Patent Publication Number: US-2019178094-A1

Title: Integrally bladed rotor

Description:
CROSS REFERENCE TO RELATED APPLICATION 
     This application claims priority to the U.S. Provisional Application No. 62/580,843, which was filed on Nov. 2, 2017, 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-energy exhaust gas flow. The high-energy 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 sections include airfoils supported on rotors. The airfoils may be separate parts assembled to a rotor or may also be integrally formed as part of the rotor. Forming the rotor and airfoils as a single part reduces the number of parts and eliminates the need for fastening systems for securing airfoils to the rotor. 
     Turbine engine manufacturers continue to seek improvements to turbine engines including improvements in assembly, manufacture, engine performance and propulsive efficiencies. 
     SUMMARY 
     In a featured embodiment, an integrally bladed rotor for a gas turbine engine includes a rotor portion with an outer periphery. At least one airfoil includes a suction side and a pressure side extending between a leading edge and a trailing edge. The at least one airfoil extends radially from the outer periphery and has an airfoil thickness between the suction side and the pressure side. A first thickness on at least one of the pressure side and suction side of the airfoil in addition to the airfoil thickness that extends radially from the outer periphery defines a crack propagation boundary. 
     In another embodiment according to the previous embodiment, the first thickness includes a first fillet between the outer periphery and the airfoil and further includes a second fillet radially outward from the first fillet. The second fillet has a second radius smaller than the first radius. 
     In another embodiment according to any of the previous embodiments, the first fillet extends radially from the outer periphery a first distance to an interface and the second fillet begins at the interface. 
     In another embodiment according to any of the previous embodiments, the interface is spaced apart from at least one of the suction side and pressure side of the airfoil a first width. 
     In another embodiment according to any of the previous embodiments, the second fillet extends from the interface to one of the pressure side and suction side of the airfoil. 
     In another embodiment according to any of the previous embodiments, the first distance includes a smooth transition from the outer periphery to the interface. 
     In another embodiment according to any of the previous embodiments, the first fillet and the second fillet are on one of the pressure side and the suction side. 
     In another embodiment according to any of the previous embodiments, the first fillet and the second fillet are disposed on both the suction side and the pressure side. 
     In another embodiment according to any of the previous embodiments, the first thickness includes a patch portion disposed on one of the suction side and the pressure side extending partway between the leading edge and the trailing edge. 
     In another featured embodiment, an integrally bladed rotor for a gas turbine engine includes a rotor portion with an outer periphery. At least one airfoil includes a suction side and a pressure side extending between a leading edge and a trailing edge. The at least one airfoil extends radially from the outer periphery. The airfoil includes an airfoil thickness between the pressure side and the suction side. A patch portion disposed on one of the suction side and the pressure side extends partway between the leading edge and the trailing edge. The patch portion includes a first thickness added to the airfoil thickness. 
     In another embodiment according to the previous embodiment, the patch portion includes a first fillet providing a smooth transition from the outer periphery. 
     In another embodiment according to any of the previous embodiments, the patch portion is disposed on the suction side at the leading edge. 
     In another embodiment according to any of the previous embodiments, the patch portion is spaced apart from the trailing edge. 
     In another embodiment according to any of the previous embodiments, the patch portion is spaced apart from the leading edge. 
     In another featured embodiment, a method of fabricating an integrally bladed rotor for a gas turbine engine includes forming a rotor portion with an outer periphery. At least one airfoil is formed extending radially from the outer periphery to include a suction side and a pressure side extending between a leading edge and a trailing edge. The at least one airfoil is formed to include an airfoil thickness between the pressure side and the suction side. A first thickness is formed on at least one of the pressure side and suction side of the airfoil in addition to the airfoil thickness. 
     In another embodiment according to the previous embodiment, forming the first thickness as a patch portion disposed on one of the suction side and the pressure side extending partway between the leading edge and the trailing edge. 
     In another embodiment according to any of the previous embodiments, forming the patch portion includes forming the patch portion on the suction side at the leading edge and spaced apart from the trailing edge. 
     In another embodiment according to any of the previous embodiments, forming the first thickness includes forming a first fillet between the outer periphery and at least one of the pressure side and suction side of the airfoil to have a first radius and forming a second fillet radially outward from the first fillet to have a second radius smaller than the first radius. 
     In another embodiment according to any of the previous embodiments, forming the first fillet to extend radially from the outer periphery a first distance to an interface and forming the second fillet to begin at the interface. 
     Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure 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. 
     These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic view of an example gas turbine engine. 
         FIG. 2  is a perspective view of an example integrally bladed rotor. 
         FIG. 3  is a perspective view of an example airfoil embodiment. 
         FIG. 4  is a cross-sectional view of a portion of the example airfoil embodiment. 
         FIG. 5  is an enlarged perspective view of a portion of an example airfoil embodiment. 
         FIG. 6  is another perspective view of an example airfoil embodiment. 
         FIG. 7  is a perspective view of another example airfoil embodiment. 
         FIG. 8  is a cross-sectional view of a portion of the example airfoil shown in  FIG. 7 . 
         FIG. 9  is a cross-section of another portion of the example airfoil shown in  FIG. 7 . 
     
    
    
     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  18 , 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  58  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  58  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  58  includes airfoils  60  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 (‘TSFCT’)”—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). 
     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. 
     Referring to  FIG. 2  with continued reference to  FIG. 1 , the example gas turbine engine  20  includes the compressor section  52  that includes a plurality of integrally bladed rotors  62  (IBR). Each of the IBRs  62  includes a rotor portion  76  defining an outer periphery  64 . A plurality of airfoils  66  extend upward from the outer periphery  64 . The IBR  62  is a one-piece part with portions that define, among other features, the rotor  76 , periphery  64  and the airfoils  66 . 
     Referring to  FIG. 3  with continued reference to  FIG. 2 , each of the plurality of airfoils  66  includes a leading edge  68 , a trailing edge  70 , a suction side  72  and a pressure side  74 . Each of the airfoils  66  extends radially outward from the periphery  64  defined in the IBR  62 . The airfoils  66  extend from a root portion  80  to a tip portion  78 . The root portion  80  is defined at the periphery  64  of the rotor portion  76 . A thickness  84  is disposed near the root  80  and extends between the peripheral surface  64  and side surfaces of the airfoil  66 . 
     Referring to  FIG. 4  with continued reference to  FIGS. 2 and 3 , the example thickness  84  is disposed at the root portion  80  about the airfoil  66  and provides a boundary to prevent crack propagation from the airfoil  66  into the rotor portion  76  of the IBR  62 . In this example, the thickness  84  extends outward from an airfoil thickness  82  defined between the pressure side  72  and the suction side  74 . 
     Referring to  FIG. 5  with continued reference to  FIGS. 3 and 4 , in one disclosed embodiment, the first thickness  84  includes a first fillet  86  that extends from the rotor periphery  64  to an interface  98  and a second fillet  88  that extends radially outward from the interface  98 . The first fillet  86  includes a transition surface  100  that extends from the periphery  64  to the interface  98 . The interface  98  is disposed a distance  94  above the peripheral surface  64  and a width  96  away from the pressure side  72  and suction side  74  of the airfoil  66 . The disclosed interface  98  is the interface between the first fillet  86  and the second fillet and is the location where the surface  100  transitions from a first radius  90  to a second radius  92 . Accordingly, the second fillet  88  begins at the interface  98  and extends upward radially at the second radius  92  into a smooth transition that merges with the pressure and suction surfaces of the airfoil  66 . 
     In the disclosed example embodiment, the first radius  90  is larger than the second radius  92 . In one disclosed embodiment, the first radius  90  is between one third and one half greater than the second radius  92 . In another disclosed embodiment, the first radius  90  is about 0.120 inches (3.048 mm) and the second radius is 0.080 inches (2.032 mm). In another disclosed embodiment the first radius is about 0.090 inches (2.286 mm) and the second radius is about 0.050 inches (1.27 mm). In another disclosed dimensional embodiment, the first radius is about 0.150 inches (3.81 mm) and the second radius is about 0.0120 inches (0.3048 mm). Moreover, in one example embodiment, the width  96  is between about 0.010 inches (0.254 mm) and 0.030 inches (0.762 mm). In other disclosed embodiment the width  96  is about 0.020 inches (0.508 mm). It should be understood, that the disclosed dimensional embodiment is provided by way of example and other radiuses and widths could be utilized and are within the contemplation of this disclosure. 
     It should be understood that although dimensional embodiments are disclosed by way of example, the first fillet  86  is larger than the second fillet  88 . The specific relative size between the first fillet  86  and the second fillet  88  may be different to provide a predefined stress propagation path that prevents crack propagation radially inward into the rotor  76 . 
     Referring to  FIG. 6 , with continued reference to  FIG. 5 , the separation at the interface  98  between the first fillet  86  and the second fillet  88  defines a crack propagation boundary schematically shown at  135 . A potential crack schematically referred to as  140  is prevented from propagating radially inward toward the rotor portion  76  by the thicker portions of the airfoil defined by the first fillet  86 . Instead, the crack  140  propagates in direction substantially along and parallel to the interface  98 . Accordingly, the interface  98  defines the crack propagation boundary and prevents cracks from propagating into the rotor portion  76 . 
     Referring to  FIGS. 7, 8 and 9 , another example IBR  102  embodiment is schematically illustrated and includes an airfoil  106  with a suction side  112  and a pressure side  114  that extends between a leading edge  108  and a trailing edge  110 . The airfoil  106  extends upward from a rotor peripheral surface  122  and includes a patch portion  104 . In this example, the patch portion  104  is disposed on the pressure side  114  and extends a distance  128  from the leading edge  108  toward the trailing edge  110 . The patch  104  is therefore spaced apart a distance  130  away from the trailing edge  130 . The patch  104  is disposed at a location between the airfoil  102  and peripheral surface  122  that defines a boundary that prevents crack propagation into the rotor  120 . Moreover, the area of the increased thickness  126  provided by the patch  104  is based on stress analysis of potential crack propagation and may vary in location and thickness. 
     The patch  104  is an increased thickness indicated at  126  ( FIG. 8 ) that is greater than the airfoil thickness  124 . The airfoil thickness  124  may vary depending on the airfoil shape between the leading edge and the trailing edge. The patch portion  104  includes a thickness  126  in addition to the airfoil thickness  124  of the airfoil  102  in a specific location near the leading edge  108 . The location of the thickness  126  is provided based on analysis of stresses inflicted on the airfoil  102  during operation. The increased thickness  126  is shown on one side of the airfoil  102 , but may extend to both sides of the airfoil  102 . Moreover, the patch  104  may extend different distances  128  toward the trailing edge  110  as determined to define a boundary to possible crack propagation radially inward. Moreover, the increased thickness  126  provided by the patch  104  could be spaced from the leading edge  108  or any position along the interface between the surface  122  and the airfoil  102  where a reduction in operating stress is required to direct crack propagation way from the rotor  120 . 
     Accordingly, the example disclosed IBR  62  includes airfoils  66 ,  102  with features that define crack propagation boundaries to prevent cracks from propagating radially inward to into rotor portions. 
     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 this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.