. 
         FIG. 6  is a graphical representation of a curvature inflection point on the airfoil at 25% span. 
         FIG. 7  is a graphical representation of a curvature inflection point on the airfoil at 50% span. 
         FIG. 8  is a graphical representation of a curvature inflection point on the airfoil at 75% span. 
         FIG. 9  is a graphical representation of a curvature inflection point on the airfoil at 90% span. 
     
    
    
     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 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 , and also 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 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 (10,668 meters). 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. 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). 
     Referring to  FIG. 2 , a cross-sectional view through a high pressure turbine section  54  is illustrated. In the example high pressure turbine section  54 , first and second arrays of circumferentially spaced fixed vanes  60 ,  62  are axially spaced apart from one another. A first stage array of circumferentially spaced turbine blades  64 , mounted to a rotor disk  68 , is arranged axially between the first and second fixed vane arrays. A second stage array of circumferentially spaced turbine blades  66  is arranged aft of the second array of fixed vanes  62 . It should be understood that any number of stages may be used. Moreover, the disclosed airfoil may be used in a compressor section, turbine section and/or fixed or rotating stages. 
     The turbine blades each include a tip  80  adjacent to a blade outer air seal  70  of a case structure  72 , which provides an outer flow path. In the illustrated embodiment, the tip  80  is unshrouded. The first and second stage arrays of turbine vanes and first and second stage arrays of turbine blades are arranged within a core flow path C and are operatively connected to a spool  32 , for example. 
     Each blade  64  includes an inner platform  76  respectively defining inner flow path. The platform inner platform  76  supports an airfoil  78  that extends in a radial direction R, as shown in  FIG. 3 . It should be understood that the turbine vanes  60 ,  62  may be discrete from one another or arranged in integrated clusters. The airfoil  78  includes a leading edge  82  and a trailing edge  84 . 
     The airfoil  78  is provided between pressure side  94  (predominantly concave) and suction side (predominantly convex)  96  in an airfoil thickness direction ( FIG. 4 ), which is generally perpendicular to a chord-wise direction provided between the leading and trailing edges  82 ,  84 . Multiple turbine blades  64  are arranged in a circumferentially spaced apart manner in a circumferential direction Y ( FIG. 4 ). The airfoil  78  includes multiple film cooling holes  90 ,  92  respectively schematically illustrated on the leading edge  82  and the pressure side  94  ( FIG. 4 ). 
     The turbine blades  64  are constructed from a high strength, heat resistant material such as a nickel-based or cobalt-based superalloy, or of a high temperature, stress resistant ceramic or composite material. In cooled configurations, internal fluid passages and external cooling apertures provide for a combination of impingement and film cooling. Other cooling approaches may be used such as trip strips, pedestals or other convective cooling techniques. In addition, one or more thermal barrier coatings, abrasion-resistant coatings or other protective coatings may be applied to the turbine vanes  62 . 
       FIGS. 3 and 4  schematically illustrates an airfoil including pressure and suction sides joined at leading and trailing edges  82 ,  84 . A root  74  supports the platform  76 . The root  74  may include a fir tree that is received in a correspondingly shaped slot in the rotor  68 , as is known. The airfoil extends a span from a support, such as an inner platform  76  to an end, such as the tip  80  in a radial direction R. The 0% span and the 100% span positions, respectively, correspond to the radial airfoil positions at the support and the end. The leading and trailing edges  82 ,  84  are spaced apart from one another and an axial chord b x  length ( FIG. 4 ) extends in the axial direction X. 
       FIGS. 5-9  graphically represent changes in curvature at different axial positions of the airfoil  78  along the axial chord b x . The relative position along the axial chord length b x  is represented by X/b x  such that the graphical representations can apply to airfoils of varying axial chord length b x . 
     A cross-section of the airfoil  78 , as illustrated in  FIG. 4 , at a particular span location includes a curvature inflection point  100  that corresponds to the location of the smallest radius of curvature r and the greatest amount of curvature. In general, the curvature of the pressure side  94  increases moving downstream along the pressure side  94  in the axial chord length bx until reaching the curvature inflection point  100 . Downstream of the inflection point, the amount of curvature begins to decrease. In the illustrated embodiment, the curvature inflection point  100  occurs between 30% and 70% of the axial chord bx. 
       FIG. 5  illustrates a non-limiting embodiment of a line graph representing the change in curvature at 0% span for varying percentages of the axial chord length b x  of the airfoil for three different examples. As shown in  FIG. 5 , the curvature inflection point  100  occurs at or approximately below 70% of the axial chord length b x . 
       FIG. 6  illustrates a non-limiting embodiment of a line graph representing the change in curvature at 25% span for varying percentages of the axial chord length b x  of the airfoil for three different examples. As shown in  FIG. 6 , the curvature inflection point  100  occurs between 50% and 60% of the axial chord length b x  or at approximately 55% of the axial chord length b x . 
       FIG. 7  illustrates a non-limiting embodiment of a line graph representing the change in curvature at 50% span for varying percentages of the axial chord length b x  of the airfoil for three different examples. As shown in  FIG. 7 , the curvature inflection point  100  occurs between 40% and 50% of the axial chord length b x  or at approximately 45% of the axial chord length b x . 
       FIG. 8  illustrates a non-limiting embodiment of a line graph representing the change in curvature at 75% span for varying percentages of the axial chord length b x  of the airfoil for three different examples. As shown in  FIG. 8 , the curvature inflection point  100  occurs between 30% and 40% of the axial chord length b x  or at approximately 35% of the axial chord length b x . 
       FIG. 9  illustrates a non-limiting embodiment of a line graph representing the change in curvature at 90% span for varying percentages of the axial chord length b x  of the airfoil for three different examples. As shown in  FIG. 9 , the curvature inflection point  100  occurs between 30% and 40% of the axial chord length b x  or at approximately 35% of the axial chord length b x . 
     As shown in  FIGS. 5-9 , the curvature inflection point  100  moves axially forward when moving from 0% span radially outward. By locating the curvature inflection point  100  further aft along the axial chord length, a region immediately downstream of the leading edge  82  on the pressure side  94  includes non-stagnated flow. By reducing the amount of stagnated flow on the pressure side  94 , more efficient ways of cooling the airfoil  78  can be utilized to reduce the overall cooling load required by the airfoil  78  and improve overall gas turbine engine efficiency. 
     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.