Patent Publication Number: US-11391294-B2

Title: Gas turbine engine airfoil

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of U.S. application Ser. No. 15/113,870 filed Jul. 25, 2016, which is a U.S. National Phase of International Application No. PCT/US2015/016032, filed Feb. 16, 2015, which claims priority to U.S. Provisional Application No. 61/941,706, which was filed on Feb. 19, 2014 and is incorporated herein by reference. 
    
    
     BACKGROUND 
     This disclosure relates to gas turbine engine airfoils. More particularly, the disclosure relates to airfoil tangential stacking offset in, for example, a gas turbine engine compressor. 
     A turbine engine such as 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 at least low and high pressure compressors, and the turbine section includes at least low and high pressure turbines. 
     Direct drive gas turbine engines include a fan section that is driven directly by one of the turbine shafts. Rotor blades in the fan section and a low pressure compressor of the compressor section of direct drive engines rotate in the same direction. 
     Gas turbine engines have been proposed in which a geared architecture is arranged between the fan section and at least some turbines in the turbine section. The geared architecture enables the associated compressor of the compressor section to be driven at much higher rotational speeds, improving overall efficiency of the engine. The propulsive efficiency of a gas turbine engine depends on many different factors, such as the design of the engine and the resulting performance debits on the fan that propels the engine and the compressor section downstream from the fan. Physical interaction between the fan and the air causes downstream turbulence and further losses. Although some basic principles behind such losses are understood, identifying and changing appropriate design factors to reduce such losses for a given engine architecture has proven to be a complex and elusive task. 
     Prior compressor airfoil geometries may not be suitable for the compressor section of gas turbine engines using a geared architecture, since the significantly different speeds of the compressor changes the desired aerodynamics of the airfoils within the compressor section. Counter-rotating fan and compressor blades, which may be used in geared architecture engines, also present design challenges. 
     SUMMARY 
     In one exemplary embodiment, a compressor airfoil of a turbine engine having a geared architecture includes pressure and suction sides that extend in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between a tangential stacking offset and span position that defines a curve that is non-linear. 
     In a further embodiment of the above, the curve has a negative slope from 90% span to 100% span. The negative slope leans toward the pressure side. 
     In a further embodiment of any of the above, the curve has a negative slope from 80% span to 100% span. 
     In a further embodiment of any of the above, the curve has a positive slope beginning at 0% span. The positive slope leans toward the suction side. 
     In a further embodiment of any of the above, the positive slope extends from 0% span to 40% span. 
     In a further embodiment of any of the above, the curve transitions from the positive slope to the negative slope in a range of 40% span to 75% span. 
     In another exemplary embodiment, a gas turbine engine includes a combustor section that is arranged between a compressor section and a turbine section. A fan section has an array of twenty-six or fewer fan blades. The fan section has a fan pressure ratio of less than 1.55. A geared architecture couples the fan section to the turbine section or the compressor section. An airfoil is arranged in the compressor section and includes pressure and suction sides that extend in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between a tangential stacking offset and span position that defines a curve that is non-linear. 
     In a further embodiment of any of the above, the compressor section includes at least a low pressure compressor and a high pressure compressor. The high pressure compressor is arranged immediately upstream of the combustor section. 
     In a further embodiment of any of the above, the airfoil is provided in a compressor outside the high pressure compressor. 
     In a further embodiment of any of the above, the low pressure compressor is counter-rotating relative to the fan blades. 
     In a further embodiment of any of the above, the gas turbine engine is a two-spool configuration. 
     In a further embodiment of any of the above, the low pressure compressor is immediately downstream from the fan section. 
     In a further embodiment of any of the above, the airfoil is rotatable relative to an engine static structure. 
     In a further embodiment of any of the above, the curve has a negative slope from 90% span to 100% span. 
     In a further embodiment of any of the above, the curve has a negative slope from 80% span to 100% span. 
     In a further embodiment of any of the above, the curve has a positive slope beginning at 0% span. The positive slope leans toward the suction side. 
     In a further embodiment of any of the above, the positive slope extends from 0% span to 40% span. 
     In a further embodiment of any of the above, the curve transitions from the positive slope to the negative slope in a range of 40% span to 75% span. 
    
    
     
       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 with a geared architecture. 
         FIG. 2  schematically illustrates a low pressure compressor section of the gas turbine engine of  FIG. 1 . 
         FIG. 3  is a schematic view of airfoil span positions. 
         FIG. 4  is a schematic view of a cross-section of an airfoil sectioned at a particular span position and depicting directional indicators. 
         FIG. 5  graphically depicts curves for several example airfoil tangential stacking offset relative to span, including two prior art curves and several inventive curves according to this disclosure. 
     
    
    
     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. That is, the disclosed airfoils may be used for engine configurations such as, for example, direct fan drives, or two- or three-spool engines with a speed change mechanism coupling the fan with a compressor or a turbine sections. 
     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 epicyclic 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. 
     The example gas turbine engine includes the fan  42  that comprises in one non-limiting embodiment less than about twenty-six (26) fan blades. In another non-limiting embodiment, the fan section  22  includes less than about twenty (20) fan blades. Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about six (6) turbine rotors schematically indicated at  34 . In another non-limiting example embodiment the low pressure turbine  46  includes about three (3) turbine rotors. A ratio between the number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotors  34  in the low pressure turbine  46  and the number of blades  42  in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
     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 1.55. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. In another non-limiting embodiment the low fan pressure ratio is from 1.1 to 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 1200 ft/second (365.7 meters/second). 
     Referring to  FIG. 2 , which schematically illustrates an example low pressure compressor (LPC)  44 , a variable inlet guide vane (IGV) is arranged downstream from a fan exit stator (FES). The figure is highly schematic, and the geometry and orientation of various features may be other than shown. An actuator driven by a controller actuates the IGV about their respective axes. Multiple airfoils are arranged downstream from the IGV. The airfoils include alternating stages of rotors (ROTOR1, ROTOR2, ROTOR3, ROTOR4) and stators (STATOR1, STATOR2, STATOR3, STATOR4). In the example shown in  FIG. 2 , the LPC includes four rotors alternating with four stators. However, in another example, a different number of rotors and a different number of stators may be used. Moreover, the IGV and stator stages may all be variable, fixed or a combination thereof. 
     The disclosed airfoils may be used in a low pressure compressor of a two spool engine or in portions of other compressor configurations, such as low, intermediate and/or high pressure areas of a three spool engine. However, it should be understood that any of the disclosed airfoils may be used for blades or vanes, and in any of the compressor section, turbine section and fan section. The low pressure compressor may be counter-rotating relative to the fan blades. 
     Referring to  FIG. 3 , span positions on an airfoil  64  are schematically illustrated from 0% to 100% in 10% increments. Each section at a given span position is provided by a conical cut that corresponds to the shape of the core flow path, as shown by the large dashed lines. In the case of an airfoil with an integral platform, the 0% span position corresponds to the radially innermost location where the airfoil meets the fillet joining the airfoil to the inner platform. In the case of an airfoil without an integral platform, the 0% span position corresponds to the radially innermost location where the discrete platform meets the exterior surface of the airfoil. For airfoils having no outer platform, such as blades, the 100% span position corresponds to the tip  66 . For airfoils having no platform at the inner diameter, such as cantilevered stators, the 0% span position corresponds to the inner diameter location of the airfoil. For stators, the 100% span position corresponds to the outermost location where the airfoil meets the fillet joining the airfoil to the outer platform. 
     Airfoils in each stage of the LPC are specifically designed radially from an inner airfoil location (0% span) to an outer airfoil location (100% span) and along circumferentially opposite pressure and suction sides  72 ,  74  extending in chord between a leading and trailing edges  68 ,  70  (see  FIG. 4 ). Each airfoil is specifically twisted with a corresponding stagger angle and bent with specific sweep and/or dihedral angles along the airfoil. Airfoil geometric shapes, stacking offsets, chord profiles, stagger angles, sweep and dihedral angles, among other associated features, are incorporated individually or collectively to improve characteristics such as aerodynamic efficiency, structural integrity, and vibration mitigation, for example, in a gas turbine engine with a geared architecture in view of the higher LPC rotational speeds. 
     The airfoil  64  has an exterior surface  76  providing a contour that extends from a leading edge  68  generally aftward in a chord-wise direction H to a trailing edge  70 , as shown in  FIG. 4 . Pressure and suction sides  72 ,  74  join one another at the leading and trailing edges  68 ,  70  and are spaced apart from one another in an airfoil thickness direction T. An array of airfoils  64  are positioned about the axis X (corresponding to an X direction) in a circumferential or tangential direction Y. Any suitable number of airfoils may be used for a particular stage in a given engine application. 
     A tangential stacking offset Y d  corresponds to the location of the center of gravity Y CG  for a particular section at a given span location relative to a reference point  80  in the Y direction. The reference point  80  is a location such as the tangential center of the root, for example. The value Y d  corresponds to the tangential distance from the reference point  80  to the center of gravity. A positive Y value corresponds to the opposite rotational direction as the hub&#39;s rotation, or toward the suction side of the airfoil. A negative Y d  value corresponds to the same rotational direction as the hub&#39;s rotation, or toward the pressure side of the airfoil. A positive slope to the tangential stacking offset profile leans toward the suction side  74 , and a negative slope to the tangential stacking offset profile leans toward the pressure side  72 . 
     The exterior surface  76  of the airfoil  64  generates lift based upon its geometry and directs flow along the core flow path C. The airfoil  64  may be constructed from a composite material, or an aluminum alloy or titanium alloy, or a combination of one or more of these. Abrasion-resistant coatings or other protective coatings may be applied to the airfoil. The rotor stages may constructed as an integrally bladed rotor, if desired, or discrete blades having roots secured within corresponding rotor slots of a disc. The stators may be provided by individual vanes, clusters of vanes, or a full ring of vanes. 
     Airfoil geometries can be described with respect to various parameters provided. The disclosed graph(s) illustrate the relationships between the referenced parameters within 10% of the desired values, which correspond to a hot aerodynamic design point for the airfoil. In another example, the referenced parameters are within 5% of the desired values, and in another example, the reference parameters are within 2% of the desired values. It should be understood that the airfoils may be oriented differently than depicted, depending on the rotational direction of the blades. The signs (positive or negative) used, if any, in the graphs of this disclosure are controlling and the drawings should then be understood as a schematic representation of one example airfoil if inconsistent with the graphs. The signs in this disclosure, including any graphs, comply with the “right hand rule.” The tangential stacking offset varies with position along the span, and varies between a hot, running condition and a cold, static (“on the bench”) condition. 
     The geared architecture  48  of the disclosed example permits the fan  42  to be driven by the low pressure turbine  46  through the low speed spool  30  at a lower angular speed than the low pressure turbine  46 , which enables the LPC  44  to rotate at higher, more useful speeds. The tangential stacking offset in a hot, running condition along the span of the airfoils  64  provides necessary compressor operation in cruise at the higher speeds enabled by the geared architecture  48 , to thereby enhance aerodynamic functionality and thermal efficiency. As used herein, the hot, running condition is the condition during cruise of the gas turbine engine  20 . For example, the tangential stacking offset in the hot, running condition can be determined in a known manner using finite element analysis. 
       FIG. 5  illustrates the relationship between the tangential stacking offset and the average span (AVERAGE SPAN %), which is the average of the radial position at the leading and trailing edges  68 ,  70 . The tangential stacking offset assumes a center of gravity based upon a homogeneous material throughout the airfoil cross-section. In one example, the airfoils are LPC rotor blades. Two prior art curves (“PRIOR ART”) are illustrated as well as several example inventive curves  88 ,  90 ,  92 ,  94 ,  96 . The airfoil  64  has a relationship between an tangential stacking offset and span position that defines a curve that is non-linear, unlike the prior art curves. 
     The curves have a negative slope from 90% span to 100% span, wherein the negative slope leans toward the pressure side. In one example, the curve has a negative slope from 80% span to 100% span. 
     The curve has a positive slope beginning at 0% span, where the positive slope leans toward the suction side. In one example, the positive slope extends from 0% span to 40% span. The curve transitions from the positive slope to the negative slope in a range of 40% span to 75% span, for example. 
     The prior art has generally used straight, tangentially linearly stacked LPC blades that lean gradually toward the suction side. The disclosed airfoils include significant pressure side-leaning to improve the aerodynamic efficiency of the high speed LPC blades downstream from a counter-rotating fan. 
     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.