Patent Publication Number: US-2016238021-A1

Title: Compressor Airfoil

Description:
FIELD 
     The present disclosure relates generally to a gas turbine engine, and more specifically, to a compressor airfoil for gas turbine engines. 
     BACKGROUND 
     In the compressor of a gas turbine engine, rotor airfoils are surrounded by an abradable circumferential outer airfoil seal (OAS). The abradable material is worn away by abrasive on the tips of the rotor airfoils, during normal run-in operation at the beginning of engine life, and also during more aggressive rub incursions during unplanned events such as foreign object ingestion and compressor surges. Sufficiently aggressive radial and/or axial rubs will result in the leading edge (LE) and trailing edge (TE) portions of the airfoil tip contacting the abradable seal. Traditionally the geometry of the LE tip and TE tip are sharp corners, designed primarily for aerodynamic function. As a result, the contact stresses, such as Hertzian stresses, are very high during these severe rubs, which causes abrasive wear, scoring, pitting, and/or spalling of the airfoil and/or abrasive coating as well as significant local temperature spikes due to sliding friction. 
     SUMMARY 
     A compressor airfoil is disclosed. The compressor airfoil includes a leading edge having a leading edge tip profile with one of a leading edge tip radius and a leading edge tip chamfer and a trailing edge having a trailing edge tip profile with one of a trailing edge tip radius and a trailing edge tip chamfer. The compressor airfoil also includes a radially outboard edge including a surface connecting the leading edge and the trailing edge. 
     The compressor airfoil may have an abrasive coating disposed over at least a portion of the compressor airfoil. The abrasive coating is disposed over at least a portion of the leading edge tip profile and/or at least a portion of the trailing edge tip profile. The abrasive coating is disposed over at least a portion of the radially outboard edge. 
     The leading edge tip radius includes at least one of polygonal curvature, polyhedral curvature, complex curvature, conventional curvature and compound curvature. 
     The trailing edge tip radius includes at least one of polygonal curvature, polyhedral curvature, complex curvature, conventional curvature and compound curvature. 
     A compressor airfoil is disclosed including a leading edge including a leading edge tip profile, a trailing edge including a trailing edge tip profile, a radially outboard edge including a surface connecting the leading edge and the trailing edge, and an abrasive coating disposed over at least a portion of the radially outboard edge and at least a portion of at least one of the leading edge tip profile and the trailing edge tip profile. 
     Similarly, the leading edge tip radius includes at least one of: polygonal curvature, polyhedral curvature, complex curvature, conventional curvature and compound curvature. The abrasive coating can cover entirely the leading edge tip radius. 
     Also similarly, the trailing edge tip radius includes at least one of: polygonal curvature, polyhedral curvature, complex curvature, conventional curvature and compound curvature. The abrasive coating can cover entirely the trailing edge tip radius. 
     The leading edge tip profile may include a leading edge tip chamfer. The trailing edge tip profile may include a trailing edge tip chamfer. The abrasive coating may entirely cover the leading edge tip chamfer and/or trailing edge tip chamfer. 
     A method is disclosed. The method may include forming a compressor airfoil bounded by a leading edge, a trailing edge, and a radially outboard edge connecting the leading edge and the trailing edge, forming a leading edge tip profile on the leading edge, and forming a trailing edge tip profile on the trailing edge. The leading edge tip profile includes a leading edge tip radius or a leading edge tip chamfer and the trailing edge tip profile includes a trailing edge tip radius or a trailing edge tip chamfer. The method may also include depositing an abrasive coating on the radially outboard edge and at least a portion of the trailing edge tip profile and/or may also include depositing an abrasive coating on the radially outboard edge and at least a portion of the leading edge tip profile. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, is best be obtained by referring to the detailed description and claims when considered in connection with the drawing figures, wherein like numerals denote like elements. 
         FIG. 1  illustrates cross-sectional view of an exemplary gas turbine engine; 
         FIG. 2  illustrates an example compressor airfoil having a radiused leading edge tip profile and a radiused trailing edge tip profile; 
         FIG. 3  illustrates an example compressor airfoil having a chamfered leading edge tip profile and a chamfered trailing edge tip profile; and 
         FIG. 4  depicts a method of forming a compressor airfoil. 
     
    
    
     DETAILED DESCRIPTION 
     With reference to  FIG. 1 , an exemplary gas turbine engine  2  is provided. Gas turbine engine  2  is a two-spool turbofan that generally incorporates a fan section  4 , a compressor section  6 , a combustor section  8  and a turbine section  10 . Alternative engines include, for example, an augmentor section among other systems or features. In operation, fan section  4  drives air along a bypass flow-path B while compressor section  6  drives air along a core flow-path C for compression and communication into combustor section  8  then expansion through turbine section  10 . Although depicted as a turbofan gas turbine engine  2  herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings is applicable to other types of turbine engines including three-spool architectures. 
     Gas turbine engine  2  generally comprises a low speed spool  12  and a high speed spool  14  mounted for rotation about an engine central longitudinal axis X-X′ relative to an engine static structure  16  via several bearing systems  18 - 1 ,  18 - 2 , and  18 - 3 . It should be understood that bearing systems is alternatively or additionally provided at locations, including for example, bearing system  18 - 1 , bearing system  18 - 2 , and bearing system  18 - 3 . 
     Low speed spool  12  generally comprises an inner shaft  20  that interconnects a fan  22 , a low pressure compressor section  24 , e.g., a first compressor section, and a low pressure turbine section  26 , e.g., a first turbine section. Inner shaft  20  is connected to fan  22  through a geared architecture  28  that drives the fan  22  at a lower speed than low speed spool  12 . Geared architecture  28  comprises a gear assembly  42  enclosed within a gear housing  44 . Gear assembly  42  couples the inner shaft  20  to a rotating fan structure. High speed spool  14  comprises an outer shaft  30  that interconnects a high pressure compressor section  32 , e.g., second compressor section, and high pressure turbine section  34 , e.g., second turbine section. A combustor  36  is located between high pressure compressor section  32  and high pressure turbine section  34 . A mid-turbine frame  38  of engine static structure  16  is located generally between high pressure turbine section  34  and low pressure turbine section  26 . Mid-turbine frame  38  supports one or more bearing systems  18 , such as  18 - 3 , in turbine section  10 . Inner shaft  20  and outer shaft  30  is concentric and rotates via bearing systems  18  about the engine central longitudinal axis X-X′, 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 low pressure compressor section  24  then high pressure compressor section  32 , mixed and burned with fuel in combustor  36 , then expanded over high pressure turbine section  34  and low pressure turbine section  26 . Mid-turbine frame  38  includes airfoils  40 , which are in the core airflow path. Turbines  26 ,  34  rotationally drive the respective low speed spool  12  and high speed spool  14  in response to the expansion. 
     Gas turbine engine  2  is, for example, a high-bypass geared aircraft engine. The bypass ratio of gas turbine engine  2  is optionally greater than about six (6). The bypass ratio of gas turbine engine  2  is optionally greater than ten (10). Geared architecture  28  is an epicyclic gear train, such as a star gear system, e.g., sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear, or other gear system. Geared architecture  28  has a gear reduction ratio of greater than about 2.3 and low pressure turbine section  26  has a pressure ratio that is greater than about 5. The bypass ratio of gas turbine engine  2  is greater than about ten (10:1). The diameter of fan  22  is significantly larger than that of the low pressure compressor section  24 , and the low pressure turbine section  26  has a pressure ratio that is greater than about 5:1. Low pressure turbine section  26  pressure ratio is measured prior to inlet of low pressure turbine section  26  as related to the pressure at the outlet of low pressure turbine section  26  prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of a suitable geared architecture engine and that the present disclosure contemplates other turbine engines including direct drive turbofans. 
     The next generation of turbofan engines is designed for higher efficiency, which is associated with higher pressure ratios and higher temperatures in the high speed spool  14  and other engine sections. These higher operating temperatures and pressure ratios create operating environments that cause thermal loads that are higher than thermal loads conventionally encountered. Operating conditions in high pressure compressor section  32  are often approximately 1400° F. (approximately 760° C.) or more, and operating conditions in combustor  36  are often higher. Moreover, during transient events and incursions, such as during periods of acceleration, or during events such as bird strike events, temperatures and pressures rapidly vary, such shocks varying Hertzian stresses and rub contact of compressor airfoils. 
     The present disclosure contemplates considering Hertzian stresses during rub contact when producing the leading edge (LE) and trailing edge (TE) portions of an airfoil. Per Hertzian theory, the maximum contact pressure is in part a function of the radius of both the curved bodies that are in contact in compression. All other parameters being equal, the larger the radii in contact, the smaller the stresses. During radial and/or axial rub events the LE and TE tips cut into an abradable seal, imparting their own small-radius geometry in the seal at the points of LE tip and TE tip contact and resulting in even higher stresses during subsequent rubs. Compressor airfoil geometries, such as that disclosed herein and such as illustrated in the Figures, address Hertzian stresses during rub in gas turbine engine airfoil design resulting in reducing these contact stresses. The resulting design includes a rounded LE tip and TE tip, wherein specific geometries are unique to each part based on independent analysis of each rotor stage, for example, a compressor airfoil  100  disposed in a high pressure compressor section  32  and a compressor airfoil  100  disposed in a low pressure compressor section  24  of compressor section  6 . Moreover, the resulting reduced susceptibility to damage facilitates the implementation of tighter clearances between the compressor airfoil  100  and the abradable seal  122  and/or thinner compressor airfoil  100  architectures. 
     For example, with reference to  FIGS. 2-3 , a compressor airfoil  100  comprises an airfoil configured to compress and spin a fluid flow, such as a rotor. The compressor airfoil  100  has a leading edge  102 . For example, the leading edge  102  of a spinning compressor airfoil  100  such as a rotor is the edge of the compressor airfoil  100  toward which the compressor airfoil  100  is spinning. 
     The compressor airfoil  100  comprises a trailing edge  104 . For example, the trailing edge  104  of a spinning compressor airfoil  100  such as a rotor is the edge of the compressor airfoil  100  away from which the compressor airfoil  100  is spinning. 
     The compressor airfoil  100  has a radially outboard edge  106  that includes a tip of the compressor airfoil  100 , e.g., a surface connecting the trailing edge  104  and the leading edge  102  and forming the radially outermost boundary of the compressor airfoil  100 . 
     According to the disclosed embodiments, the compressor airfoil  100  has an abrasive coating  108 . An abrasive coating  108  may be disposed along a portion of at least one of the leading edge  102 , radially outboard edge  106 , and the trailing edge  104  of the compressor airfoil  100 . The abrasive coating  108  covers the radially outboard edge  106  and extends over a portion of the leading edge  102  and/or the trailing edge  104 . For instance, the abrasive coating  108  extends over the leading edge radius  114  ( FIG. 3 ) and/or the trailing edge radius  118  ( FIG. 3 ) and/or extends over a leading edge chamfer  116  ( FIG. 2 ) and/or a trailing edge chamfer  120  ( FIG. 2 ). 
     As used herein, “extend over” an article, means to coat a portion of the article, extending outward therefrom. For instance, to “extend over” the leading edge radius  114  ( FIG. 3 ) means to coat at least the portion of the leading edge  102  made up of the leading edge radius  114  ( FIG. 3 ) and to extend outward therefrom, locally thickening the compressor airfoil  100 . 
     The abrasive coating  108  comprises a uniform thickness, or is thicker in areas corresponding to expected deeper incursions into the abradable seal  122  or where greater cutting capability is desired over the life of the compressor airfoil  100 . For instance, the tip coating is a uniform thickness, or the tip coating has an increased thickness over a leading edge radius  114  ( FIG. 3 ) and/or the trailing edge radius  118  ( FIG. 3 ) and/or a leading edge chamfer  116  ( FIG. 2 ) and/or a trailing edge chamfer  120  ( FIG. 2 ). 
     The abrasive coating  108  comprises a frictional material selected to grind away a portion of an abradable seal  122 . For instance, the turbine engine  2  comprises an abradable seal  122  disposed annularly outward of the path of the compressor airfoil  100 . As the compressor airfoil  100  rotates, the leading edge  102 , radially outboard edge  106  and/or trailing edge  104  at least periodically impinge the abradable seal  122 . The abrasive coating  108  facilitates the grinding away of the abradable seal  122  proximate to the impingement of the compressor airfoil  100 . In this manner, the abradable seal  122  is worn away at the site of the impingement, after which the compressor airfoil  100  rotates with relative freedom, and yet a fluidic seal is formed about the radially outboard edge  106  of the compressor airfoil  100 , diminishing the leakage of fluid flow around the compressor airfoil  100  and directing the fluid flow through the rotational plane of the compressor airfoil  100 . 
     The leading edge  102  comprises a leading edge tip profile  110 . The leading edge tip profile  110  comprises a shape selected to ameliorate Hertzian stresses. For example, with reference to  FIG. 2 , the leading edge tip profile  110  comprises a leading edge radius  114 . With reference to  FIG. 3 , the leading edge tip profile  110  comprises a leading edge chamfer  116 . As such, the leading edge tip profile  110  comprises any shape selected to ameliorate Hertzian stress. 
     The trailing edge  104  comprises a trailing edge tip profile  112 . The trailing edge tip profile  112  comprises a shape selected to ameliorate Hertzian stresses. For example, with reference to  FIG. 2 , the trailing edge tip profile  112  comprises a trailing edge radius  118 . For example, with reference to  FIG. 3 , the trailing edge tip profile  112  comprises a trailing edge chamfer  120 . As such, the trailing edge tip profile  112  comprises any shape selected to ameliorate Hertzian stress. 
     With reference to both  FIGS. 2 and 3 , the trailing edge tip profile  112  and the leading edge tip profile  110  are one of same and different. For example, the trailing edge tip profile  112  comprises a trailing edge chamfer  120  or a trailing edge radius  118  or any other desired shape, while the leading edge tip profile  110  comprises a leading edge chamfer  116  or a leading edge radius  114 , or any other desired shape, and the shapes are optionally dissimilar, for instance a leading edge chamfer  116  and a trailing edge radius  118  or a leading edge radius  114  and a trailing edge chamfer  120 , or the like. 
     The leading edge radius  114  and/or trailing edge radius  118  comprise a variety of geometries or combinations of geometries. For instance, the leading edge radius  114  and/or trailing edge radius  118  comprise a radius that is at least one of polygonal, polyhedral, a complex curvature, conventional curvature, compound curvature, and/or any curvature. The radius is optionally at least one of a 2D and/or 3D projection of a relief cut such as having a curved shape when viewed in a plane perpendicular to the axis of rotation of the compressor airfoil  100 . Similarly, a leading edge chamfer  116  and/or a trailing edge chamfer  120  are optionally a variety of geometries or combinations of geometries such as a chamfer having a single angle, or a chamfer having multiple angles and/or the like. 
     In summary, with reference to  FIGS. 1-4 , a compressor airfoil  100  may have a leading edge  102  that has a leading edge tip profile  110  and a trailing edge  104  that has a trailing edge tip profile  112 . Each profile  110 ,  112  may have a radius and/or a chamfer. For instance, the leading edge tip profile  110  may have a leading edge chamfer  116  or a leading edge radius  114  and the trailing edge tip profile  112  may have a trailing edge chamfer  120  or a trailing edge radius  118 . An abrasive coating  108  may cover a portion of one or both profile  110 ,  112 , and also all or a portion of the radially outboard edge  106 . As the compressor airfoil  100  spins, the chamfer or radius of each profile  110 ,  112  ameliorates Hertzian stresses caused by incursions of the compressor airfoil  100  into the abradable seal  122 . The abrasive coating  108  further ameliorates the Hertzian stresses as well as assists the grinding away of the abradable seal  122 . This enhances the aerodynamic efficiency of the compressor airfoil  100  by ameliorating fluid leakage around the compressor airfoil  100 . 
     Methods of forming a compressor airfoil  100  are also presented. For instance, with reference to  FIG. 4 , a method  400  includes forming a compressor airfoil  100  bounded by a leading edge  102 , a trailing edge  104  and a radially outboard edge  106  connecting the leading edge  102  and the trailing edge  104  (step  402 ). The method  400  further includes forming a leading edge tip profile  110  comprising one of a leading edge radius  114  or a leading edge chamfer  116  (step  404 ) and forming a trailing edge tip profile  112  comprising one of a trailing edge radius  118  or a trailing edge chamfer  120  (step  406 ). The method  400  includes depositing an abrasive coating  108  on at least one of the trailing edge  104 , the leading edge  102  and the radially outboard edge  106  (step  408 ). The steps are carried out in parallel, or in sequences as desired. 
     Having discussed aspects of a compressor airfoil  100 , a compressor airfoil  100  is made of a single material, or different materials, or combinations of materials. For example, components of the system are made from metal. For example, aspects of a compressor airfoil  100  are metal, such as titanium, aluminum, steel, or stainless steel, though it alternatively comprises numerous other materials configured to provide support. Components of the system are optionally made from other materials such as, for example, composite, ceramic, plastics, polymers, alloys, glass, binder, epoxy, polyester, acrylic, or any material or combination of materials having desired material properties, such as heat tolerance, strength, stiffness, or weight. Portions of a compressor airfoil  100  as disclosed herein are optionally made of different materials or combinations of materials, and/or comprise coatings. Moreover, components of the compressor airfoil  100  are in some instances mixtures of different materials according to different ratios, gradients, and the like. 
     A compressor airfoil  100  thus has multiple materials, or any material configuration suitable to enhance or reinforce the resiliency and/or support of the system when subjected to wear in an aircraft operating environment or to satisfy other desired electromagnetic, chemical, physical, or material properties, for example coefficient of thermal expansion, ductility, weight, flexibility, strength, or heat tolerance. 
     One such material is an austenitic nickel-chromium-based alloy such as Inconel®, which is available from Special Metals Corporation of New Hartford, N.Y., USA. Other such material includes ceramic matrix composite (CMC). Further material includes refractory metal, for example, an alloy of titanium, such as, titanium-zirconium-molybdenum (TZM). 
     Benefits, other advantages, and solutions to problems have been described herein with regard to specific embodiments. The scope of the disclosure, however, is provided in the appended claims